X-29 in Protective Cover Being Transported by Truck to Dryden

NASA Technical Reports Server (NTRS)

1988-01-01

In a stark juxtaposition of nature and technology, the second X-29 forward-swept-wing research aircraft is shown here passing by one of the classic, spiny Joshua trees that populate the Mojave desert while being transported by truck to NASA's Ames-Dryden Flight Research Facility (later the Dryden Flight Research Center), Edwards, California, on November 7, 1988. The aircraft, with its protective covering, traveled by ship from the manufacturer's plant on Long Island through the Panama Canal to Port Hueneme and then was trucked to Dryden. X-29 No. 2 was used in a high angle-of-attack research program which began in spring 1989. Two X-29 aircraft, featuring one of the most unusual designs in aviation history, flew at the Ames-Dryden Flight Research Facility (now the Dryden Flight Research Center, Edwards, California) from 1984 to 1992. The fighter-sized X-29 technology demonstrators explored several concepts and technologies including: the use of advanced composites in aircraft construction; variable-camber wing surfaces; a unique forward- swept wing and its thin supercritical airfoil; strakes; close-coupled canards; and a computerized fly-by-wire flight control system used to maintain control of the otherwise unstable aircraft. Research results showed that the configuration of forward-swept wings, coupled with movable canards, gave pilots excellent control response at angles of attack of up to 45 degrees. During its flight history, the X-29 aircraft flew 422 research missions and a total of 436 missions. Sixty of the research flights were part of the X-29 follow-on 'vortex control' phase. The forward-swept wing of the X-29 resulted in reverse airflow, toward the fuselage rather than away from it, as occurs on the usual aft-swept wing. Consequently, on the forward-swept wing, the ailerons remained unstalled at high angles of attack. This provided better airflow over the ailerons and prevented stalling (loss of lift) at high angles of attack. Introduction of composite materials in the 1970s opened a new field of aircraft construction. It also made possible the construction of the X-29's thin supercritical wing. State-of-the-art composites allowed aeroelastic tailoring which, in turn, allowed the wing some bending but limited twisting and eliminated structural divergence within the flight envelope (i.e. deformation of the wing or the wing breaking off in flight). Additionally, composite materials allowed the wing to be sufficiently rigid for safe flight without adding an unacceptable weight penalty. The X-29 project consisted of two phases plus the follow-on vortex-control phase. Phase 1 demonstrated that the forward sweep of the X-29 wings kept the wing tips unstalled at the moderate angles of attack flown in that phase (a maximum of 21 degrees). Phase I also demonstrated that the aeroelastic tailored wing prevented structural divergence of the wing within the flight envelope, and that the control laws and control-surface effectiveness were adequate to provide artificial stability for an otherwise unstable aircraft. Phase 1 further demonstrated that the X-29 configuration could fly safely and reliably, even in tight turns. During Phase 2 of the project, the X-29, flying at an angle of attack of up to 67 degrees, demonstrated much better control and maneuvering qualities than computational methods and simulation models had predicted . During 120 research flights in this phase, NASA, Air Force, and Grumman project pilots reported the X-29 aircraft had excellent control response to an angle of attack of 45 degrees and still had limited controllability at a 67-degree angle of attack. This controllability at high angles of attack can be attributed to the aircraft's unique forward-swept wing- canard design. The NASA/Air Force-designed high-gain flight control laws also contributed to the good flying qualities. During the Air Force-initiated vortex-control phase, the X-29 successfully demonstrated vortex flow control (VFC). This VFC was more effective than expected in generating yaw forces, especially in high angles of attack where the rudder is less effective. VFC was less effective in providing control when sideslip (wind pushing on the side of the aircraft) was present, and it did little to decrease rocking oscillation of the aircraft. The X-29 vehicle was a single-engine aircraft, 48.1 feet long with a wing span of 27.2 feet. Each aircraft was powered by a General Electric F404-GE-400 engine producing 16,000 pounds of thrust. The program was a joint effort of the Department of Defense's Defense Advanced Research Projects Agency (DARPA), the U.S. Air Force, the Ames-Dryden Flight Research Facility, the Air Force Flight Test Center, and the Grumman Corporation. The program was managed by the Air Force's Wright Laboratory, Wright Patterson Air Force Base, Ohio.

An Overview of Controls and Flying Qualities Technology on the F/A-18 High Alpha Research Vehicle

NASA Technical Reports Server (NTRS)

Pahle, Joseph W.; Wichman, Keith D.; Foster, John V.; Bundick, W. Thomas

1996-01-01

The NASA F/A-18 High Alpha Research Vehicle (HARV) has been the flight test bed of a focused technology effort to significantly increase maneuvering capability at high angles of attack. Development and flight test of control law design methodologies, handling qualities metrics, performance guidelines, and flight evaluation maneuvers are described. The HARV has been modified to include two research control effectors, thrust vectoring, and actuated forebody strakes in order to provide increased control power at high angles of attack. A research flight control system has been used to provide a flexible, easily modified capability for high-angle-of-attack research controls. Different control law design techniques have been implemented and flight-tested, including eigenstructure assignment, variable gain output feedback, pseudo controls, and model-following. Extensive piloted simulation has been used to develop nonlinear performance guide-lines and handling qualities criteria for high angles of attack. This paper reviews the development and evaluation of technologies useful for high-angle-of-attack control. Design, development, and flight test of the research flight control system, control laws, flying qualities specifications, and flight test maneuvers are described. Flight test results are used to illustrate some of the lessons learned during flight test and handling qualities evaluations.

X-31 Mishap: Lessons Learned

NASA Technical Reports Server (NTRS)

Larson, Richard R.

2007-01-01

The experimental X-31 High Angle of Attack Research Aircraft crashed during a 1995 test mission flight conducted by NASA at Edwards Air Force Base, California. The pilot lost control of the airplane and was forced to eject, sustaining a permanent back injury that ended his flying career. Prior to this incident the airplane had a perfect record of several hundred non-eventful flights supported by an experienced team. During the subsequent investigation by a mishap committee it was discovered that a series of cascading events contributed to this accident. Some of the identified contributing factors that resulted in this mishap are common to aircraft design and to flight-test in general. The mistakes and the solutions are presented here so that the flight-test community may consider and learn from them. The primary cause of the crash was icing and, ultimately, a complete blockage of the pitot-static nose probe. The icing was caused by a freak weather phenomenon that was neither expected nor known to exist on the day of the mishap. The normal probe had been replaced with a special Kiel probe to allow total pressure measurements of up to 70 degrees angle of attack for flight-test purposes. The Kiel probe did not include a heater, because it was assumed that the airplane would not be flown in the clouds or in conditions conducive to icing. This assumption was later proven to be incorrect. The iced Kiel probe caused incorrect gain scheduling in the flight control system, resulting in an unstable aircraft. This failure was essentially undetected because of a faulty design in the flight control system architecture. There were, however, also a number of other issues that lead up to this situation that never should have happened. This presentation discusses what the issues were that contributed to the incident. After the incident was investigated, some of these issues were addressed and some changes were made. The second X-31 aircraft flew the remainder of the flight tests, and the program was successfully completed without incident. This presentation also shows a video of the mishap including lessons learned, and the changes that were made to resume the flight-test program are presented.

Flight Testing of Novel Compliant Spines for Passive Wing Morphing on Ornithopters

NASA Technical Reports Server (NTRS)

Wissa, Aimy; Guerreiro, Nelson; Grauer, Jared; Altenbuchner, Cornelia; Hubbard, James E., Jr.; Tummala, Yashwanth; Frecker, Mary; Roberts, Richard

2013-01-01

Unmanned Aerial Vehicles (UAVs) are proliferating in both the civil and military markets. Flapping wing UAVs, or ornithopters, have the potential to combine the agility and maneuverability of rotary wing aircraft with excellent performance in low Reynolds number flight regimes. The purpose of this paper is to present new free flight experimental results for an ornithopter equipped with one degree of freedom (1DOF) compliant spines that were designed and optimized in terms of mass, maximum von-Mises stress, and desired wing bending deflections. The spines were inserted in an experimental ornithopter wing spar in order to achieve a set of desired kinematics during the up and down strokes of a flapping cycle. The ornithopter was flown at Wright Patterson Air Force Base in the Air Force Research Laboratory Small Unmanned Air Systems (SUAS) indoor flight facility. Vicon motion tracking cameras were used to track the motion of the vehicle for five different wing configurations. The effect of the presence of the compliant spine on wing kinematics and leading edge spar deflection during flight is presented. Results show that the ornithopter with the compliant spine inserted in its wing reduced the body acceleration during the upstroke which translates into overall lift gains.

Ares I Aerodynamic Testing at the Boeing Polysonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Pinier, Jeremy T.; Niskey, Charles J.; Hanke, Jeremy L.; Tomek, William G.

2011-01-01

Throughout three full design analysis cycles, the Ares I project within the Constellation program has consistently relied on the Boeing Polysonic Wind Tunnel (PSWT) for aerodynamic testing of the subsonic, transonic and supersonic portions of the atmospheric flight envelope (Mach=0.5 to 4.5). Each design cycle required the development of aerodynamic databases for the 6 degree-of-freedom (DOF) forces and moments, as well as distributed line-loads databases covering the full range of Mach number, total angle-of-attack, and aerodynamic roll angle. The high fidelity data collected in this facility has been consistent with the data collected in NASA Langley s Unitary Plan Wind Tunnel (UPWT) at the overlapping condition ofMach=1.6. Much insight into the aerodynamic behavior of the launch vehicle during all phases of flight was gained through wind tunnel testing. Important knowledge pertaining to slender launch vehicle aerodynamics in particular was accumulated. In conducting these wind tunnel tests and developing experimental aerodynamic databases, some challenges were encountered and are reported as lessons learned in this paper for the benefit of future crew launch vehicle aerodynamic developments.

Effects of Type and Strength of Force Feedback on Movement Time in a Target Selection Task

NASA Technical Reports Server (NTRS)

Rorie, Robert Conrad; Vu, Kim-Phuong L.; Marayong, Panadda; Robles, Jose; Strybel, Thomas Z.; Battiste, Vernol

2013-01-01

Future cockpits will likely include new onboard technologies, such as cockpit displays of traffic information, to help support future flight deck roles and responsibilities. These new technologies may benefit from multimodal feedback to aid pilot information processing. The current study investigated the effects of multiple levels of force feedback on operator performance in an aviation task. Participants were presented with two different types of force feedback (gravitational and spring force feedback) for a discrete targeting task, with multiple levels of gain examined for each force feedback type. Approach time and time in target were recorded. Results suggested that the two highest levels of gravitational force significantly reduced approach times relative to the lowest level of gravitational force. Spring force level only affected time in target. Implications of these findings for the design of future cockpit displays will be discussed.

Digital control of highly augmented combat rotorcraft

NASA Technical Reports Server (NTRS)

Tischler, Mark B.

1987-01-01

Proposed concepts for the next generation of combat helicopters are to be embodied in a complex, highly maneuverable, multiroled vehicle with avionics systems. Single pilot and nap-of-the-Earth operations require handling qualities which minimize the involvement of the pilot in basic stabilization tasks. To meet these requirements will demand a full authority, high-gain, multimode, multiply-redundant, digital flight-control system. The gap between these requirements and current low-authority, low-bandwidth operational rotorcraft flight-control technology is considerable. This research aims at smoothing the transition between current technology and advanced concept requirements. The state of the art of high-bandwidth digital flight-control systems are reviewed; areas of specific concern for flight-control systems of modern combat are exposed; and the important concepts are illustrated in design and analysis of high-gain, digital systems with a detailed case study involving a current rotorcraft system. Approximate and exact methods are explained and illustrated for treating the important concerns which are unique to digital systems.

Field Flight Dynamics of Hummingbirds during Territory Encroachment and Defense

PubMed Central

Sholtis, Katherine M.; Shelton, Ryan M.; Hedrick, Tyson L.

2015-01-01

Hummingbirds are known to defend food resources such as nectar sources from encroachment by competitors (including conspecifics). These competitive intraspecific interactions provide an opportunity to quantify the biomechanics of hummingbird flight performance during ecologically relevant natural behavior. We recorded the three-dimensional flight trajectories of Ruby-throated Hummingbirds defending, being chased from and freely departing from a feeder. These trajectories allowed us to compare natural flight performance to earlier laboratory measurements of maximum flight speed, aerodynamic force generation and power estimates. During field observation, hummingbirds rarely approached the maximal flight speeds previously reported from wind tunnel tests and never did so during level flight. However, the accelerations and rates of change in kinetic and potential energy we recorded indicate that these hummingbirds likely operated near the maximum of their flight force and metabolic power capabilities during these competitive interactions. Furthermore, although birds departing from the feeder while chased did so faster than freely-departing birds, these speed gains were accomplished by modulating kinetic and potential energy gains (or losses) rather than increasing overall power output, essentially trading altitude for speed during their evasive maneuver. Finally, the trajectories of defending birds were directed toward the position of the encroaching bird rather than the feeder. PMID:26039101

Field Flight Dynamics of Hummingbirds during Territory Encroachment and Defense.

PubMed

Sholtis, Katherine M; Shelton, Ryan M; Hedrick, Tyson L

2015-01-01

Hummingbirds are known to defend food resources such as nectar sources from encroachment by competitors (including conspecifics). These competitive intraspecific interactions provide an opportunity to quantify the biomechanics of hummingbird flight performance during ecologically relevant natural behavior. We recorded the three-dimensional flight trajectories of Ruby-throated Hummingbirds defending, being chased from and freely departing from a feeder. These trajectories allowed us to compare natural flight performance to earlier laboratory measurements of maximum flight speed, aerodynamic force generation and power estimates. During field observation, hummingbirds rarely approached the maximal flight speeds previously reported from wind tunnel tests and never did so during level flight. However, the accelerations and rates of change in kinetic and potential energy we recorded indicate that these hummingbirds likely operated near the maximum of their flight force and metabolic power capabilities during these competitive interactions. Furthermore, although birds departing from the feeder while chased did so faster than freely-departing birds, these speed gains were accomplished by modulating kinetic and potential energy gains (or losses) rather than increasing overall power output, essentially trading altitude for speed during their evasive maneuver. Finally, the trajectories of defending birds were directed toward the position of the encroaching bird rather than the feeder.

F-8C adaptive flight control laws

NASA Technical Reports Server (NTRS)

Hartmann, G. L.; Harvey, C. A.; Stein, G.; Carlson, D. N.; Hendrick, R. C.

1977-01-01

Three candidate digital adaptive control laws were designed for NASA's F-8C digital flyby wire aircraft. Each design used the same control laws but adjusted the gains with a different adaptative algorithm. The three adaptive concepts were: high-gain limit cycle, Liapunov-stable model tracking, and maximum likelihood estimation. Sensors were restricted to conventional inertial instruments (rate gyros and accelerometers) without use of air-data measurements. Performance, growth potential, and computer requirements were used as criteria for selecting the most promising of these candidates for further refinement. The maximum likelihood concept was selected primarily because it offers the greatest potential for identifying several aircraft parameters and hence for improved control performance in future aircraft application. In terms of identification and gain adjustment accuracy, the MLE design is slightly superior to the other two, but this has no significant effects on the control performance achievable with the F-8C aircraft. The maximum likelihood design is recommended for flight test, and several refinements to that design are proposed.

Adaptive Augmenting Control Flight Characterization Experiment on an F/A-18

NASA Technical Reports Server (NTRS)

VanZwieten, Tannen S.; Orr, Jeb S.; Wall, John H.; Gilligan, Eric T.

2014-01-01

This paper summarizes the Adaptive Augmenting Control (AAC) flight characterization experiments performed using an F/A-18 (TN 853). AAC was designed and developed specifically for launch vehicles, and is currently part of the baseline autopilot design for NASA's Space Launch System (SLS). The scope covered here includes a brief overview of the algorithm (covered in more detail elsewhere), motivation and benefits of flight testing, top-level SLS flight test objectives, applicability of the F/A-18 as a platform for testing a launch vehicle control design, test cases designed to fully vet the AAC algorithm, flight test results, and conclusions regarding the functionality of AAC. The AAC algorithm developed at Marshall Space Flight Center is a forward loop gain multiplicative adaptive algorithm that modifies the total attitude control system gain in response to sensed model errors or undesirable parasitic mode resonances. The AAC algorithm provides the capability to improve or decrease performance by balancing attitude tracking with the mitigation of parasitic dynamics, such as control-structure interaction or servo-actuator limit cycles. In the case of the latter, if unmodeled or mismodeled parasitic dynamics are present that would otherwise result in a closed-loop instability or near instability, the adaptive controller decreases the total loop gain to reduce the interaction between these dynamics and the controller. This is in contrast to traditional adaptive control logic, which focuses on improving performance by increasing gain. The computationally simple AAC attitude control algorithm has stability properties that are reconcilable in the context of classical frequency-domain criteria (i.e., gain and phase margin). The algorithm assumes that the baseline attitude control design is well-tuned for a nominal trajectory and is designed to adapt only when necessary. Furthermore, the adaptation is attracted to the nominal design and adapts only on an as-needed basis (see Figure 1). The MSFC algorithm design was formulated during the Constellation Program and reached a high maturity level during SLS through simulation-based development and internal and external analytical review. The AAC algorithm design has three summary-level objectives: (1) "Do no harm;" return to baseline control design when not needed, (2) Increase performance; respond to error in ability of vehicle to track command, and (3) Regain stability; respond to undesirable control-structure interaction or other parasitic dynamics. AAC has been successfully implemented as part of the Space Launch System baseline design, including extensive testing in high-fidelity 6-DOF simulations the details of which are described in [1]. The Dryden Flight Research Center's F/A-18 Full-Scale Advanced Systems Testbed (FAST) platform is used to conduct an algorithm flight characterization experiment intended to fully vet the aforementioned design objectives. FAST was specifically designed with this type of test program in mind. The onboard flight control system has full-authority experiment control of ten aerodynamic effectors and two throttles. It has production and research sensor inputs and pilot engage/disengage and real-time configuration of up to eight different experiments on a single flight. It has failure detection and automatic reversion to fail-safe mode. The F/A-18 aircraft has an experiment envelope cleared for full-authority control and maneuvering and exhibits characteristics for robust recovery from unusual attitudes and configurations aided by the presence of a qualified test pilot. The F/A-18 aircraft has relatively high mass and inertia with exceptional performance; the F/A-18 also has a large thrust-to-weight ratio, owing to its military heritage. This enables the simulation of a portion of the ascent trajectory with a high degree of dynamic similarity to a launch vehicle, and the research flight control system can simulate unstable longitudinal dynamics. Parasitic dynamics such as slosh and bending modes, as well as atmospheric disturbances, are being produced by the airframe via modification of bending filters and the use of secondary control surfaces, including leading and trailing edge flaps, symmetric ailerons, and symmetric rudders. The platform also has the ability to inject signals in flight to simulate structural mode resonances or other challenging dynamics. This platform also offers more test maneuvers and longer maneuver times than a single rocket or missile test, which provides ample opportunity to fully and repeatedly exercise all aspects of the algorithm. Prior to testing on an F/A-18, AAC was the only component of the SLS autopilot design that had not been flight tested. The testing described in this paper raises the Technology Readiness Level (TRL) early in the SLS Program and is able to demonstrate its capabilities and robustness in a flight environment.

Eigenstructure Assignment for Fault Tolerant Flight Control Design

NASA Technical Reports Server (NTRS)

Sobel, Kenneth; Joshi, Suresh (Technical Monitor)

2002-01-01

In recent years, fault tolerant flight control systems have gained an increased interest for high performance military aircraft as well as civil aircraft. Fault tolerant control systems can be described as either active or passive. An active fault tolerant control system has to either reconfigure or adapt the controller in response to a failure. One approach is to reconfigure the controller based upon detection and identification of the failure. Another approach is to use direct adaptive control to adjust the controller without explicitly identifying the failure. In contrast, a passive fault tolerant control system uses a fixed controller which achieves acceptable performance for a presumed set of failures. We have obtained a passive fault tolerant flight control law for the F/A-18 aircraft which achieves acceptable handling qualities for a class of control surface failures. The class of failures includes the symmetric failure of any one control surface being stuck at its trim value. A comparison was made of an eigenstructure assignment gain designed for the unfailed aircraft with a fault tolerant multiobjective optimization gain. We have shown that time responses for the unfailed aircraft using the eigenstructure assignment gain and the fault tolerant gain are identical. Furthermore, the fault tolerant gain achieves MIL-F-8785C specifications for all failure conditions.

A variable-gain output feedback control design approach

NASA Technical Reports Server (NTRS)

Haylo, Nesim

1989-01-01

A multi-model design technique to find a variable-gain control law defined over the whole operating range is proposed. The design is formulated as an optimal control problem which minimizes a cost function weighing the performance at many operating points. The solution is obtained by embedding into the Multi-Configuration Control (MCC) problem, a multi-model robust control design technique. In contrast to conventional gain scheduling which uses a curve fit of single model designs, the optimal variable-gain control law stabilizes the plant at every operating point included in the design. An iterative algorithm to compute the optimal control gains is presented. The methodology has been successfully applied to reconfigurable aircraft flight control and to nonlinear flight control systems.

High-speed civil transport flight- and propulsion-control technological issues

NASA Technical Reports Server (NTRS)

Ray, J. K.; Carlin, C. M.; Lambregts, A. A.

1992-01-01

Technology advances required in the flight and propulsion control system disciplines to develop a high speed civil transport (HSCT) are identified. The mission and requirements of the transport and major flight and propulsion control technology issues are discussed. Each issue is ranked and, for each issue, a plan for technology readiness is given. Certain features are unique and dominate control system design. These features include the high temperature environment, large flexible aircraft, control-configured empennage, minimizing control margins, and high availability and excellent maintainability. The failure to resolve most high-priority issues can prevent the transport from achieving its goals. The flow-time for hardware may require stimulus, since market forces may be insufficient to ensure timely production. Flight and propulsion control technology will contribute to takeoff gross weight reduction. Similar technology advances are necessary also to ensure flight safety for the transport. The certification basis of the HSCT must be negotiated between airplane manufacturers and government regulators. Efficient, quality design of the transport will require an integrated set of design tools that support the entire engineering design team.

The aerodynamics of flight in an insect flight-mill

PubMed Central

Barkan, Shay; Soroker, Victoria

2017-01-01

Predicting the dispersal of pest insects is important for pest management schemes. Flight-mills provide a simple way to evaluate the flight potential of insects, but there are several complications in relating tethered-flight to natural flight. We used high-speed video to evaluate the effect of flight-mill design on flight of the red palm weevil (Rynchophorous ferruginneus) in four variants of a flight-mill. Two variants had the rotating radial arm pivoted on the main shaft of the rotation axis, allowing freedom to elevate the arm as the insect applied lift force. Two other variants had the pivot point fixed, restricting the radial arm to horizontal motion. Beetles were tethered with their lateral axis horizontal or rotated by 40°, as in a banked turn. Flight-mill type did not affect flight speed or wing-beat frequency, but did affect flapping kinematics. The wingtip internal to the circular trajectory was always moved faster relative to air, suggesting that the beetles were attempting to steer in the opposite direction to the curved trajectory forced by the flight-mill. However, banked beetles had lower flapping asymmetry, generated higher lift forces and lost more of their body mass per time and distance flown during prolonged flight compared to beetles flying level. The results indicate, that flapping asymmetry and low lift can be rectified by tethering the beetle in a banked orientation, but the flight still does not correspond directly to free-flight. This should be recognized and taken into account when designing flight-mills and interoperating their data. PMID:29091924

The aerodynamics of flight in an insect flight-mill.

PubMed

Ribak, Gal; Barkan, Shay; Soroker, Victoria

2017-01-01

Predicting the dispersal of pest insects is important for pest management schemes. Flight-mills provide a simple way to evaluate the flight potential of insects, but there are several complications in relating tethered-flight to natural flight. We used high-speed video to evaluate the effect of flight-mill design on flight of the red palm weevil (Rynchophorous ferruginneus) in four variants of a flight-mill. Two variants had the rotating radial arm pivoted on the main shaft of the rotation axis, allowing freedom to elevate the arm as the insect applied lift force. Two other variants had the pivot point fixed, restricting the radial arm to horizontal motion. Beetles were tethered with their lateral axis horizontal or rotated by 40°, as in a banked turn. Flight-mill type did not affect flight speed or wing-beat frequency, but did affect flapping kinematics. The wingtip internal to the circular trajectory was always moved faster relative to air, suggesting that the beetles were attempting to steer in the opposite direction to the curved trajectory forced by the flight-mill. However, banked beetles had lower flapping asymmetry, generated higher lift forces and lost more of their body mass per time and distance flown during prolonged flight compared to beetles flying level. The results indicate, that flapping asymmetry and low lift can be rectified by tethering the beetle in a banked orientation, but the flight still does not correspond directly to free-flight. This should be recognized and taken into account when designing flight-mills and interoperating their data.

Rotor-state feedback in the design of flight control laws for a hovering helicopter

NASA Technical Reports Server (NTRS)

Takahashi, Marc D.

1994-01-01

The use of rigid-body and rotor-state feedback gains in the design of helicopter flight control laws was investigated analytically on a blade element, articulated rotor, helicopter model. The study was conducted while designing a control law to meet an existing military rotorcraft handling qualities design specification (ADS-33C) in low-speed flight. A systematic approach to meet this specification was developed along with an assessment of the function of these gains in the feedback loops. Using the results of this assessment, the pitch and roll crossover behavior was easily modified by adjusting the body attitude and rotor-flap feedback gains. Critical to understanding the feedback gains is that the roll and pitch rate dynamics each have second-order behavior, not the classic first-order behavior, which arises from a quasi-static rotor, six degree-of-freedom model.

Flight test of the X-29A at high angle of attack: Flight dynamics and controls

NASA Technical Reports Server (NTRS)

Bauer, Jeffrey E.; Clarke, Robert; Burken, John J.

1995-01-01

The NASA Dryden Flight Research Center has flight tested two X-29A aircraft at low and high angles of attack. The high-angle-of-attack tests evaluate the feasibility of integrated X-29A technologies. More specific objectives focus on evaluating the high-angle-of-attack flying qualities, defining multiaxis controllability limits, and determining the maximum pitch-pointing capability. A pilot-selectable gain system allows examination of tradeoffs in airplane stability and maneuverability. Basic fighter maneuvers provide qualitative evaluation. Bank angle captures permit qualitative data analysis. This paper discusses the design goals and approach for high-angle-of-attack control laws and provides results from the envelope expansion and handling qualities testing at intermediate angles of attack. Comparisons of the flight test results to the predictions are made where appropriate. The pitch rate command structure of the longitudinal control system is shown to be a valid design for high-angle-of-attack control laws. Flight test results show that wing rock amplitude was overpredicted and aileron and rudder effectiveness were underpredicted. Flight tests show the X-29A airplane to be a good aircraft up to 40 deg angle of attack.

Cosmos 2229

NASA Technical Reports Server (NTRS)

Edgerton, V. Reggie; Roy, Roland R.; Hodgson, John A.

1993-01-01

The 6 weeks preflight activities of the Cosmos project during 1993 included: modification of EMG connector to improve the reliability of EMG recording; 24 hour cage activity recording from all but two of the flight animals (monkeys); attempts to record from flight candidates during foot lever task; and force transducer calibrations on all flight candidate animals. The 4 week postflight recordings included: postflight recordings from flight animals; postflight recordings on 3 control (non-flight) animals; postflight recalibration of force transducers on 1 flight and 4 control (non-flight) animals; and attempts to record EMG and video data from the flight animals during postflight locomotion and postural activity. The flight EMG recordings suggest that significant changes in muscle control may occur in spaceflight. It is also clear from recordings that levels of EMG recorded during spaceflight can attain values similar to those measured on earth. Amplifier gain settings should therefore probably not be changed for spaceflight.

Longitudinal-control design approach for high-angle-of-attack aircraft

NASA Technical Reports Server (NTRS)

Ostroff, Aaron J.; Proffitt, Melissa S.

1993-01-01

This paper describes a control synthesis methodology that emphasizes a variable-gain output feedback technique that is applied to the longitudinal channel of a high-angle-of-attack aircraft. The aircraft is a modified F/A-18 aircraft with thrust-vectored controls. The flight regime covers a range up to a Mach number of 0.7; an altitude range from 15,000 to 35,000 ft; and an angle-of-attack (alpha) range up to 70 deg, which is deep into the poststall region. A brief overview is given of the variable-gain mathematical formulation as well as a description of the discrete control structure used for the feedback controller. This paper also presents an approximate design procedure with relationships for the optimal weights for the selected feedback control structure. These weights are selected to meet control design guidelines for high-alpha flight controls. Those guidelines that apply to the longitudinal-control design are also summarized. A unique approach is presented for the feed-forward command generator to obtain smooth transitions between load factor and alpha commands. Finally, representative linear analysis results and nonlinear batch simulation results are provided.

Flight Force Measurements on a Spacecraft to Launch Vehicle Interface

NASA Astrophysics Data System (ADS)

Kaufman, Daniel S.; Gordon, Scott A.

2012-07-01

For several years we had wanted to measure interface forces between a launch vehicle and the Payload. Finally in July 2006 a proposal was made and funded to evaluate the use of flight force measurements (FFM) to improve the loads process of a Spacecraft in its design and test cycle. A NASA/Industry team was formed, the core Team consisted of 20 people. The proposal identified two questions that this assessment would attempt to address by obtaining the flight forces. These questions were: 1) Is flight correlation and reconstruction with acceleration methods sufficient? 2) How much can the loads and therefore the design and qualification be reduced by having force measurements? The objective was to predict the six interface driving forces between the Spacecraft and the Launch Vehicle throughout the boost phase. Then these forces would be compared with reconstructed loads analyses for evaluation in an attempt to answer them. The paper will present the development of a strain based force measurement system and also an acceleration method, actual flight results, post flight evaluations and lessons learned.

The brain adjusts grip forces differently according to gravity and inertia: a parabolic flight experiment

PubMed Central

White, Olivier

2015-01-01

In everyday life, one of the most frequent activities involves accelerating and decelerating an object held in precision grip. In many contexts, humans scale and synchronize their grip force (GF), normal to the finger/object contact, in anticipation of the expected tangential load force (LF), resulting from the combination of the gravitational and the inertial forces. In many contexts, GF and LF are linearly coupled. A few studies have examined how we adjust the parameters–gain and offset–of this linear relationship. However, the question remains open as to how the brain adjusts GF regardless of whether LF is generated by different combinations of weight and inertia. Here, we designed conditions to generate equivalent magnitudes of LF by independently varying mass and movement frequency. In a control experiment, we directly manipulated gravity in parabolic flights, while other factors remained constant. We show with a simple computational approach that, to adjust GF, the brain is sensitive to how LFs are produced at the fingertips. This provides clear evidence that the analysis of the origin of LF is performed centrally, and not only at the periphery. PMID:25717293

The effect of aircraft control forces on pilot performance during instrument landings in a flight simulator.

PubMed

Hewson, D J; McNair, P J; Marshall, R N

2001-07-01

Pilots may have difficulty controlling aircraft at both high and low force levels due to larger variability in force production at these force levels. The aim of this study was to measure the force variability and landing performance of pilots during an instrument landing in a flight simulator. There were 12 pilots who were tested while performing 5 instrument landings in a flight simulator, each of which required different control force inputs. Pilots can produce the least force when pushing the control column to the right, therefore the force levels for the landings were set relative to each pilot's maximum aileron-right force. The force levels for the landings were 90%, 60%, and 30% of maximal aileron-right force, normal force, and 25% of normal force. Variables recorded included electromyographic activity (EMG), aircraft control forces, aircraft attitude, perceived exertion and deviation from glide slope and heading. Multivariate analysis of variance was used to test for differences between landings. Pilots were least accurate in landing performance during the landing at 90% of maximal force (p < 0.05). There was also a trend toward decreased landing performance during the landing at 25% of normal force. Pilots were more variable in force production during the landings at 60% and 90% of maximal force (p < 0.05). Pilots are less accurate at performing instrument landings when control forces are high due to the increased variability of force production. The increase in variability at high force levels is most likely associated with motor unit recruitment, rather than rate coding. Aircraft designers need to consider the reduction in pilot performance at high force levels, as well as pilot strength limits when specifying new standards.

F-4 Beryllium Rudders; A Precis of the Design, Fabrication, Ground and Flight Test Demonstrations

DTIC Science & Technology

1975-05-01

Wright-Patterson Air Force Base , Ohio 45433. AIR FORCE FLIGHT DYNAMICS LABORATORY AIR FORCE SYSTEMS COMMAND WRIGHT-PATTERSON AIR FORCE BASE , OHIO 45433...rudder. These sequential ground tests include: - A 50,000 cycle fatigue test of upper balance weight support structure. A static test to...Design Details 6. Design Analysis 7. Rudder Mass Balance 8, Rudder Moment of Inertia 9, Rudder Weight RUDDER FABRICATION AND ASSEMBLY 1. 2

Straight-line climbing flight aerodynamics of a fruit bat

NASA Astrophysics Data System (ADS)

Viswanath, K.; Nagendra, K.; Cotter, J.; Frauenthal, M.; Tafti, D. K.

2014-02-01

From flight data obtained on a fruit bat, Cynopterus brachyotis, a kinematic model for straight-line flapping motion is extracted and analyzed in a computational fluid dynamics (CFD) framework to gain insight into the complexity of bat flight. The intricate functional mechanics and architecture of the bat wings set it apart from other vertebrate flight. The extracted kinematic model is simulated for a range of Reynolds numbers, to observe the effect these phenomena have on the unsteady transient mechanisms of the flow produced by the flapping wings. The Strouhal number calculated from the data is high indicating that the oscillatory motion dominates the flow physics. From the obtained data, the bat exhibits fine control of its mechanics by actively varying wing camber, wing area, torsional rotation of the wing, forward and backward translational sweep of the wing, and wing conformation to dictate the fluid dynamics. As is common in flapping flight, the primary force generation is through the attached unsteady vortices on the wing surface. The bat through varying the wing camber and the wing area modulates this force output. The power requirement for the kinematics is analyzed and correlated with the aerodynamic performance.

X-57 Power and Command System Design

NASA Technical Reports Server (NTRS)

Clarke, Sean; Redifer, Matthew; Papathakis, Kurt; Samuel, Aamod; Foster, Trevor

2017-01-01

This paper describes the power and command system architecture of the X-57 Maxwell flight demonstrator aircraft. The X-57 is an experimental aircraft designed to demonstrate radically improved aircraft efficiency with a 3.5 times aero-propulsive efficiency gain at a "high-speed cruise" flight condition for comparable general aviation aircraft. These gains are enabled by integrating the design of a new, optimized wing and a new electric propulsion system. As a result, the X-57 vehicle takes advantage of the new capabilities afforded by electric motors as primary propulsors. Integrating new technologies into critical systems in experimental aircraft poses unique challenges that require careful design considerations across the entire vehicle system, such as qualification of new propulsors (motors, in the case of the X-57 aircraft), compatibility of existing systems with a new electric power distribution bus, and instrumentation and monitoring of newly qualified propulsion system devices.

Flexible body stability analysis of Space Shuttle ascent flight control system by using lambda matrix solution techniques

NASA Technical Reports Server (NTRS)

Bown, R. L.; Christofferson, A.; Lardas, M.; Flanders, H.

1980-01-01

A lambda matrix solution technique is being developed to perform an open loop frequency analysis of a high order dynamic system. The procedure evaluates the right and left latent vectors corresponding to the respective latent roots. The latent vectors are used to evaluate the partial fraction expansion formulation required to compute the flexible body open loop feedback gains for the Space Shuttle Digital Ascent Flight Control System. The algorithm is in the final stages of development and will be used to insure that the feedback gains meet the design specification.

X-33 Ascent Flight Controller Design by Trajectory Linearization: A Singular Perturbational Approach

NASA Technical Reports Server (NTRS)

Zhu, J. Jim; Banker, Brad D.; Hall, Charles E.

2000-01-01

The flight control of X-33 poses a challenge to conventional gain-scheduled flight controllers due to its large attitude maneuvers from liftoff to orbit and reentry. In addition, a wide range of uncertainties in vehicle handling qualities and disturbances must be accommodated by the attitude control system. Nonlinear tracking and decoupling control by trajectory linearization can be viewed as the ideal gain-scheduling controller designed at every point on the flight trajectory. Therefore it provides robust stability and performance at all stages of flight without interpolation of controller gains and eliminates costly controller redesigns due to minor airframe alteration or mission reconfiguration. In this paper, a prototype trajectory linearization design for an X-33 ascent flight controller is presented along with 3-DOF and 6-DOF simulation results. It is noted that the 6-DOF results were obtained from the 3-DOF design with only a few hours of tuning, which demonstrates the inherent robustness of the design technique. It is this "plug-and-play" feature that is much needed by NASA for the development, test and routine operations of the RLV'S. Plans for further research are also presented, and refined 6-DOF simulation results will be presented in the final version of the paper.

A neural based intelligent flight control system for the NASA F-15 flight research aircraft

NASA Technical Reports Server (NTRS)

Urnes, James M.; Hoy, Stephen E.; Ladage, Robert N.; Stewart, James

1993-01-01

A flight control concept that can identify aircraft stability properties and continually optimize the aircraft flying qualities has been developed by McDonnell Aircraft Company under a contract with the NASA-Dryden Flight Research Facility. This flight concept, termed the Intelligent Flight Control System, utilizes Neural Network technology to identify the host aircraft stability and control properties during flight, and use this information to design on-line the control system feedback gains to provide continuous optimum flight response. This self-repairing capability can provide high performance flight maneuvering response throughout large flight envelopes, such as needed for the National Aerospace Plane. Moreover, achieving this response early in the vehicle's development schedule will save cost.

Design and Flight Evaluation of a New Force-Based Flow Angle Probe

NASA Technical Reports Server (NTRS)

Corda, Stephen; Vachon, Michael Jacob

2006-01-01

A novel force-based flow angle probe was designed and flight tested on the NASA F-15B Research Testbed aircraft at NASA Dryden Flight Research Center. The prototype flow angle probe is a small, aerodynamic fin that has no moving parts. Forces on the prototype flow angle probe are measured with strain gages and correlated with the local flow angle. The flow angle probe may provide greater simplicity, greater robustness, and better access to flow measurements in confined areas relative to conventional moving vane-type flow angle probes. Flight test data were obtained at subsonic, transonic, and supersonic Mach numbers to a maximum of Mach 1.70. Flight conditions included takeoff, landing, straight and level flight, flight at higher aircraft angles of attack, and flight at elevated g-loadings. Flight test maneuvers included angle-of-attack and angle-of-sideslip sweeps. The flow angle probe-derived flow angles are compared with those obtained with a conventional moving vane probe. The flight tests validated the feasibility of a force-based flow angle measurement system.

MIDAS: Lessons learned from the first spaceborne atomic force microscope

NASA Astrophysics Data System (ADS)

Bentley, Mark Stephen; Arends, Herman; Butler, Bart; Gavira, Jose; Jeszenszky, Harald; Mannel, Thurid; Romstedt, Jens; Schmied, Roland; Torkar, Klaus

2016-08-01

The Micro-Imaging Dust Analysis System (MIDAS) atomic force microscope (AFM) onboard the Rosetta orbiter was the first such instrument launched into space in 2004. Designed only a few years after the technique was invented, MIDAS is currently orbiting comet 67P Churyumov-Gerasimenko and producing the highest resolution 3D images of cometary dust ever made in situ. After more than a year of continuous operation much experience has been gained with this novel instrument. Coupled with operations of the Flight Spare and advances in terrestrial AFM a set of "lessons learned" has been produced, cumulating in recommendations for future spaceborne atomic force microscopes. The majority of the design could be reused as-is, or with incremental upgrades to include more modern components (e.g. the processor). Key additional recommendations are to incorporate an optical microscope to aid the search for particles and image registration, to include a variety of cantilevers (with different spring constants) and a variety of tip geometries.

YF-12C on ramp

NASA Technical Reports Server (NTRS)

1973-01-01

The so-called YF-12C on the NASA Flight Research Center ramp. Following the loss of a YF-12A in a non-fatal accident in June 1971, NASA acquired the second production SR-71A (61-7951) from the Air Force. Because the SR-71 program was shrouded in the highest secrecy, the Air Force restricted NASA to using the aircraft solely for propulsion testing with YF-12A inlets and engines. It was designated the YF-12C, and given a bogus tail number (06937). The two YF-12As in the program had actual tail numbers 06935 and 06936. The first NASA flight of the YF-12C took place on 24 May 1972. The Flight Research Center's involvement with the YF-12A, an interceptor version of the Lockheed A-12, began in 1967. Ames Research Center was interested in using wind tunnel data that had been generated at Ames under extreme secrecy. Also, the Office of Advanced Research and Technology (OART) saw the YF-12A as a means to advance high-speed technology, which would help in designing the Supersonic Transport (SST). The Air Force needed technical assistance to get the latest reconnaissance version of the A-12 family, the SR-71A, fully operational. Eventually, the Air Force offered NASA the use of two YF-12A aircraft, 60-6935 and 606936. A joint NASA-USAF program was mapped out in June 1969. NASA and Air Force technicians spent three months readying 935 for flight. On 11 December 1969, the flight program got underway with a successful maiden flight piloted by Col. Joe Rogers and Maj. Gary Heidelbaugh of the SR-71/F-12 Test Force. During the program, the Air Force concentrated on military applications, and NASA pursued a loads research program. NASA studies included inflight heating, skin-friction cooling, 'coldwall' research (a heat transfer experiment), flowfield studies, shaker vane research, and tests in support of the Space Shuttle landing program. Ultimately, 935 became the workhorse of the program, with 146 flights between 11 December 1969 and 7 November 1979. The second YF-12A, 936, made 62 flights. It was lost in a non-fatal crash on 24 June 1971. It was replaced by the YF-12C. The Lockheed A-12 family, known as the Blackbirds, were designed by Clarence 'Kelly' Johnson. They were constructed mostly of titanium to withstand aerodynamic heating. Fueled by JP-7, the Blackbirds were capable of cruising at Mach 3.2 and attaining altitudes in excess of 80,000 feet. The first version, a CIA reconnaissance aircraft that first flew in April 1962 was called the A-12. An interceptor version was developed in 1963 under the designation YF-12A. A USAF reconnaissance variant, called the SR-71, was first flown in 1964. The A-12 and SR-71 designs included leading and trailing edges made of high-temperature fiberglass-asbestos laminates. The NASA YF-12 research program was ambitious; the aircraft flew an average of once a week unless down for extended maintenance or modification. Program expenses averaged $3.1 million per year just to run the flight tests. NASA crews for the YF-12 included pilots Fitzhugh Fulton and Donald Mallick, anf flight test engineers Victor Horton and Ray Young. Other NASA test pilots checked out in the YF-12A included John Manke, William Dana, Gary Krier, Einar Enevoldson, Tom McMurtry, Steve Ishmael, and Michael Swann. Only Fulton, Mallick, Ray, and Horton flew the YF-12C.

Stability Result For Dynamic Inversion Devised to Control Large Flexible Aircraft

NASA Technical Reports Server (NTRS)

Gregory, Irene M.

2001-01-01

High performance aircraft of the future will be designed lighter, more maneuverable, and operate over an ever expanding flight envelope. One of the largest differences from the flight control perspective between current and future advanced aircraft is elasticity. Over the last decade, dynamic inversion methodology has gained considerable popularity in application to highly maneuverable fighter aircraft, which were treated as rigid vehicles. This paper is an initial attempt to establish global stability results for dynamic inversion methodology as applied to a large, flexible aircraft. This work builds on a previous result for rigid fighter aircraft and adds a new level of complexity that is the flexible aircraft dynamics, which cannot be ignored even in the most basic flight control. The results arise from observations of the control laws designed for a new generation of the High-Speed Civil Transport aircraft.

Limits to Open Class Performance?

NASA Technical Reports Server (NTRS)

Bowers, Albion H.

2008-01-01

This presentation discusses open or unlimited class aircraft performance limitations and design solutions. Limitations in this class of aircraft include slow climbing flight which requires low wing loading, high cruise speed which requires high wing loading, gains in induced or viscous drag alone which result in only half the gain overall and other structural problems (yaw inertia and spins, flutter and static loads integrity). Design solutions include introducing minimum induced drag for a given span (elliptical span load or winglets) and introducing minimum induced drag for a bell shaped span load. It is concluded that open class performance limits (under current rules and technologies) is very close to absolute limits, though some gains remain to be made from unexplored areas and new technologies.

The strange flight behaviour of slowly spinning soccer balls

NASA Astrophysics Data System (ADS)

Mizota, Taketo; Kurogi, Kouhei; Ohya, Yuji; Okajima, Atsushi; Naruo, Takeshi; Kawamura, Yoshiyuki

2013-05-01

The strange three-dimensional flight behaviour of slowly spinning soccer balls is one of the most interesting and unknown phenomenon associated with the trajectories of sports balls. Many spectators have experienced numerous exciting and emotional instances while observing the curious flight behaviour of these balls. We examine the aerodynamic mechanisms of erratic ball behaviours through real flight observations, unsteady force measurements and flow pattern visualisations. The strange behaviour is elucidated by the relationship between the unsteady forces on the ball and the wake flow. The irregular changes in position for twin longitudinal vortices have already been discovered in the supercritical Reynolds number region of a sphere with a smooth surface. This finding is applicable to the strange behaviour of the flight of soccer balls with this supercritical flow. The players, spectators, and television viewers will gain greater insight into the effects of soccer ball flights.

ARES I-X: The First Test Flight of a New Era

NASA Technical Reports Server (NTRS)

Smith, R. Marshall; Davis, Stephan R.; Bryant, Richard Barry; Cook, Steve

2010-01-01

On October 28th, 2009, the National Aeronautics and Space Administration (NASA) launched the Ares I-X Flight Test Vehicle (FTV) from pad 39B, providing the first set of flight test data for NASA's Ares I vehicle design team. This test was critical in providing insight into areas were significant design challenges existed. This paper discusses the objectives of the mission and how they were satisfied. It discusses the overall results of the flight test and look at the data retrieved from the flight. Ares I-X was highly instrumented with over 700 channels of Developmental Flight Instrumentation (DFI). Significant insight was gained in the areas of thrust oscillation, vibro-acoustics, predicting jet interactions and slag ejection from solid rocket systems with submerged nozzles. The paper outlines results from the Guidance Navigation & Control (GN&C), Thermal, Vibro-acoustic, Structures, Aero, Aero-Acoustic and Trajectory teams.

Modern digital flight control system design for VTOL aircraft

NASA Technical Reports Server (NTRS)

Broussard, J. R.; Berry, P. W.; Stengel, R. F.

1979-01-01

Methods for and results from the design and evaluation of a digital flight control system (DFCS) for a CH-47B helicopter are presented. The DFCS employed proportional-integral control logic to provide rapid, precise response to automatic or manual guidance commands while following conventional or spiral-descent approach paths. It contained altitude- and velocity-command modes, and it adapted to varying flight conditions through gain scheduling. Extensive use was made of linear systems analysis techniques. The DFCS was designed, using linear-optimal estimation and control theory, and the effects of gain scheduling are assessed by examination of closed-loop eigenvalues and time responses.

Status and trends in active control technology

NASA Technical Reports Server (NTRS)

Rediess, H. A.; Szalai, K. J.

1975-01-01

The emergence of highly reliable fly-by-wire flight control systems makes it possible to consider a strong reliance on automatic control systems in the design optimization of future aircraft. This design philosophy has been referred to as the control configured vehicle approach or the application of active control technology. Several studies and flight tests sponsored by the Air Force and NASA have demonstrated the potential benefits of control configured vehicles and active control technology. The present status and trends of active control technology are reviewed and the impact it will have on aircraft designs, design techniques, and the designer is predicted.

Design and Analysis of Morpheus Lander Flight Control System

NASA Technical Reports Server (NTRS)

Jang, Jiann-Woei; Yang, Lee; Fritz, Mathew; Nguyen, Louis H.; Johnson, Wyatt R.; Hart, Jeremy J.

2014-01-01

The Morpheus Lander is a vertical takeoff and landing test bed vehicle developed to demonstrate the system performance of the Guidance, Navigation and Control (GN&C) system capability for the integrated autonomous landing and hazard avoidance system hardware and software. The Morpheus flight control system design must be robust to various mission profiles. This paper presents a design methodology for employing numerical optimization to develop the Morpheus flight control system. The design objectives include attitude tracking accuracy and robust stability with respect to rigid body dynamics and propellant slosh. Under the assumption that the Morpheus time-varying dynamics and control system can be frozen over a short period of time, the flight controllers are designed to stabilize all selected frozen-time control systems in the presence of parametric uncertainty. Both control gains in the inner attitude control loop and guidance gains in the outer position control loop are designed to maximize the vehicle performance while ensuring robustness. The flight control system designs provided herein have been demonstrated to provide stable control systems in both Draper Ares Stability Analysis Tool (ASAT) and the NASA/JSC Trick-based Morpheus time domain simulation.

Highly Maneuverable Aircraft Technology (HiMAT) flight-flutter test program

NASA Technical Reports Server (NTRS)

Kehoe, M. W.

1984-01-01

The highly maneuverable aircraft technology (HiMAT) vehicle was evaluated in a joint NASA and Air Force flight test program. The HiMAT vehicle is a remotely piloted research vehicle. Its design incorporates the use of advanced composite materials in the wings, and canards for aeroelastic tailoring. A flight-flutter test program was conducted to clear a sufficient flight envelope to allow for performance, stability and control, and loads testing. Testing was accomplished with and without flight control-surface dampers. Flutter clearance of the vehicle indicated satisfactory damping and damping trends for the structural modes of the HiMAT vehicle. The data presented include frequency and damping plotted as a function of Mach number.

X-37 Flight Demonstrator Project: Capabilities for Future Space Transportation System Development

NASA Technical Reports Server (NTRS)

Dumbacher, Daniel L.

2004-01-01

The X-37 Approach and Landing Vehicle (ALTV) is an automated (unmanned) spacecraft designed to reduce technical risk in the descent and landing phases of flight. ALTV mission requirements and Orbital Vehicle (OV) technology research and development (R&D) goals are formulated to validate and mature high-payoff ground and flight technologies such as Thermal Protection Systems (TPS). It has been more than three decades since the Space Shuttle was designed and built. Real-world hardware experience gained through the multitude of X-37 Project activities has expanded both Government and industry knowledge of the challenges involved in developing new generations of spacecraft that can fulfill the Vision for Space Exploration.

Force Limited Vibration Testing

NASA Technical Reports Server (NTRS)

Scharton, Terry; Chang, Kurng Y.

2005-01-01

This slide presentation reviews the concept and applications of Force Limited Vibration Testing. The goal of vibration testing of aerospace hardware is to identify problems that would result in flight failures. The commonly used aerospace vibration tests uses artificially high shaker forces and responses at the resonance frequencies of the test item. It has become common to limit the acceleration responses in the test to those predicted for the flight. This requires an analysis of the acceleration response, and requires placing accelerometers on the test item. With the advent of piezoelectric gages it has become possible to improve vibration testing. The basic equations have are reviewed. Force limits are analogous and complementary to the acceleration specifications used in conventional vibration testing. Just as the acceleration specification is the frequency spectrum envelope of the in-flight acceleration at the interface between the test item and flight mounting structure, the force limit is the envelope of the in-flight force at the interface . In force limited vibration tests, both the acceleration and force specifications are needed, and the force specification is generally based on and proportional to the acceleration specification. Therefore, force limiting does not compensate for errors in the development of the acceleration specification, e.g., too much conservatism or the lack thereof. These errors will carry over into the force specification. Since in-flight vibratory force data are scarce, force limits are often derived from coupled system analyses and impedance information obtained from measurements or finite element models (FEM). Fortunately, data on the interface forces between systems and components are now available from system acoustic and vibration tests of development test models and from a few flight experiments. Semi-empirical methods of predicting force limits are currently being developed on the basis of the limited flight and system test data. A simple two degree of freedom system is shown and the governing equations for basic force limiting results for this system are reviewed. The design and results of the shuttle vibration forces (SVF) experiments are reviewed. The Advanced Composition Explorer (ACE) also was used to validate force limiting. Test instrumentation and supporting equipment are reviewed including piezo-electric force transducers, signal processing and conditioning systems, test fixtures, and vibration controller systems. Several examples of force limited vibration testing are presented with some results.

Evolution of Space Shuttle Range Safety Ascent Flight Envelope Design

NASA Technical Reports Server (NTRS)

Brewer, Joan; Davis, Jerel; Glenn, Christopher

2011-01-01

For every space vehicle launch from the Eastern Range in Florida, the range user must provide specific Range Safety (RS) data products to the Air Force's 45th Space Wing in order to obtain flight plan approval. One of these data products is a set of RS ascent flight envelope trajectories that define the normal operating region of the vehicle during powered flight. With the Shuttle Program launching 135 manned missions over a 30-year period, 135 envelope sets were delivered to the range. During this time, the envelope methodology and design process evolved to support mission changes, maintain high data quality, and reduce costs. The purpose of this document is to outline the shuttle envelope design evolution and capture the lessons learned that could apply to future spaceflight endeavors.

A high speed and high gain CMOS receiver chip for a pulsed time-of-flight laser rangefinder

NASA Astrophysics Data System (ADS)

Yu, Jin-jin; Deng, Ruo-han; Yuan, Hong-hui; Chen, Yong-ping

2011-06-01

An integrated receiver channel for a pulsed time-of-flight (TOF) laser rangefinder has been designed. Pulsed TOF laser range finding devices using a laser diode transmitter can achieve millimeter-level distance measurement accuracy in a measurement range of several tens of meters to non-cooperative targets. The amplifier exploits the regulated cascade (RGC) configuration as the input-stage, thus achieving as large effective input trans-conductance as that of Si Bipolar or GaAs MESFET. The RGC input configuration isolates the input parasitic capacitance including photodiode capacitance from the bandwidth determination better than common-gate TIA. To enlarge the bandwidth, inductive peaking technology has been adopted. An active inductor (MOS-L) is used instead of spiral inductor in CMOS process. An R-2R resistor ladder is inserting between per-amplifier and post-amplifier as the variable attenuator for digital gain control purpose. The gain-bandwidth of a basic differential pair with resistive load is not large enough for broad band operation. A circuit solution to improve both gain and bandwidth of an amplifying stage is proposed. Traditional and modified Cherry-Hooper amplifiers are discussed and the cascading of several stages to constitute the post-amplifier is designed. The fully integrated one-chip solution is designed with Cadence IC design platform. The simulation result shows the bandwidth of the trans-impedance amplifier is 215MHz with the presence of a 2pF input capacitor and 5pF load capacitor. And the maximum trans-impedance gain is 136dB. The walk error is less than 1ns in 1:1000 dynamic range. The responsive time is less than 2.2ns.

The F-15B Propulsion Flight Test Fixture: A New Flight Facility For Propulsion Research

NASA Technical Reports Server (NTRS)

Corda, Stephen; Vachon, M. Jake; Palumbo, Nathan; Diebler, Corey; Tseng, Ting; Ginn, Anthony; Richwine, David

2001-01-01

The design and development of the F-15B Propulsion Flight Test Fixture (PFTF), a new facility for propulsion flight research, is described. Mounted underneath an F-15B fuselage, the PFTF provides volume for experiment systems and attachment points for propulsion devices. A unique feature of the PFTF is the incorporation of a six-degree-of-freedom force balance. Three-axis forces and moments can be measured in flight for experiments mounted to the force balance. The NASA F-15B airplane is described, including its performance and capabilities as a research test bed aircraft. The detailed description of the PFTF includes the geometry, internal layout and volume, force-balance operation, available instrumentation, and allowable experiment size and weight. The aerodynamic, stability and control, and structural designs of the PFTF are discussed, including results from aerodynamic computational fluid dynamic calculations and structural analyses. Details of current and future propulsion flight experiments are discussed. Information about the integration of propulsion flight experiments is provided for the potential PFTF user.

Computational analysis of forebody tangential slot blowing on the high alpha research vehicle

NASA Technical Reports Server (NTRS)

Gee, Ken

1994-01-01

Current and future fighter aircraft can maneuver in the high-angle-of-attack flight regime while flying at low subsonic and transonic freestream Mach numbers. However, at any flight speed, the ability of the vertical tails to generate yawing moment is limited in high-angle-of-attack flight. Thus, any system designed to provide the pilot with additional side force and yawing moment must work in both low subsonic and transonic flight. However, previous investigations of the effectiveness of forebody tangential slot blowing in generating the desired control forces and moments have been limited to the low subsonic freestream flow regime. In order to investigate the effectiveness of tangential slot blowing in transonic flight, a computational fluid dynamics analysis was carried out during the grant period. Computational solutions were obtained at three different freestream Mach numbers and at various jet mass flow ratios. All results were obtained using the isolated F/A-18 forebody grid geometry at 30.3 degrees angle of attack. One goal of the research was to determine the effect of freestream Mach number on the effectiveness of forebody tangential slot blowing in generating yawing moment. The second part of the research studied the force onset time lag associated with blowing. The time required for the yawing moment to reach a steady-state value from the onset of blowing may have an impact on the implementation of a pneumatic system on a flight vehicle.

Results of the first complete static calibration of the RSRA rotor-load-measurement system

NASA Technical Reports Server (NTRS)

Acree, C. W., Jr.

1984-01-01

The compound Rotor System Research Aircraft (RSRA) is designed to make high-accuracy, simultaneous measurements of all rotor forces and moments in flight. Physical calibration of the rotor force- and moment-measurement system when installed in the aircraft is required to account for known errors and to ensure that measurement-system accuracy is traceable to the National Bureau of Standards. The first static calibration and associated analysis have been completed with good results. Hysteresis was a potential cause of static calibration errors, but was found to be negligible in flight compared to full-scale loads, and analytical methods have been devised to eliminate hysteresis effects on calibration data. Flight tests confirmed that the calibrated rotor-load-measurement system performs as expected in flight and that it can dependably make direct measurements of fuselage vertical drag in hover.

Impact of aeroelasticity on propulsion and longitudinal flight dynamics of an air-breathing hypersonic vehicle

NASA Technical Reports Server (NTRS)

Raney, David L.; Mcminn, John D.; Pototzky, Anthony S.; Wooley, Christine L.

1993-01-01

Many air-breathing hypersonic aerospacecraft design concepts incorporate an elongated fuselage forebody acting as the aerodynamic compression surface for a hypersonic combustion module, or scram jet. This highly integrated design approach creates the potential for an unprecedented form of aero-propulsive-elastic interaction in which deflections of the vehicle fuselage give rise to propulsion transients, producing force and moment variations that may adversely impact the rigid body flight dynamics and/or further excite the fuselage bending modes. To investigate the potential for such interactions, a math model was developed which included the longitudinal flight dynamics, propulsion system, and first seven elastic modes of a hypersonic air-breathing vehicle. Perturbation time histories from a simulation incorporating this math model are presented that quantify the propulsive force and moment variations resulting from aeroelastic vehicle deflections. Root locus plots are presented to illustrate the effect of feeding the propulsive perturbations back into the aeroelastic model. A concluding section summarizes the implications of the observed effects for highly integrated hypersonic air-breathing vehicle concepts.

Impact of aeroelasticity on propulsion and longitudinal flight dynamics of an air-breathing hypersonic vehicle

NASA Astrophysics Data System (ADS)

Raney, David L.; McMinn, John D.; Pototzky, Anthony S.; Wooley, Christine L.

1993-04-01

Many air-breathing hypersonic aerospacecraft design concepts incorporate an elongated fuselage forebody acting as the aerodynamic compression surface for a hypersonic combustion module, or scram jet. This highly integrated design approach creates the potential for an unprecedented form of aero-propulsive-elastic interaction in which deflections of the vehicle fuselage give rise to propulsion transients, producing force and moment variations that may adversely impact the rigid body flight dynamics and/or further excite the fuselage bending modes. To investigate the potential for such interactions, a math model was developed which included the longitudinal flight dynamics, propulsion system, and first seven elastic modes of a hypersonic air-breathing vehicle. Perturbation time histories from a simulation incorporating this math model are presented that quantify the propulsive force and moment variations resulting from aeroelastic vehicle deflections. Root locus plots are presented to illustrate the effect of feeding the propulsive perturbations back into the aeroelastic model. A concluding section summarizes the implications of the observed effects for highly integrated hypersonic air-breathing vehicle concepts.

Ares I Flight Control System Design

NASA Technical Reports Server (NTRS)

Jang, Jiann-Woei; Alaniz, Abran; Hall, Robert; Bedrossian, Nazareth; Hall, Charles; Ryan, Stephen; Jackson, Mark

2010-01-01

The Ares I launch vehicle represents a challenging flex-body structural environment for flight control system design. This paper presents a design methodology for employing numerical optimization to develop the Ares I flight control system. The design objectives include attitude tracking accuracy and robust stability with respect to rigid body dynamics, propellant slosh, and flex. Under the assumption that the Ares I time-varying dynamics and control system can be frozen over a short period of time, the flight controllers are designed to stabilize all selected frozen-time launch control systems in the presence of parametric uncertainty. Flex filters in the flight control system are designed to minimize the flex components in the error signals before they are sent to the attitude controller. To ensure adequate response to guidance command, step response specifications are introduced as constraints in the optimization problem. Imposing these constraints minimizes performance degradation caused by the addition of the flex filters. The first stage bending filter design achieves stability by adding lag to the first structural frequency to phase stabilize the first flex mode while gain stabilizing the higher modes. The upper stage bending filter design gain stabilizes all the flex bending modes. The flight control system designs provided here have been demonstrated to provide stable first and second stage control systems in both Draper Ares Stability Analysis Tool (ASAT) and the MSFC 6DOF nonlinear time domain simulation.

User Selection Criteria of Airspace Designs in Flexible Airspace Management

NASA Technical Reports Server (NTRS)

Lee, Hwasoo E.; Lee, Paul U.; Jung, Jaewoo; Lai, Chok Fung

2011-01-01

A method for identifying global aerodynamic models from flight data in an efficient manner is explained and demonstrated. A novel experiment design technique was used to obtain dynamic flight data over a range of flight conditions with a single flight maneuver. Multivariate polynomials and polynomial splines were used with orthogonalization techniques and statistical modeling metrics to synthesize global nonlinear aerodynamic models directly and completely from flight data alone. Simulation data and flight data from a subscale twin-engine jet transport aircraft were used to demonstrate the techniques. Results showed that global multivariate nonlinear aerodynamic dependencies could be accurately identified using flight data from a single maneuver. Flight-derived global aerodynamic model structures, model parameter estimates, and associated uncertainties were provided for all six nondimensional force and moment coefficients for the test aircraft. These models were combined with a propulsion model identified from engine ground test data to produce a high-fidelity nonlinear flight simulation very efficiently. Prediction testing using a multi-axis maneuver showed that the identified global model accurately predicted aircraft responses.

YF-12 in flight

NASA Technical Reports Server (NTRS)

1972-01-01

The Flight Research Center's involvement with the YF-12A, an interceptor version of the Lockheed A-12, began in 1967. Ames Research Center was interested in using wind tunnel data that had been generated at Ames under extreme secrecy. Also, the Office of Advanced Research and Technology (OART) saw the YF-12A as a means to advance high-speed technology, which would help in designing the Supersonic Transport (SST). The Air Force needed technical assistance to get the latest reconnaissance version of the A-12 family, the SR-71A, fully operational. Eventually, the Air Force offered NASA the use of two YF-12A aircraft, 60-6935 and 60-6936. A joint NASA-USAF program was mapped out in June 1969. NASA and Air Force technicians spent three months readying 935 for flight. On 11 December 1969, the flight program got underway with a successful maiden flight piloted by Col. Joe Rogers and Maj. Gary Heidelbaugh of the SR-71/F-12 Test Force. During the program, the Air Force concentrated on military applications, and NASA pursued a loads research program. NASA studies included inflight heating, skin-friction cooling, 'coldwall' research (a heat transfer experiment), flowfield studies, shaker vane research, and tests in support of the Space Shuttle landing program. Ultimately, 935 became the workhorse of the program, with 146 flights between 11 December 1969 and 7 November 1979. The second YF-12A, 936, made 62 flights. It was lost in a non-fatal crash on 24 June 1971. It was replaced by the so-called YF-12C (SR-71A 61-7951, modified with YF-12A inlets and engines and a bogus tail number 06937). The Lockheed A-12 family, known as the Blackbirds, were designed by Clarence 'Kelly' Johnson. They were constructed mostly of titanium to withstand aerodynamic heating. Fueled by JP-7, the Blackbirds were capable of cruising at Mach 3.2 and attaining altitudes in excess of 80,000 feet. The first version, a CIA reconnaissance aircraft that first flew in April 1962 was called the A-12. An interceptor version was developed in 1963 under the designation YF-12A. A USAF reconnaissance variant, called the SR-71, was first flown in 1964. The A-12 and SR-71 designs included leading and trailing edges made of high-temperature fiberglass-asbestos laminates. The NASA YF-12 research program was ambitious; the aircraft flew an average of once a week unless down for extended maintenance or modification. Program expenses averaged $3.1 million per year just to run the flight tests. NASA crews for the YF-12 included pilots Fitzhugh Fulton and Donald Mallick, anf flight test engineers Victor Horton and Ray Young. Other NASA test pilots checked out in the YF-12A included John Manke, William Dana, Gary Krier, Einar Enevoldson, Tom McMurtry, Steve Ishmael, and Michael Swann. Except for its delivery and departure, the YF-12C was only flown by Fulton, Mallick, Horton, and Ray. It made 90 flights between 16 July 1971 and 22 December 1978.

Bi-directional thruster development and test report

NASA Technical Reports Server (NTRS)

Jacot, A. D.; Bushnell, G. S.; Anderson, T. M.

1990-01-01

The design, calibration and testing of a cold gas, bi-directional throttlable thruster are discussed. The thruster consists of an electro-pneumatic servovalve exhausting through opposite nozzles with a high gain pressure feedback loop to optimize performance. The thruster force was measured to determine hysteresis and linearity. Integral gain was used to maximize performance for linearity, hysteresis, and minimum thrust requirements. Proportional gain provided high dynamic response (bandwidth and phase lag). Thruster performance is very important since the thrusters are intended to be used for active control.

Analytical modeling and experimental evaluation of a passively morphing ornithopter wing

NASA Astrophysics Data System (ADS)

Wissa, Aimy A.

Ornithopters or flapping wing Unmanned Aerial Vehicles (UAVs) have potential applications in both civil and military sectors. Amongst all categories of UAVs, ornithopters have a unique ability to fly in low Reynolds number flight regimes and have the agility and maneuverability of rotary wing aircraft. In nature, birds achieve such performance by exploiting various wing kinematics known as gaits. The objective of this work was to improve the steady level flight wing performance of an ornithopter by implementing the Continuous Vortex Gait (CVG) using a novel passive compliant spine. The CVG is a set of bio-inspired kinematics that natural flyers use to produce lift and thrust during steady level flight. A significant contribution of this work was the recognition that the CVG is an avian gait that could be achieved using a passive morphing mechanism. In contrast to rigid-link mechanisms and active approaches, reported by other researchers in the open literature, passive morphing mechanisms require no additional energy expenditure, while introducing minimal weight addition and complexity. During the execution of the CVG, the avian wing wrist is the primary joint responsible for the wing shape changes. Thus a compliant mechanism, called a compliant spine, was fabricated, and integrated in the ornithopter's wing leading edge spar where an avian wrist would normally exist, namely at 37% of the wing half span. Each compliant spine was designed to be flexible in bending during the wing upstroke and stiff in bending during the wing downstroke. Inserting a variable stiffness compliant mechanism in the leading edge (LE) spar of the ornithopter could affect its structural stability. An analytical model was developed to determine the structural stability of the ornithopter LE spar. The model was validated using experimental measurements. The LE spar equations of motion were then reformulated into Mathieu's equation and the LE spar was proven to be structurally stable with a compliant spine design insert. A research ornithopter platform was tested in air and in vacuum as well as in free and constrained flight with various compliant spine designs inserted in its wings. Results from the constrained flight tests indicated that the ornithopter with a compliant spine inserted in its wings consumed 45% less electrical power and produced 16% of its weight in additional lift, without incurring any thrust penalties. Results from, the vacuum constrained tests attributed these benefits to aerodynamic effects rather than inertial effects. Free flight tests were performed at Wright Patterson Air Force Base, which houses the largest indoor flight laboratory in the country. The wing kinematics along with the vehicle dynamics were captured during this testing using ViconRTM motion tracking cameras. These flight tests proved to be successful in producing consistent and repeatable flight data over more than eight free flight flapping cycles of free flight and validated a new and novel testing technique. The ornithopter body dynamics were shown to be significant, i.e. +/-4gs. Inserting the compliant spine into the leading edge spar of the ornithopter during free flight reduced the baseline configuration body vertical center of mass positive acceleration by 69%, which translates into overall lift gains. It also increased the horizontal propulsive force by 300%, which translates into thrust gains.

Laboratory outreach: student assessment of flow cytometer fluidics in zero gravity.

PubMed

Crucian, B; Norman, J; Brentz, J; Pietrzyk, R; Sams, C

2000-10-01

Due to the the clinical utility of the flow cytometer, the National Aeronautics and Space Administration (NASA) is interested in the design of a space flight-compatible cytometer for use on long-duration space missions. Because fluid behavior is altered dramatically during space flight, it was deemed necessary to validate the principles of hydrodynamic focusing and laminar flow (cytometer fluidics) in a true microgravity environment. An experiment to validate these properties was conducted by 12 students from Sweetwater High School (Sweetwater, TX) participating in the NASA Reduced Gravity Student Flight Opportunity, Class of 2000. This program allows high school students to gain scientific experience by conducting an experiment on the NASA KC-135 zero gravity laboratory aircraft. The KC-135 creates actual zero-gravity conditions in 30-second intervals by flying a highly inclined parabolic flight path. The experiment was designed by their mentor in the program, the Johnson Space Center's flow cytometrist Brian Crucian, PhD, MT(ASCP). The students performed the experiment, with the mentor, onboard the NASA zero-gravity research aircraft in April 2000.

Aircraft flight flutter testing at the NASA Ames-Dryden Flight Research Facility

NASA Technical Reports Server (NTRS)

Kehoe, Michael W.

1988-01-01

Many parameter identification techniques have been used at the NASA Ames Research Center, Dryden Research Facility at Edwards Air Force Base to determine the aeroelastic stability of new and modified research vehicles in flight. This paper presents a summary of each technique used with emphasis on fast Fourier transform methods. Experiences gained from application of these techniques to various flight test programs are discussed. Also presented are data-smoothing techniques used for test data distorted by noise. Data are presented for various aircraft to demonstrate the accuracy of each parameter identification technique discussed.

Design and control of a vertical takeoff and landing fixed-wing unmanned aerial vehicle

NASA Astrophysics Data System (ADS)

Malang, Yasir

With the goal of extending capabilities of multi-rotor unmanned aerial vehicles (UAVs) for wetland conservation missions, a novel hybrid aircraft design consisting of four tilting rotors and a fixed wing is designed and built. The tilting rotors and nonlinear aerodynamic effects introduce a control challenge for autonomous flight, and the research focus is to develop and validate an autonomous transition flight controller. The overall controller structure consists of separate cascaded Proportional Integral Derivative (PID) controllers whose gains are scheduled according to the rotors' tilt angle. A control mechanism effectiveness factor is used to mix the multi-rotor and fixed-wing control actuators during transition. A nonlinear flight dynamics model is created and transition stability is shown through MATLAB simulations, which proves gain-scheduled control is a good fit for tilt-rotor aircraft. Experiments carried out using the prototype UAV validate simulation results for VTOL and tilted-rotor flight.

NACA Aircraft in hangar 1953 - L-R: Three D-558-2s, D-558-1, B-47, wing of YF-84A, background are th

NASA Technical Reports Server (NTRS)

1953-01-01

The aircraft in this 1953 photo of the National Advisory Committee for Aeronautics (NACA) hangar at South Base of Edwards Air Force Base showed the wide range of research activities being undertaken. On the left side of the hanger are the three D-558-2 research aircraft. These were designed to test swept wings at supersonic speeds approaching Mach 2. The front D-558-2 is the third built (NACA 145/Navy 37975). It has been modified with a leading-edge chord extension. This was one of a number of wing modifications, using different configurations of slats and/or wing fences, to ease the airplane's tendency to pitch-up. NACA 145 had both a jet and a rocket engine. The middle aircraft is NACA 144 (Navy 37974), the second built. It was all-rocket powered, and Scott Crossfield made the first Mach 2 flight in this aircraft on November 20, 1953. The aircraft in the back is D-558-2 number 1. NACA 143 (Navy 37973) was also carried both a jet and a rocket engine in 1953. It had been used for the Douglas contractor flights, then was turned over to the NACA. The aircraft was not converted to all-rocket power until June 1954. It made only a single NACA flight before NACA's D-558-2 program ended in 1956. Beside the three D-558-2s is the third D-558-1. Unlike the supersonic D-558-2s, it was designed for flight research at transonic speeds, up to Mach 1. The D-558-1 was jet-powered, and took off from the ground. The D-558-1's handling was poor as it approached Mach 1. Given the designation NACA 142 (Navy 37972), it made a total of 78 research flights, with the last in June 1953. In the back of the hangar is the X-4 (Air Force 46-677). This was a Northrop-built research aircraft which tested a swept wing design without horizontal stabilizers. The aircraft proved unstable in flight at speeds above Mach 0.88. The aircraft showed combined pitching, rolling, and yawing motions, and the design was considered unsuitable. The aircraft, the second X-4 built, was then used as a pilot trainer for approach and landing studies. This data was used in designing later rocket-powered aircraft. Almost hidden in the back of the hangar is the ETF-51D (NACA 148/Air Force 44-84958). This two-seat trainer was used as a low-speed chase aircraft, as well as for support flights and liaison missions. It arrived at the NACA High Speed Flight Research Station on September 5, 1950, and was retired from NASA service after a taxi accident on April 15, 1959. The U.S. Army unit at Edwards AFB repaired the aircraft, and used it for helicopter chase work. On the right side of the photo the B-47A (NACA 150/Air Force 49-1900) and YF-84A (NACA 134/Air Force 45-59490) are visible. The B-47A was the first production aircraft built by Boeing. The aircraft was transferred from Langley Memorial Aeronautical Laboratory to the High-Speed Flight Research Station on March 17, 1953, where it was used for a wide range of research, including handling qualities, dynamic stability, gust loads, noise level measurements, aeroelasticity (the bending of the wings in flight), and a survey of the X-15 High Range. The YF-84A, in front of the B-47A, was used for vortex generator studies. The Dryden Flight Research Center, NASA's premier installation for aeronautical flight research, celebrated its 50th anniversary in 1996. Dryden is the 'Center of Excellence' for atmospheric flight operations. The Center's charter is to research, develop, verify, and transfer advanced aeronautics, space, and related technologies. It is located at Edwards, Calif., on the western edge of the Mojave Desert, 80 miles north of Los Angeles. Dryden's history dates back to the early fall of 1946, when a group of five aeronautical engineers arrived at what is now Edwards from the NACA's Langley Memorial Aeronautical Laboratory, Hampton, Va. Their goal was to prepare for the X-l supersonic research flights in a joint NACA-U.S. Army Air Forces-Bell Aircraft Corp. program. NACA--the National Advisory Committee for Aeronautics--was the predecessor of today's NASA. Since the days of the X-l, the first aircraft to fly faster than the speed of sound, the installation has grown in size and significance and is associated with many important developments in aviation -- supersonic and hypersonic flight, wingless lifting bodies, digital fly-by-wire, supercritical and forward-swept wings, and the space shuttles. Its name has changed many times over the years. From 14 November 1949 to 1 July 1954 it bore the name NACA High-Speed Flight Research Station.

The chocolate-colored expanse of Rogers Dry Lake frames the sleek lines of the Boeing / NASA X-48B subscale demonstrator during a test flight at Edwards AFB

NASA Image and Video Library

2007-08-14

Boeing Phantom Works' subscale Blended Wing Body technology demonstration aircraft began its initial flight tests from NASA's Dryden Flight Research Center at Edwards Air Force Base, Calif. in the summer of 2007. The 8.5 percent dynamically scaled unmanned aircraft, designated the X-48B by the Air Force, is designed to mimic the aerodynamic characteristics of a full-scale large cargo transport aircraft with the same blended wing body shape. The initial flight tests focused on evaluation of the X-48B's low-speed flight characteristics and handling qualities. About 25 flights were planned to gather data in these low-speed flight regimes. Based on the results of the initial flight test series, a second set of flight tests was planned to test the aircraft's low-noise and handling characteristics at transonic speeds.

Flight Test Analysis of the Forces and Moments Imparted on a B737-100 Airplane During Wake Vortex Encounters

NASA Technical Reports Server (NTRS)

Roberts, Chistopher L.

2001-01-01

Aircraft travel has become a major form of transportation. Several of our major airports are operating near their capacity limit, increasing congestion and delays for travelers. As a result, the National Aeronautics and Space Administration (NASA) has been working in conjunction with the Federal Aviation Administration (FAA), airline operators, and the airline industry to increase airport capacity without sacrificing public safety. One solution to the problem is to increase the number of airports and build new. runways; yet, this solution is becoming increasingly difficult due to limited space. A better solution is to increase the production per runway. This solution increases the possibility that one aircraft will encounter the trailing wake of another aircraft. Hazardous wake vortex encounters occur when an aircraft encounters the wake produced by a heavier aircraft. This heavy-load aircraft produces high-intensity wake turbulence that redistributes the aerodynamic loads of trailing smaller aircraft. This situation is particularly hazardous for smaller aircraft during takeoffs and landings. In order to gain a better understanding of the wake-vortex/aircraft encounter phenomena, NASA Langley Research Center conducted a series of flight tests from 1995 through 1997. These tests were designed to gather data for the development a wake encounter and wake-measurement data set with the accompanying atmospheric state information. This data set is being compiled into a database that can be used by wake vortex researchers to compare with experimental and computational results. The purpose of this research is to derive and implement a procedure for calculating the wake-vortex/aircraft interaction portion of that database by using the data recorded during those flight tests. There were three objectives to this research. Initially, the wake-induced forces and moments from each flight were analyzed based on varying flap deflection angles. The flap setting alternated between 15 and 30 degrees while the separation distance remained constant. This examination was performed to determine if increases in flap deflection would increase or decrease the effects of the wake-induced forces and moments. Next, the wake-induced forces and moments from each flight were analyzed based on separation distances of 1-3 nautical miles. In this comparison, flap deflection was held constant at 30 degrees. The purpose of this study was to determine if increased separation distances reduced the effects of the wake vortex on the aircraft. The last objective compared the wake-induced forces and moments of each flight as it executed a series of maneuvers through the wake-vortex. This analysis was conducted to examine the impact of the wake on the B737 as it traversed the wake horizontally and vertically. Results from the first analysis indicated that there was no difference in wake effect at flap deflections of 15 and 30 degrees. This conclusion is evidenced in the cases of the wake-induced sideforce, rolling moment, and yawing moment. The wake-induced lift, drag, and pitching moment cases yielded less conclusive results. The second analysis compared the wake-induced forces and moments at separation distances of 1-3 nautical miles. Results indicated that there was no significant difference in the wake-induced lift, drag, sideforce, or yawing moment coefficients. The analysis compared the wake-induced forces and moments based on different flight maneuvers. It was found that the wake-induced forces and moments had the greatest impact on out-to-in and in-to-out maneuvers.

Space Launch System Implementation of Adaptive Augmenting Control

NASA Technical Reports Server (NTRS)

Wall, John H.; Orr, Jeb S.; VanZwieten, Tannen S.

2014-01-01

Given the complex structural dynamics, challenging ascent performance requirements, and rigorous flight certification constraints owing to its manned capability, the NASA Space Launch System (SLS) launch vehicle requires a proven thrust vector control algorithm design with highly optimized parameters to provide stable and high-performance flight. On its development path to Preliminary Design Review (PDR), the SLS flight control system has been challenged by significant vehicle flexibility, aerodynamics, and sloshing propellant. While the design has been able to meet all robust stability criteria, it has done so with little excess margin. Through significant development work, an Adaptive Augmenting Control (AAC) algorithm has been shown to extend the envelope of failures and flight anomalies the SLS control system can accommodate while maintaining a direct link to flight control stability criteria such as classical gain and phase margin. In this paper, the work performed to mature the AAC algorithm as a baseline component of the SLS flight control system is presented. The progress to date has brought the algorithm design to the PDR level of maturity. The algorithm has been extended to augment the full SLS digital 3-axis autopilot, including existing load-relief elements, and the necessary steps for integration with the production flight software prototype have been implemented. Several updates which have been made to the adaptive algorithm to increase its performance, decrease its sensitivity to expected external commands, and safeguard against limitations in the digital implementation are discussed with illustrating results. Monte Carlo simulations and selected stressing case results are also shown to demonstrate the algorithm's ability to increase the robustness of the integrated SLS flight control system.

History of suborbital spaceflight: medical and performance issues.

PubMed

Campbell, Mark R; Garbino, Alejandro

2011-04-01

The development of manned sub-orbital commercial space vehicles is rapidly occurring and flight testing followed by operational flights will soon begin. The experience of manned suborbital spaceflight at the designated altitude (100 km/62.14 mi) is very limited--two Mercury-Redstone flights, two X-15 flights, one inadvertent Soyuz launch abort, and three recent SpaceShipOne flights, with only 15 min of critical flight time each. All indications were that the sequence of acceleration-weightlessness-deceleration was well tolerated with minimal neurovestibular dysfunction. However, there are some indications that distraction and spatial disorientation did occur. Vertigo on transition from the boost phase to weightlessness was reported on most high-altitude X-15 flights. +Gz tolerance to re-entry deceleration forces (as high as 6 + Gz) after 4 min of weightlessness is still unknown. Only further suborbital spaceflight experience will clarify if pilot performance will be affected.

F-8C adaptive flight control extensions. [for maximum likelihood estimation

NASA Technical Reports Server (NTRS)

Stein, G.; Hartmann, G. L.

1977-01-01

An adaptive concept which combines gain-scheduled control laws with explicit maximum likelihood estimation (MLE) identification to provide the scheduling values is described. The MLE algorithm was improved by incorporating attitude data, estimating gust statistics for setting filter gains, and improving parameter tracking during changing flight conditions. A lateral MLE algorithm was designed to improve true air speed and angle of attack estimates during lateral maneuvers. Relationships between the pitch axis sensors inherent in the MLE design were examined and used for sensor failure detection. Design details and simulation performance are presented for each of the three areas investigated.

Sliding Mode Control of the X-33 with an Engine Failure

NASA Technical Reports Server (NTRS)

Shtessel, Yuri B.; Hall, Charles E.

2000-01-01

Ascent flight control of the X-3 is performed using two XRS-2200 linear aerospike engines. in addition to aerosurfaces. The baseline control algorithms are PID with gain scheduling. Flight control using an innovative method. Sliding Mode Control. is presented for nominal and engine failed modes of flight. An easy to implement, robust controller. requiring no reconfiguration or gain scheduling is demonstrated through high fidelity flight simulations. The proposed sliding mode controller utilizes a two-loop structure and provides robust. de-coupled tracking of both orientation angle command profiles and angular rate command profiles in the presence of engine failure, bounded external disturbances (wind gusts) and uncertain matrix of inertia. Sliding mode control causes the angular rate and orientation angle tracking error dynamics to be constrained to linear, de-coupled, homogeneous, and vector valued differential equations with desired eigenvalues. Conditions that restrict engine failures to robustness domain of the sliding mode controller are derived. Overall stability of a two-loop flight control system is assessed. Simulation results show that the designed controller provides robust, accurate, de-coupled tracking of the orientation angle command profiles in the presence of external disturbances and vehicle inertia uncertainties, as well as the single engine failed case. The designed robust controller will significantly reduce the time and cost associated with flying new trajectory profiles or orbits, with new payloads, and with modified vehicles

Analytical and flight investigation of the influence of rotor and other high-order dynamics on helicopter flight-control system bandwidth

NASA Technical Reports Server (NTRS)

Chen, R. T. N.; Hindson, W. S.

1985-01-01

The increasing use of highly augmented digital flight-control systems in modern military helicopters prompted an examination of the influence of rotor dynamics and other high-order dynamics on control-system performance. A study was conducted at NASA Ames Research Center to correlate theoretical predictions of feedback gain limits in the roll axis with experimental test data obtained from a variable-stability research helicopter. Feedback gains, the break frequency of the presampling sensor filter, and the computational frame time of the flight computer were systematically varied. The results, which showed excellent theoretical and experimental correlation, indicate that the rotor-dynamics, sensor-filter, and digital-data processing delays can severely limit the usable values of the roll-rate and roll-attitude feedback gains.

Design and Development of a 200-kW Turbo-Electric Distributed Propulsion Testbed

NASA Technical Reports Server (NTRS)

Papathakis, Kurt V.; Kloesel, Kurt J.; Lin, Yohan; Clarke, Sean; Ediger, Jacob J.; Ginn, Starr

2016-01-01

The National Aeronautics and Space Administration (NASA) Armstrong Flight Research Center (AFRC) (Edwards, California) is developing a Hybrid-Electric Integrated Systems Testbed (HEIST) Testbed as part of the HEIST Project, to study power management and transition complexities, modular architectures, and flight control laws for turbo-electric distributed propulsion technologies using representative hardware and piloted simulations. Capabilities are being developed to assess the flight readiness of hybrid electric and distributed electric vehicle architectures. Additionally, NASA will leverage experience gained and assets developed from HEIST to assist in flight-test proposal development, flight-test vehicle design, and evaluation of hybrid electric and distributed electric concept vehicles for flight safety. The HEIST test equipment will include three trailers supporting a distributed electric propulsion wing, a battery system and turbogenerator, dynamometers, and supporting power and communication infrastructure, all connected to the AFRC Core simulation. Plans call for 18 high performance electric motors that will be powered by batteries and the turbogenerator, and commanded by a piloted simulation. Flight control algorithms will be developed on the turbo-electric distributed propulsion system.

YF-12C in flight at sunset

NASA Technical Reports Server (NTRS)

1974-01-01

The so-called YF-12C in flight at sunset. The YF-12C was the second production SR-71A (61-7951), modified with YF-12A inlets and engines, and given a bogus tail number (06937). It replaced a YF-12A (60-6936) that crashed during a joint USAF-NASA research program. The Flight Research Center's involvement with the YF-12A, an interceptor version of the Lockheed A-12, began in 1967. Ames Research Center was interested in using wind tunnel data that had been generated at Ames under extreme secrecy. Also, the Office of Advanced Research and Technology (OART) saw the YF-12A as a means to advance high-speed technology, which would help in designing the Supersonic Transport (SST). The Air Force needed technical assistance to get the latest reconnaissance version of the A-12 family, the SR-71A, fully operational. Eventually, the Air Force offered NASA the use of two YF-12A aircraft, 60-6935 and 60-6936. A joint NASA-USAF program was mapped out in June 1969. NASA and Air Force technicians spent three months readying 935 for flight. On 11 December 1969, the flight program got underway with a successful maiden flight piloted by Col. Joe Rogers and Maj. Gary Heidelbaugh of the SR-71/F-12 Test Force. During the program, the Air Force concentrated on military applications, and NASA pursued a loads research program. NASA studies included inflight heating, skin-friction cooling, 'coldwall' research (a heat transfer experiment), flowfield studies, shaker vane research, and tests in support of the Space Shuttle landing program. Ultimately, 935 became the workhorse of the program, with 146 flights between 11 December 1969 and 7 November 1979. The second YF-12A, 936, made 62 flights. It was lost in a non-fatal crash on 24 June 1971. It was replaced by the YF-12C. The YF-12C was delivered to NASA on 16 July 1971. From then until 22 December 1978, it made 90 flights. The Lockheed A-12 family, known as the Blackbirds, were designed by Clarence 'Kelly' Johnson. They were constructed mostly of titanium to withstand aerodynamic heating. Fueled by JP-7, the Blackbirds were capable of cruising at Mach 3.2 and attaining altitudes in excess of 80,000 feet. The first version, a CIA reconnaissance aircraft that first flew in April 1962 was called the A-12. An interceptor version was developed in 1963 under the designation YF-12A. A USAF reconnaissance variant, called the SR-71, was first flown in 1964. The A-12 and SR-71 designs included leading and trailing edges made of high-temperature fiberglass-asbestos laminates. The NASA YF-12 research program was ambitious; the aircraft flew an average of once a week unless down for extended maintenance or modification. Program expenses averaged $3.1 million per year just to run the flight tests. NASA crews for the YF-12 included pilots Fitzhugh Fulton and Donald Mallick, anf flight test engineers Victor Horton and Ray Young. Other NASA test pilots checked out in the YF-12A included John Manke, William Dana, Gary Krier, Einar Enevoldson, Tom McMurtry, Steve Ishmael, and Michael Swann. The YF-12C was flown only by Fulton, Mallick, Horton and Young during its NASA research missions.

Aircraft Flutter Testing

NASA Technical Reports Server (NTRS)

1997-01-01

Wilmer Reed gained international recognition for his innovative research, contributions and patented ideas relating to flutter and aeroelasticity of aerospace vehicles at Langley Research Center. In the early 1980's, Reed retired from Langley and joined the engineering staff of Dynamic Engineering Inc. While at DEI, Reed conceived and patented the DEI Flutter Exciter, now used world-wide in flight flutter testing of new or modified aircraft designs. When activated, the DEI Flutter Exciter alternately deflects the airstream upward and downward in a rapid manner, creating a force similar to that produced by an oscillating trailing edge flap. The DEI Flutter Exciter is readily adaptable to a variety of aircraft.

Hypersonic rarefied-flow aerodynamics inferred from Shuttle Orbiter acceleration measurements

NASA Technical Reports Server (NTRS)

Blanchard, R. C.; Hinson, E. W.

1989-01-01

Data obtained from multiple flights of sensitive accelerometers on the Space Shuttle Orbiter during reentry have been used to develop an improved aerodynamic model for the Orbiter normal- and axial-force coefficients in hypersonic rarefied flow. The lack of simultaneous atmospheric density measurements was overcome in part by using the ratio of normal-to-axial acceleration, in which density cancels, as a constraint. Differences between the preflight model and the flight-acceleration-derived model in the continuum regime are attributed primarily to real gas effects. New insights are gained into the variation of the force coefficients in the transition between the continuum regime and free molecule flow.

ACSYNT inner loop flight control design study

NASA Technical Reports Server (NTRS)

Bortins, Richard; Sorensen, John A.

1993-01-01

The NASA Ames Research Center developed the Aircraft Synthesis (ACSYNT) computer program to synthesize conceptual future aircraft designs and to evaluate critical performance metrics early in the design process before significant resources are committed and cost decisions made. ACSYNT uses steady-state performance metrics, such as aircraft range, payload, and fuel consumption, and static performance metrics, such as the control authority required for the takeoff rotation and for landing with an engine out, to evaluate conceptual aircraft designs. It can also optimize designs with respect to selected criteria and constraints. Many modern aircraft have stability provided by the flight control system rather than by the airframe. This may allow the aircraft designer to increase combat agility, or decrease trim drag, for increased range and payload. This strategy requires concurrent design of the airframe and the flight control system, making trade-offs of performance and dynamics during the earliest stages of design. ACSYNT presently lacks means to implement flight control system designs but research is being done to add methods for predicting rotational degrees of freedom and control effector performance. A software module to compute and analyze the dynamics of the aircraft and to compute feedback gains and analyze closed loop dynamics is required. The data gained from these analyses can then be fed back to the aircraft design process so that the effects of the flight control system and the airframe on aircraft performance can be included as design metrics. This report presents results of a feasibility study and the initial design work to add an inner loop flight control system (ILFCS) design capability to the stability and control module in ACSYNT. The overall objective is to provide a capability for concurrent design of the aircraft and its flight control system, and enable concept designers to improve performance by exploiting the interrelationships between aircraft and flight control system design parameters.

Rapid perceptual adaptation to high gravitoinertial force levels Evidence for context-specific adaptation

NASA Technical Reports Server (NTRS)

Lackner, J. R.; Graybiel, A.

1982-01-01

Subjects exposed to periodic variations in gravitoinertial force (2-G peak) in parabolic flight maneuvers quickly come to perceive the peak force level as having decreased in intensity. By the end of a 40-parabola flight, the decrease in apparent force is approximately 40%. On successive flight days, the apparent intensity of the force loads seems to decrease as well, indicating a cumulative adaptive effect. None of the subjects reported feeling abnormally 'light' for more than a minute or two after return to 1-G background force levels. The pattern of findings suggests a context-specific adaptation to high-force levels.

This photo shows a head-on view of NASA's SR-71B on the ramp at the Air Force's Plant 42 in Palmdale, California, shortly before delivery to DFRC

NASA Image and Video Library

1991-07-24

This photo shows a head-on view of NASA's SR-71B, used for pilot proficiency and training, on the ramp at the Air Force's Plant 42 in Palmdale, California, shortly before delivery to the Ames-Dryden Flight Research Facility (later, Dryden Flight Research Center) at Edwards, California. NASA operated two of these unique aircraft, an SR-71A, for high-speed, high altitude research, and this SR- 71B pilot trainer for most of the decade of the 1990s. The "B" model is special because of its raised rear cockpit, which provided a second pilot position so a trainer and an experienced pilot could both see what was going on during flights. The SR-71 was designed and built by the Lockheed Skunk Works, now the Lockheed Martin Skunk Works. Studies have shown that less than 20 percent of the total thrust used to fly at Mach 3 is produced by the basic engine itself. The balance of the total thrust is produced by the unique design of the engine inlet and "moveable spike" system at the front of the engine nacelles, and by the ejector nozzles at the exhaust which burn air compressed in the engine bypass system. Data from the SR-71 high speed research program will be used to aid designers of future supersonic/hypersonic aircraft and propulsion systems, including a high speed civil transport.

Redesign of a Variable-Gain Output Feedback Longitudinal Controller Flown on the High-Alpha Research Vehicle (HARV)

NASA Technical Reports Server (NTRS)

Ostroff, Aaron J.

1998-01-01

This paper describes a redesigned longitudinal controller that flew on the High-Alpha Research Vehicle (HARV) during calendar years (CY) 1995 and 1996. Linear models are developed for both the modified controller and a baseline controller that was flown in CY 1994. The modified controller was developed with three gain sets for flight evaluation, and several linear analysis results are shown comparing the gain sets. A Neal-Smith flying qualities analysis shows that performance for the low- and medium-gain sets is near the level 1 boundary, depending upon the bandwidth assumed, whereas the high-gain set indicates a sensitivity problem. A newly developed high-alpha Bode envelope criterion indicates that the control system gains may be slightly high, even for the low-gain set. A large motion-base simulator in the United Kingdom was used to evaluate the various controllers. Desired performance, which appeared to be satisfactory for flight, was generally met with both the low- and medium-gain sets. Both the high-gain set and the baseline controller were very sensitive, and it was easy to generate pilot-induced oscillation (PIO) in some of the target-tracking maneuvers. Flight target-tracking results varied from level 1 to level 3 and from no sensitivity to PIO. These results were related to pilot technique and whether actuator rate saturation was encountered.

Prediction of forces and moments for flight vehicle control effectors. Part 1: Validation of methods for predicting hypersonic vehicle controls forces and moments

NASA Technical Reports Server (NTRS)

Maughmer, Mark D.; Ozoroski, L.; Ozoroski, T.; Straussfogel, D.

1990-01-01

Many types of hypersonic aircraft configurations are currently being studied for feasibility of future development. Since the control of the hypersonic configurations throughout the speed range has a major impact on acceptable designs, it must be considered in the conceptual design stage. The ability of the aerodynamic analysis methods contained in an industry standard conceptual design system, APAS II, to estimate the forces and moments generated through control surface deflections from low subsonic to high hypersonic speeds is considered. Predicted control forces and moments generated by various control effectors are compared with previously published wind tunnel and flight test data for three configurations: the North American X-15, the Space Shuttle Orbiter, and a hypersonic research airplane concept. Qualitative summaries of the results are given for each longitudinal force and moment and each control derivative in the various speed ranges. Results show that all predictions of longitudinal stability and control derivatives are acceptable for use at the conceptual design stage. Results for most lateral/directional control derivatives are acceptable for conceptual design purposes; however, predictions at supersonic Mach numbers for the change in yawing moment due to aileron deflection and the change in rolling moment due to rudder deflection are found to be unacceptable. Including shielding effects in the analysis is shown to have little effect on lift and pitching moment predictions while improving drag predictions.

Interaction between graviception and carotid baroreflex function in humans during parabolic flight-induced microgravity.

PubMed

Ogoh, Shigehiko; Marais, Michaël; Lericollais, Romain; Denise, Pierre; Raven, Peter B; Normand, Hervé

2018-05-10

The aim of the present study was to assess carotid baroreflex (CBR) during acute changes in otolithic activity in humans. To address this question, we designed a set of experiments to identify the modulatory effects of microgravity on CBR function at a tilt angle of -2{degree sign}, which was identified to minimize changes in central blood volume during parabolic flight. During parabolic flight at 0g and 1g, CBR function curves were modelled from the heart rate (HR) and mean arterial pressure (MAP) responses to rapid pulse trains of neck pressure (NP) and neck suction (NS) ranging from +40 to -80 Torr; CBR control of HR (carotid-HR) and MAP (carotid-MAP) baroreflex function curves, respectively. The maximal gain (G max ) of both carotid-HR and carotid-MAP baroreflex function curves were augmented during microgravity compared to 1g (carotid-HR, -0.53 to -0.80 beats/min/mmHg, P<0.05; carotid-MAP, -0.24 to -0.30 mmHg/mmHg, P<0.05). These findings suggest that parabolic flight-induced acute change of otolithic activity may modify CBR function and identifies that the vestibular system contributes to blood pressure regulation under fluctuations in gravitational forces.

A study of helicopter stability and control including blade dynamics

NASA Technical Reports Server (NTRS)

Zhao, Xin; Curtiss, H. C., Jr.

1988-01-01

A linearized model of rotorcraft dynamics has been developed through the use of symbolic automatic equation generating techniques. The dynamic model has been formulated in a unique way such that it can be used to analyze a variety of rotor/body coupling problems including a rotor mounted on a flexible shaft with a number of modes as well as free-flight stability and control characteristics. Direct comparison of the time response to longitudinal, lateral and directional control inputs at various trim conditions shows that the linear model yields good to very good correlation with flight test. In particular it is shown that a dynamic inflow model is essential to obtain good time response correlation, especially for the hover trim condition. It also is shown that the main rotor wake interaction with the tail rotor and fixed tail surfaces is a significant contributor to the response at translational flight trim conditions. A relatively simple model for the downwash and sidewash at the tail surfaces based on flat vortex wake theory is shown to produce good agreement. Then, the influence of rotor flap and lag dynamics on automatic control systems feedback gain limitations is investigated with the model. It is shown that the blade dynamics, especially lagging dynamics, can severly limit the useable values of the feedback gain for simple feedback control and that multivariable optimal control theory is a powerful tool to design high gain augmentation control system. The frequency-shaped optimal control design can offer much better flight dynamic characteristics and a stable margin for the feedback system without need to model the lagging dynamics.

X-31 Unloading Returning from Paris Air Show

NASA Technical Reports Server (NTRS)

1995-01-01

After being flown in the Paris Air Show in June 1995, the X-31 Enhanced Fighter Maneuverability Technology Demonstrator Aircraft, based at the NASA Dryden Flight Research Center, Edwards Air Force Base, California, is off-loaded from an Air Force Reserve C-5 transport after the ferry flight back to Edwards. At the air show, the X-31 demonstrated the value of using thrust vectoring (directing engine exhaust flow) coupled with advanced flight control systems to provide controlled flight at very high angles of attack. The X-31 Enhanced Fighter Maneuverability (EFM) demonstrator flew at the Ames- Dryden Flight Research Facility, Edwards, California (redesignated the Dryden Flight Research Center in 1994) from February 1992 until 1995 and before that at the Air Force's Plant 42 in Palmdale, California. The goal of the project was to provide design information for the next generation of highly maneuverable fighter aircraft. This program demonstrated the value of using thrust vectoring (directing engine exhaust flow) coupled with an advanced flight control system to provide controlled flight to very high angles of attack. The result was a significant advantage over most conventional fighters in close-in combat situations. The X-31 flight program focused on agile flight within the post-stall regime, producing technical data to give aircraft designers a better understanding of aerodynamics, effectiveness of flight controls and thrust vectoring, and airflow phenomena at high angles of attack. Stall is a condition of an airplane or an airfoil in which lift decreases and drag increases due to the separation of airflow. Thrust vectoring compensates for the loss of control through normal aerodynamic surfaces that occurs during a stall. Post-stall refers to flying beyond the normal stall angle of attack, which in the X-31 was at a 30-degree angle of attack. During Dryden flight testing, the X-31 aircraft established several milestones. On November 6, 1992, the X-31 achieved controlled flight at a 70-degree angle of attack. On April 29, 1993, the second X-31 successfully executed a rapid minimum-radius, 180-degree turn using a post-stall maneuver, flying well beyond the aerodynamic limits of any conventional aircraft. This revolutionary maneuver has been called the 'Herbst Maneuver' after Wolfgang Herbst, a German proponent of using post-stall flight in air-to-air combat. It is also called a 'J Turn' when flown to an arbitrary heading change. The aircraft was flown in tactical maneuvers against an F/A-18 and other tactical aircraft as part of the test flight program. During November and December 1993, the X-31 reached a supersonic speed of Mach 1.28. In 1994, the X-31 program installed software to demonstrate quasi-tailless operation. The X-31 flight test program was conducted by an international test organization (ITO) managed by the Advanced Research Projects Office (ARPA), known as the Defense Advanced Research Projects Office (DARPA) before March 1993. The ITO included the U.S. Navy and U.S. Air Force, Rockwell Aerospace, the Federal Republic of Germany, Daimler-Benz (formerly Messerschmitt-Bolkow-Blohm and Deutsche Aerospace), and NASA. Gary Trippensee was the ITO director and NASA Project Manager. Pilots came from participating organizations. The X-31 was 43.33 feet long with a wingspan of 23.83 feet. It was powered by a single General Electric P404-GE-400 turbofan engine that produced 16,000 pounds of thrust in afterburner.

Design and evaluation of a deformable wing configuration for economical hovering flight of an insect-like tailless flying robot.

PubMed

Phan, Hoang Vu; Park, Hoon Cheol

2018-04-18

Studies on wing kinematics indicate that flapping insect wings operate at higher angles of attack (AoAs) than conventional rotary wings. Thus, effectively flying an insect-like flapping-wing micro air vehicle (FW-MAV) requires appropriate wing design for achieving low power consumption and high force generation. Even though theoretical studies can be performed to identify appropriate geometric AoAs for a wing for achieving efficient hovering flight, designing an actual wing by implementing these angles into a real flying robot is challenging. In this work, we investigated the wing morphology of an insect-like tailless FW-MAV, which was named KUBeetle, for obtaining high vertical force/power ratio or power loading. Several deformable wing configurations with various vein structures were designed, and their characteristics of vertical force generation and power requirement were theoretically and experimentally investigated. The results of the theoretical study based on the unsteady blade element theory (UBET) were validated with reference data to prove the accuracy of power estimation. A good agreement between estimated and measured results indicated that the proposed UBET model can be used to effectively estimate the power requirement and force generation of an FW-MAV. Among the investigated wing configurations operating at flapping frequencies of 23 Hz to 29 Hz, estimated results showed that the wing with a suitable vein placed outboard exhibited an increase of approximately 23.7%  ±  0.5% in vertical force and approximately 10.2%  ±  1.0% in force/power ratio. The estimation was supported by experimental results, which showed that the suggested wing enhanced vertical force by approximately 21.8%  ±  3.6% and force/power ratio by 6.8%  ±  1.6%. In addition, wing kinematics during flapping motion was analyzed to determine the reason for the observed improvement.

A flight simulator control system using electric torque motors

NASA Technical Reports Server (NTRS)

Musick, R. O.; Wagner, C. A.

1975-01-01

Control systems are required in flight simulators to provide representative stick and rudder pedal characteristics. A system has been developed that uses electric dc torque motors instead of the more common hydraulic actuators. The torque motor system overcomes certain disadvantages of hydraulic systems, such as high cost, high power consumption, noise, oil leaks, and safety problems. A description of the torque motor system is presented, including both electrical and mechanical design as well as performance characteristics. The system develops forces sufficiently high for most simulations, and is physically small and light enough to be used in most motion-base cockpits.

Wearable Accelerometers in High Performance Jet Aircraft.

PubMed

Rice, G Merrill; VanBrunt, Thomas B; Snider, Dallas H; Hoyt, Robert E

2016-02-01

Wearable accelerometers have become ubiquitous in the fields of exercise physiology and ambulatory hospital settings. However, these devices have yet to be validated in extreme operational environments. The objective of this study was to correlate the gravitational forces (G forces) detected by wearable accelerometers with the G forces detected by high performance aircraft. We compared the in-flight G forces detected by the two commercially available portable accelerometers to the F/A-18 Carrier Aircraft Inertial Navigation System (CAINS-2) during 20 flights performed by the Navy's Flight Demonstration Squadron (Blue Angels). Postflight questionnaires were also used to assess the perception of distractibility during flight. Of the 20 flights analyzed, 10 complete in-flight comparisons were made, accounting for 25,700 s of correlation between the CAINS-2 and the two tested accelerometers. Both accelerometers had strong correlations with that of the F/A-18 Gz axis, averaging r = 0.92 and r = 0.93, respectively, over 10 flights. Comparison of both portable accelerometer's average vector magnitude to each other yielded an average correlation of r = 0.93. Both accelerometers were found to be minimally distracting. These results suggest the use of wearable accelerometers is a valid means of detecting G forces during high performance aircraft flight. Future studies using this surrogate method of detecting accelerative forces combined with physiological information may yield valuable in-flight normative data that heretofore has been technically difficult to obtain and hence holds the promise of opening the door for a new golden age of aeromedical research.

A learning flight control system for the F8-DFBW aircraft. [Digital Fly-By-Wire

NASA Technical Reports Server (NTRS)

Montgomery, R. C.; Mekel, R.; Nachmias, S.

1978-01-01

This report contains a complete description of a learning control system designed for the F8-DFBW aircraft. The system is parameter-adaptive with the additional feature that it 'learns' the variation of the control system gains needed over the flight envelope. It, thus, generates and modifies its gain schedule when suitable data are available. The report emphasizes the novel learning features of the system: the forms of representation of the flight envelope and the process by which identified parameters are used to modify the gain schedule. It contains data taken during piloted real-time 6 degree-of-freedom simulations that were used to develop and evaluate the system.

Shuttle Radar Topography Mission (SRTM) Flight System Design and Operations Overview

NASA Technical Reports Server (NTRS)

Shen, Yuhsyen; Shaffer, Scott J.; Jordan, Rolando L.

2000-01-01

This paper provides an overview of the Shuttle Radar Topography Mission (SRTM), with emphasis on flight system implementation and mission operations from systems engineering perspective. Successfully flown in February, 2000, the SRTM's primary payload consists of several subsystems to form the first spaceborne dual-frequency (C-band and X-band) fixed baseline interferometric synthetic aperture radar (InSAR) system, with the mission objective to acquire data sets over 80% of Earth's landmass for height reconstruction. The paper provides system architecture, unique design features, engineering budgets, design verification, in-flight checkout and data acquisition of the SRTM payload, in particular for the C-band system. Mission operation and post-mission data processing activities are also presented. The complexity of the SRTM as a system, the ambitious mission objective, the demanding requirements and the high interdependency between multi-disciplined subsystems posed many challenges. The engineering experience and the insight thus gained have important implications for future spaceborne interferometric SAR mission design and implementation.

Space Launch System Implementation of Adaptive Augmenting Control

NASA Technical Reports Server (NTRS)

VanZwieten, Tannen S.; Wall, John H.; Orr, Jeb S.

2014-01-01

Given the complex structural dynamics, challenging ascent performance requirements, and rigorous flight certification constraints owing to its manned capability, the NASA Space Launch System (SLS) launch vehicle requires a proven thrust vector control algorithm design with highly optimized parameters to robustly demonstrate stable and high performance flight. On its development path to preliminary design review (PDR), the stability of the SLS flight control system has been challenged by significant vehicle flexibility, aerodynamics, and sloshing propellant dynamics. While the design has been able to meet all robust stability criteria, it has done so with little excess margin. Through significant development work, an adaptive augmenting control (AAC) algorithm previously presented by Orr and VanZwieten, has been shown to extend the envelope of failures and flight anomalies for which the SLS control system can accommodate while maintaining a direct link to flight control stability criteria (e.g. gain & phase margin). In this paper, the work performed to mature the AAC algorithm as a baseline component of the SLS flight control system is presented. The progress to date has brought the algorithm design to the PDR level of maturity. The algorithm has been extended to augment the SLS digital 3-axis autopilot, including existing load-relief elements, and necessary steps for integration with the production flight software prototype have been implemented. Several updates to the adaptive algorithm to increase its performance, decrease its sensitivity to expected external commands, and safeguard against limitations in the digital implementation are discussed with illustrating results. Monte Carlo simulations and selected stressing case results are shown to demonstrate the algorithm's ability to increase the robustness of the integrated SLS flight control system.

X-2 on ramp with B-50 mothership and support crew

NASA Technical Reports Server (NTRS)

1956-01-01

Air Force test pilot Capt. Iven Kincheloe stands in front of the Bell X-2 (46-674) on the ramp at Edwards Air Force Base, California. Behind the X-2 are ground support personnel, the B-50 launch aircraft and crew, chase planes, and support vehicles. Kincheloe had flown nearly 100 combat missions in Korea in an F-86 and was credited with shooting down 10 enemy aircraft. He then graduated from the Empire Test Pilot's School in Great Britain in December 1954, whereupon he was assigned to Edwards Air Force Base. He made four powered flights in the X-2. On September 7, 1956, he reached an altitude of 126,200 feet. After the death of Capt. Mel Apt and the loss of the X-2 #1 on September 27, 1956, in the first Mach 3 flight, Kincheloe was assigned as the Air Force project pilot for the X-15. Before he had a chance to fly that rocket-powered aircraft, Kincheloe himself lost his life on July 26, 1958, in an F-104 accident. The X-2 was a swept-wing, rocket-powered aircraft designed to fly faster than Mach 3 (three times the speed of sound). It was built for the U.S. Air Force by the Bell Aircraft Company, Buffalo, New York. The X-2 was flown to investigate the problems of aerodynamic heating as well as stability and control effectiveness at high altitudes and high speeds (in excess of Mach 3). Bell aircraft built two X-2 aircraft. These were constructed of K-monel (a copper and nickel alloy) for the fuselage and stainless steel for the swept wings and control surfaces. The aircraft had ejectable nose capsules instead of ejection seats because the development of ejection seats had not reached maturity at the time the X-2 was conceived. The X-2 ejection canopy was successfully tested using a German V-2 rocket. The X-2 used a skid-type landing gear to make room for more fuel. The airplane was air launched from a modified Boeing B-50 Superfortress Bomber. X-2 Number 1 made its first unpowered glide flight on Aug. 5, 1954, and made a total of 17 (4 glide and 13 powered) flights before it was lost Sept. 27, 1956. The pilot on Flight 17, Capt. Milburn Apt, had flown the aircraft to a record speed of Mach 3.2 (2,094 mph), thus becoming the first person to exceed Mach 3. During that last flight, inertial coupling occurred and the pilot was killed. The aircraft suffered little damage in the crash, resulting in proposals (never implemented) from the Langley Memorial Aeronautical Laboratory, Hampton, Virginia, to rebuild it for use in a hypersonic (Mach 5+) test program. In 1953, X-2 Number 2 was lost in an in-flight explosion while at the Bell Aircraft Company during captive flight trials and was jettisoned into Lake Ontario. The Air Force had previously flown the aircraft on three glide flights at Edwards Air Force Base, California, in 1952. Although the NACA's High-Speed Flight Station, Edwards, California, (predecessor of NASA's Dryden Flight Research Center) never actually flew the X-2 aircraft, the NACA did support the program primarily through Langley Memorial Aeronautical Laboratory wind-tunnel tests and Wallops Island, Virginia, rocket-model tests. The NACA High-Speed Flight Station also provided stability and control recording instrumentation and simulator support for the Air Force flights. In the latter regard, the NACA worked with the Air Force in using a special computer to extrapolate and predict aircraft behavior from flight data.

X-31 in flight - Double Reversal

NASA Technical Reports Server (NTRS)

1995-01-01

Two X-31 Enhanced Fighter Maneuverability (EFM) demonstrators were flown at the Rockwell International facility, Palmdale, California, and the NASA Dryden Flight Research Center, Edwards, California, to obtain data that may apply to the design of highly-maneuverable next-generation fighters. The program had its first flight on October 11, 1990, in Palmdale; it ended in June 1995. The X-31 program demonstrated the value of thrust vectoring (directing engine exhaust flow) coupled with advanced flight control systems, to provide controlled flight during close-in air combat at very high angles of attack. The result of this increased maneuverability is an airplane with a significant advantage over conventional fighters. 'Angle-of-attack' (alpha) is an engineering term to describe the angle of an aircraft body and wings relative to its actual flight path. During maneuvers, pilots often fly at extreme angles of attack -- with the nose pitched up while he aircraft continues in its original direction. This can lead to loss of control and result in the loss of the aircraft, pilot or both. Three thrust-vectoring paddles made of graphite epoxy mounted on the exhaust nozzle of the X-31 aircraft directed the exhaust flow to provide control in pitch (up and down) and yaw (right and left) to improve control. The paddles can sustain heat of up to 1,500 degrees centigrade for extended periods of time. In addition the X-31 aircraft were configured with movable forward canards and fixed aft strakes. The canards were small wing-like structures set on the wing line between the nose and the leading edge of the wing. The strakes were set on the same line between the trailing edge of the wing and the engine exhaust. Both supplied additional control in tight maneuvering situations. The X-31 research program produced technical data at high angles of attack. This information is giving engineers and aircraft designers a better understanding of aerodynamics, effectiveness of flight controls and thrust vectoring, and airflow phenomena at high angles of attack. This understanding is expected to lead to design methods that provide better maneuverability in future high performance aircraft and make them safer to fly. An international test organization of about 110 people, managed by the Advanced Research Projects Agency (ARPA), conducted the flight operations at NASA Dryden. The ARPA had requested flight research for the X-31 aircraft be moved there in February 1992. In addition to ARPA and NASA, the international test organization (ITO) included the U.S. Navy, the U.S. Air Force, Rockwell International, the Federal Republic of Germany, and Daimler-Benz Aerospace (formerly Messerschmitt-Bolkow-Blohm and Deutsche Aerospace). NASA was responsible for flight research operations, aircraft maintenance, and research engineering once the program moved to Dryden. The No. 1 X-31 aircraft was lost in an accident Jan. 19, 1995. The pilot, Karl Heinz-Lang, of the Federal Republic of Germany, ejected safely before the aircraft crashed in an unpopulated desert area just north of Edwards. The X-31 program logged an X-plane record of 580 flights during the program, including 555 research missions and 21 in Europe for the 1995 Paris Air Show. A total of 14 pilots representing all agencies of the ITO flew the aircraft. This 39-second clip begins as the X-31 performs a short loop at the top of a stall maneuver, then quickly reverses its course first left, then right by means of thrust vectoring -- thereby gaining a tactical advantage over a putative opponent in air-to-air combat.

X-31 in flight - Herbst Turn

NASA Technical Reports Server (NTRS)

1995-01-01

Two X-31 Enhanced Fighter Maneuverability (EFM) demonstrators were flown at the Rockwell International facility, Palmdale, California, and the NASA Dryden Flight Research Center, Edwards, California, to obtain data that may apply to the design of highly-maneuverable next-generation fighters. The program had its first flight on October 11, 1990, in Palmdale; it ended in June 1995. The X-31 program demonstrated the value of thrust vectoring (directing engine exhaust flow) coupled with advanced flight control systems, to provide controlled flight during close-in air combat at very high angles of attack. The result of this increased maneuverability is an airplane with a significant advantage over conventional fighters. 'Angle-of-attack' (alpha) is an engineering term to describe the angle of an aircraft body and wings relative to its actual flight path. During maneuvers, pilots often fly at extreme angles of attack -- with the nose pitched up while the aircraft continues in its original direction. This can lead to loss of control and result in the loss of the aircraft, pilot or both. Three thrust-vectoring paddles made of graphite epoxy mounted on the exhaust nozzle of the X-31 aircraft directed the exhaust flow to provide control in pitch (up and down) and yaw (right and left) to improve control. The paddles can sustain heat of up to 1,500 degrees centigrade for extended periods of time. In addition the X-31 aircraft were configured with movable forward canards and fixed aft strakes. The canards were small wing-like structures set on the wing line between the nose and the leading edge of the wing. The strakes were set on the same line between the trailing edge of the wing and the engine exhaust. Both supplied additional control in tight maneuvering situations. The X-31 research program produced technical data at high angles of attack. This information is giving engineers and aircraft designers a better understanding of aerodynamics, effectiveness of flight controls and thrust vectoring, and airflow phenomena at high angles of attack. This understanding is expected to lead to design methods that provide better maneuverability in future high performance aircraft and make them safer to fly. An international test organization of about 110 people, managed by the Advanced Research Projects Agency (ARPA), conducted the flight operations at NASA Dryden. The ARPA had requested flight research for the X-31 aircraft be moved there in February 1992. In addition to ARPA and NASA, the international test organization (ITO) included the U.S. Navy, the U.S. Air Force, Rockwell International, the Federal Republic of Germany, and Daimler-Benz Aerospace (formerly Messerschmitt-Bolkow-Blohm and Deutsche Aerospace). NASA was responsible for flight research operations, aircraft maintenance, and research engineering once the program moved to Dryden. The No. 1 X-31 aircraft was lost in an accident January 19, 1995. The pilot, Karl Heinz-Lang, of the Federal Republic of Germany, ejected safely before the aircraft crashed in an unpopulated desert area just north of Edwards. The X-31 program logged an X-plane record of 580 flights during the program, including 555 research missions and 21 in Europe for the 1995 Paris Air Show. A total of 14 pilots representing all agencies of the ITO flew the aircraft. This 32-second clip shows the aircraft at the top of a stall and then thrust vectoring itself around to attain a new heading, thereby allowing the aircraft to gain the advantage over a putative opponent in air-to-air combat. This maneuver is also known as a 'J turn.'

Design of a cusped field thruster for drag-free flight

NASA Astrophysics Data System (ADS)

Liu, H.; Chen, P. B.; Sun, Q. Q.; Hu, P.; Meng, Y. C.; Mao, W.; Yu, D. R.

2016-09-01

Drag-free flight has played a more and more important role in many space missions. The thrust control system is the key unit to achieve drag-free flight by providing a precise compensation for the disturbing force except gravity. The cusped field thruster has shown a significant potential to be capable of the function due to its long life, high efficiency, and simplicity. This paper demonstrates a cusped field thruster's feasibility in drag-free flight based on its instinctive characteristics and describes a detailed design of a cusped field thruster made by Harbin Institute of Technology (HIT). Furthermore, the performance test is conducted, which shows that the cusped field thruster can achieve a continuously variable thrust from 1 to 20 mN with a low noise and high resolution below 650 W, and the specific impulse can achieve 1800 s under a thrust of 18 mN and discharge voltage of 1000 V. The thruster's overall performance indicates that the cusped field thruster is quite capable of achieving drag-free flight. With the further optimization, the cusped field thruster will exhibit a more extensive application value.

Flight Test Results for the F-16XL With a Digital Flight Control System

NASA Technical Reports Server (NTRS)

Stachowiak, Susan J.; Bosworth, John T.

2004-01-01

In the early 1980s, two F-16 airplanes were modified to extend the fuselage length and incorporate a large area delta wing planform. These two airplanes, designated the F-16XL, were designed by the General Dynamics Corporation (now Lockheed Martin Tactical Aircraft Systems) (Fort Worth, Texas) and were prototypes for a derivative fighter evaluation program conducted by the United States Air Force. Although the concept was never put into production, the F-16XL prototypes provided a unique planform for testing concepts in support of future high-speed supersonic transport aircraft. To extend the capabilities of this testbed vehicle the F-16XL ship 1 aircraft was upgraded with a digital flight control system. The added flexibility of a digital flight control system increases the versatility of this airplane as a testbed for aerodynamic research and investigation of advanced technologies. This report presents the handling qualities flight test results covering the envelope expansion of the F-16XL with the digital flight control system.

Development and Flight Testing of a Neural Network Based Flight Control System on the NF-15B Aircraft

NASA Technical Reports Server (NTRS)

Bomben, Craig R.; Smolka, James W.; Bosworth, John T.; Silliams-Hayes, Peggy S.; Burken, John J.; Larson, Richard R.; Buschbacher, Mark J.; Maliska, Heather A.

2006-01-01

The Intelligent Flight Control System (IFCS) project at the NASA Dryden Flight Research Center, Edwards AFB, CA, has been investigating the use of neural network based adaptive control on a unique NF-15B test aircraft. The IFCS neural network is a software processor that stores measured aircraft response information to dynamically alter flight control gains. In 2006, the neural network was engaged and allowed to learn in real time to dynamically alter the aircraft handling qualities characteristics in the presence of actual aerodynamic failure conditions injected into the aircraft through the flight control system. The use of neural network and similar adaptive technologies in the design of highly fault and damage tolerant flight control systems shows promise in making future aircraft far more survivable than current technology allows. This paper will present the results of the IFCS flight test program conducted at the NASA Dryden Flight Research Center in 2006, with emphasis on challenges encountered and lessons learned.

Thin film strain transducer. [in-flight measurement of stress or strain in walls of high altitude balloons

NASA Technical Reports Server (NTRS)

Rand, J. L.

1981-01-01

Previous attempts to develop an appropriate sensor for measuring the stress or strain of high altitude balloons during flight are reviewed as well as the various conditions that must be met by such a device. The design, development and calibration of a transducer which promises to satisfy the necessary design constraints are described. The thin film strain transducer has a low effective modulus so as not to interfere with the strain that would naturally occur in the balloon. In addition, the transducer has a high sensitivity to longitudinal strain (7.216 mV/V/unit strain) which is constant for all temperature from room temperature to -80 C and all strains from 5 percent compression to 10 percent tensile strain. At the same time, the sensor is relatively insensitive (0.27 percent) to transverse forces. The device has a standard 350 ohm impedance which is compatible with available bridge balance, amplification and telemetry instrumentation now available for balloon flight. Recommendations are included for improved coatings to provide passive thermal control as well as model, tethered and full scale flight testing.

Thermoregulation of water foraging honeybees--balancing of endothermic activity with radiative heat gain and functional requirements.

PubMed

Kovac, Helmut; Stabentheiner, Anton; Schmaranzer, Sigurd

2010-12-01

Foraging honeybees are subjected to considerable variations of microclimatic conditions challenging their thermoregulatory ability. Solar heat is a gain in the cold but may be a burden in the heat. We investigated the balancing of endothermic activity with radiative heat gain and physiological functions of water foraging Apis mellifera carnica honeybees in the whole range of ambient temperatures (T(a)) and solar radiation they are likely to be exposed in their natural environment in Middle Europe. The mean thorax temperature (T(th)) during foraging stays was regulated at a constantly high level (37.0-38.5 °C) in a broad range of T(a) (3-30 °C). At warmer conditions (T(a)=30-39 °C) T(th) increased to a maximal level of 45.3 °C. The endothermic temperature excess (difference of T(body)-T(a) of living and dead bees) was used to assess the endogenously generated temperature elevation as a correlate of energy turnover. Up to a T(a) of ∼30 °C bees used solar heat gain for a double purpose: to reduce energetic expenditure and to increase T(th) by about 1-3 °C to improve force production of flight muscles. At higher T(a) they exhibited cooling efforts to get rid of excess heat. A high T(th) also allowed regulation of the head temperature high enough to guarantee proper function of the bees' suction pump even at low T(a). This shortened the foraging stays and this way reduced energetic costs. With decreasing T(a) bees also reduced arrival body weight and crop loading to do both minimize costs and optimize flight performance. Copyright © 2010 Elsevier Ltd. All rights reserved.

Post-Stall Aerodynamic Modeling and Gain-Scheduled Control Design

NASA Technical Reports Server (NTRS)

Wu, Fen; Gopalarathnam, Ashok; Kim, Sungwan

2005-01-01

A multidisciplinary research e.ort that combines aerodynamic modeling and gain-scheduled control design for aircraft flight at post-stall conditions is described. The aerodynamic modeling uses a decambering approach for rapid prediction of post-stall aerodynamic characteristics of multiple-wing con.gurations using known section data. The approach is successful in bringing to light multiple solutions at post-stall angles of attack right during the iteration process. The predictions agree fairly well with experimental results from wind tunnel tests. The control research was focused on actuator saturation and .ight transition between low and high angles of attack regions for near- and post-stall aircraft using advanced LPV control techniques. The new control approaches maintain adequate control capability to handle high angle of attack aircraft control with stability and performance guarantee.

An Overview of Active Flow Control Enhanced Vertical Tail Technology Development

NASA Technical Reports Server (NTRS)

Lin, John C.; Andino, Marlyn Y.; Alexander, Michael G.; Whalen, Edward A.; Spoor, Marc A.; Tran, John T.; Wygnanski, Israel J.

2016-01-01

This paper summarizes a joint NASA/Boeing research effort to advance Active Flow Control (AFC) technology to enhance aerodynamic efficiency of a vertical tail. Sweeping jet AFC technology was successfully tested on subscale and full-scale models as well as in flight. The subscale test was performed at Caltech on a 14% scale model. More than 50% side force enhancement was achieved by the sweeping jet actuation when the momentum coefficient was 1.7%. AFC caused significant increases in suction pressure on the actuator side and associated side force enhancement. Subsequently, a full-scale Boeing 757 vertical tail model equipped with sweeping jets was tested at the National Full-Scale Aerodynamics Complex 40- by 80-Foot Wind Tunnel at NASA Ames Research Center. There, flow separation control optimization was performed at near flight conditions. Greater than 20% increase in side force were achieved for the maximum rudder deflection of 30deg at the key sideslip angles (0deg and -7.5deg) with a 31-actuator AFC configuration. Based on these tests, the momentum coefficient is shown to be a necessary, but not sufficient parameter to use for design and scaling of sweeping jet AFC from subscale tests to full-scale applications. Leveraging the knowledge gained from the wind tunnel tests, the AFC-enhanced vertical tail technology was successfully flown on the Boeing 757 ecoDemonstrator in the spring of 2015.

Design and Flight Tests of an Adaptive Control System Employing Normal-Acceleration Command

NASA Technical Reports Server (NTRS)

McNeill, Water E.; McLean, John D.; Hegarty, Daniel M.; Heinle, Donovan R.

1961-01-01

An adaptive control system employing normal-acceleration command has been designed with the aid of an analog computer and has been flight tested. The design of the system was based on the concept of using a mathematical model in combination with a high gain and a limiter. The study was undertaken to investigate the application of a system of this type to the task of maintaining nearly constant dynamic longitudinal response of a piloted airplane over the flight envelope without relying on air data measurements for gain adjustment. The range of flight conditions investigated was between Mach numbers of 0.36 and 1.15 and altitudes of 10,000 and 40,000 feet. The final adaptive system configuration was derived from analog computer tests, in which the physical airplane control system and much of the control circuitry were included in the loop. The method employed to generate the feedback signals resulted in a model whose characteristics varied somewhat with changes in flight condition. Flight results showed that the system limited the variation in longitudinal natural frequency of the adaptive airplane to about half that of the basic airplane and that, for the subsonic cases, the damping ratio was maintained between 0.56 and 0.69. The system also automatically compensated for the transonic trim change. Objectionable features of the system were an exaggerated sensitivity of pitch attitude to gust disturbances, abnormally large pitch attitude response for a given pilot input at low speeds, and an initial delay in normal-acceleration response to pilot control at all flight conditions. The adaptive system chatter of +/-0.05 to +/-0.10 of elevon at about 9 cycles per second (resulting in a maximum airplane normal-acceleration response of from +/-0.025 g to +/- 0.035 g) was considered by the pilots to be mildly objectionable but tolerable.

Analysis of Handling Qualities Design Criteria for Active Inceptor Force-Feel Characteristics

NASA Technical Reports Server (NTRS)

Malpica, Carlos A.; Lusardi, Jeff A.

2013-01-01

The force-feel system characteristics of the cyclic inceptors of most helicopters are set based on the characteristics of the mechanical components in the control system (mass, springs, friction dampers, etc.). For these helicopters, the force-feel characteristics typically remain constant over the entire flight envelope, with perhaps a trim release to minimize control forces while maneuvering. With the advent of fly-by-wire control systems and active inceptors in helicopters, the force-feel characteristics are now determined by the closed-loop response of the active inceptor itself as defined by the inertia, force/displacement gradient, damping, breakout force and detent shape configuration parameters in the inceptor control laws. These systems give the flexibility to dynamically prescribe different feel characteristics for different control modes or flight conditions, and the ability to provide tactile cueing to the pilot through the actively controlled side-stick or center-stick cyclic inceptor. For rotorcraft, a few studies have been conducted to assess the effects of cyclic force-feel characteristics on handling qualities in flight. An early study provided valuable insight into the static force-deflection characteristics (force gradient) and the number of axes controlled by the side-stick controller for the U.S. Army's Advanced Digital/Optical Control System (ADOCS) demonstrator aircraft [1]. The first of a series of studies providing insight on the inceptor dynamic force-feel characteristics was conducted on the NASA/Army CH-47B variable-stability helicopter [2]. This work led to a proposed requirement that set boundaries based on the cyclic natural frequency and inertia, with the stipulation of a lower damping ratio limit of 0.3 [3]. A second study was conducted by the Canadian Institute for Aerospace Research using their variable-stability Bell 205A helicopter [4]. This research suggested boundaries for stick dynamics based on natural frequency and damping ratio. While these two studies produced boundaries for acceptable/unacceptable stick dynamics for rotorcraft, they were not able to provide guidance on how variations of the stick dynamics in the acceptable region impact handling qualities. More recently, a ground based simulation study [5] suggested little benefit was to be obtained from variations of the damping ratio for a side-stick controller exhibiting high natural frequencies (greater than 17 rad/s) and damping ratios (greater than 2.0). A flight test campaign was conducted concurrently on the RASCAL JUH-60A in-flight simulator and the ACT/FHS EC-135 in flight simulator [6]. Upon detailed analysis of the pilot evaluations the study identified a clear preference for a high damping ratio and natural frequency of the center stick inceptors. Side stick controllers were found to be less sensitive to the damping. While these studies have compiled a substantial amount of data, in the form of qualitative and quantitative pilot opinion, a fundamental analysis of the effect of the inceptor force-feel system on flight control is found to be lacking. The study of Ref. [6] specifically concluded that a systematic analysis was necessary, since discrepancies with the assigned handling qualities showed that proposed analytical design metrics, or criteria, were not suitable. The overall goal of the present study is to develop a clearer fundamental understanding of the underlying mechanisms associated with the inceptor dynamics that govern the handling qualities using a manageable analytical methodology.

Prediction of forces and moments for flight vehicle control effectors: Workplan

NASA Technical Reports Server (NTRS)

Maughmer, Mark D.

1989-01-01

Two research activities directed at hypersonic vehicle configurations are currently underway. The first involves the validation of a number of classical local surface inclination methods commonly employed in preliminary design studies of hypersonic flight vehicles. Unlike studies aimed at validating such methods for predicting overall vehicle aerodynamics, this effort emphasizes validating the prediction of forces and moments for flight control studies. Specifically, several vehicle configurations for which experimental or flight-test data are available are being examined. By comparing the theoretical predictions with these data, the strengths and weaknesses of the local surface inclination methods can be ascertained and possible improvements suggested. The second research effort, of significance to control during take-off and landing of most proposed hypersonic vehicle configurations, is aimed at determining the change due to ground effect in control effectiveness of highly swept delta planforms. Central to this research is the development of a vortex-lattice computer program which incorporates an unforced trailing vortex sheet and an image ground plane. With this program, the change in pitching moment of the basic vehicle due to ground proximity, and whether or not there is sufficient control power available to trim, can be determined. In addition to the current work, two different research directions are suggested for future study. The first is aimed at developing an interactive computer program to assist the flight controls engineer in determining the forces and moments generated by different types of control effectors that might be used on hypersonic vehicles. The first phase of this work would deal in the subsonic portion of the flight envelope, while later efforts would explore the supersonic/hypersonic flight regimes. The second proposed research direction would explore methods for determining the aerodynamic trim drag of a generic hypersonic flight vehicle and ways in which it can be minimized through vehicle design and trajectory optimization.

Preliminary assessment of the robustness of dynamic inversion based flight control laws

NASA Technical Reports Server (NTRS)

Snell, S. A.

1992-01-01

Dynamic-inversion-based flight control laws present an attractive alternative to conventional gain-scheduled designs for high angle-of-attack maneuvering, where nonlinearities dominate the dynamics. Dynamic inversion is easily applied to the aircraft dynamics requiring a knowledge of the nonlinear equations of motion alone, rather than an extensive set of linearizations. However, the robustness properties of the dynamic inversion are questionable especially when considering the uncertainties involved with the aerodynamic database during post-stall flight. This paper presents a simple analysis and some preliminary results of simulations with a perturbed database. It is shown that incorporating integrators into the control loops helps to improve the performance in the presence of these perturbations.

Active Aerodynamic Load Reduction on a Rotorcraft Fuselage With Rotor Effects: A CFD Validation Effort

NASA Technical Reports Server (NTRS)

Allan, Brian G.; Schaeffler, Norman W.; Jenkins, Luther N.; Yao, Chung-Sheng; Wong, Oliver D.; Tanner, Philip E.

2015-01-01

A rotorcraft fuselage is typically designed with an emphasis on operational functionality with aerodynamic efficiency being of secondary importance. This results in a significant amount of drag during high-speed forward flight that can be a limiting factor for future high-speed rotorcraft designs. To enable higher speed flight, while maintaining a functional fuselage design (i.e., a large rear cargo ramp door), the NASA Rotary Wing Project has conducted both experimental and computational investigations to assess active flow control as an enabling technology for fuselage drag reduction. This paper will evaluate numerical simulations of a flow control system on a generic rotorcraft fuselage with a rotor in forward flight using OVERFLOW, a structured mesh Reynolds-averaged Navier-Stokes flow solver developed at NASA. The results are compared to fuselage forces, surface pressures, and PN flow field data obtained in a wind tunnel experiment conducted at the NASA Langley 14-by 22-Foot Subsonic Tunnel where significant drag and download reductions were demonstrated using flow control. This comparison showed that the Reynolds-averaged Navier-Stokes flow solver was unable to predict the fuselage forces and pressure measurements on the ramp for the baseline and flow control cases. While the CFD was able to capture the flow features, it was unable to accurately predict the performance of the flow control.

Modified Dynamic Inversion to Control Large Flexible Aircraft: What's Going On?

NASA Technical Reports Server (NTRS)

Gregory, Irene M.

1999-01-01

High performance aircraft of the future will be designed lighter, more maneuverable, and operate over an ever expanding flight envelope. One of the largest differences from the flight control perspective between current and future advanced aircraft is elasticity. Over the last decade, dynamic inversion methodology has gained considerable popularity in application to highly maneuverable fighter aircraft, which were treated as rigid vehicles. This paper explores dynamic inversion application to an advanced highly flexible aircraft. An initial application has been made to a large flexible supersonic aircraft. In the course of controller design for this advanced vehicle, modifications were made to the standard dynamic inversion methodology. The results of this application were deemed rather promising. An analytical study has been undertaken to better understand the nature of the made modifications and to determine its general applicability. This paper presents the results of this initial analytical look at the modifications to dynamic inversion to control large flexible aircraft.

SHARP-B2: Flight Test Objectives, Project Implementation and Initial Results

NASA Technical Reports Server (NTRS)

Salute, Joan; Bull, Jeff; Rasky, Dan; Keese, David; Arnold, Jim (Technical Monitor)

2001-01-01

On September 28, 2000 the SHARP-B2 flight experiment was launched from Vandenberg Air Force Base, California. SHARP-B2 is the 2nd Ballistic flight test in the SHARP (Slender Hypervelocity Aerothermodynamic Research Probes) program which develops and tests new thermal protection materials and sharp body concepts. The flight tested Ultra-High Temperature Ceramics (UHTCs), which may radically change the design and performance of future aerospace vehicles. The new designs may overturn an age-old tenet of aerodynamics: that blunt-body aerospace vehicles, but not those with sharp leading edges, can survive the searing temperatures created as the vehicles tear through the atmosphere. Sharp leading edges offer numerous advantages over the blunt-body design currently in use. They could allow a space shuttle or crew return vehicle to maneuver in space more like an airplane and potentially allow astronauts to return to Earth from anywhere on orbit. They may allow improved astronaut safety by decreasing the risk of aborting into the ocean. They may reduce the electromagnetic interference that causes the communications blackouts that plague reentering blunt-body space vehicles. Reducing the amount of drag could lead to a reduction in propulsion requirements. Planetary probes could make use of sharp-body technology for aerobraking and to maximize their maneuvering capability. SHARP-B2 was a joint effort among NASA Ames, Sandia National Laboratories, the U.S. Air Force and the U.S. Army. It was funded by the Pathfinder Program at NASA's Marshall Space Flight Center. The SHARP-B2 payload was carried aboard a U.S. Air Force Minuteman III missile carrying a modified Mk 12A reentry vehicle (RV), which blasted off from Vandenberg Air Force Base near Lompoc, CA, at 3:01 a.m. PDT on Sept. 28. The RV was equipped with four 5. 1 inch-long strakes, or sharp leading edges. Each strake contained three UHTCs: ZrB2/SiC/C; ZrB2/SiC; and HfB2/SiC. Once it reached an altitude of about 400 nautical miles, the RV was released, returning through Earth's atmosphere at speeds exceeding Mach 22. One pair of strakes was designed to retract just before reaching temperatures high enough to cause the material to begin ablating. The other pair was designed to retract shortly after ablation began, at an expected temperature of nearly 5,100 degrees Fahrenheit. Sensors in the strakes measured how closely performance matched pre-flight calculations, and data was successfully collected throughout the 23-minute flight. A parachute was deployed (but not fully inflated) and the RV splashed down in a lagoon at the Kwajalein missile range in the Pacific Ocean. Within 3 hours radar track analysis showed ocean entry to be precisely at the latitude/longitude coordinates estimated during pre-flight simulation. An hour later a ship was deployed by the Army to recover the reentry vehicle, which was recovered in 165 feet of water, just 500 feet from its planned splash-down point. This is the first RV recovery in over a decade.

Gimbals Drive and Control Electronics Design, Development and Testing of the LRO High Gain Antenna and Solar Array Systems

NASA Technical Reports Server (NTRS)

Chernyakov, Boris; Thakore, Kamal

2010-01-01

Launched June 18, 2009 on an Atlas V rocket, NASA's Lunar Reconnaissance Orbiter (LRO) is the first step in NASA's Vision for Space Exploration program and for a human return to the Moon. The spacecraft (SC) carries a wide variety of scientific instruments and provides an extraordinary opportunity to study the lunar landscape at resolutions and over time scales never achieved before. The spacecraft systems are designed to enable achievement of LRO's mission requirements. To that end, LRO's mechanical system employed two two-axis gimbal assemblies used to drive the deployment and articulation of the Solar Array System (SAS) and the High Gain Antenna System (HGAS). This paper describes the design, development, integration, and testing of Gimbal Control Electronics (GCE) and Actuators for both the HGAS and SAS systems, as well as flight testing during the on-orbit commissioning phase and lessons learned.

Aircraft control forces and EMG activity in a C-130 Hercules during strength-critical maneuvers.

PubMed

Hewson, D J; McNair, P J; Marshall, R N

2001-03-01

The force levels required to operate aircraft controls should be readily generated by pilots, without undue fatigue or exertion. However, maximum pilot applied forces, as specified in aircraft design standards, were empirically derived from the subjective comments of test pilots, and may not be applicable for the majority of pilots. Further, experienced RNZAF Hercules flying instructors have indicated that endurance and fatigue are problems for Hercules pilots. The aim of this study was to quantify aircraft control forces during emergency maneuvers in a Hercules aircraft and compare these forces with design standards. In addition, EMG data were recorded as an indicator of muscle fatigue during flight. Six subjects were tested in a C-130 Hercules aircraft. The maneuvers performed were low-level dynamic flight, one engine-off straight-and-level flight, and a two-engines-off simulated approach. The variables recorded were pilot-applied forces and EMG activity. Left rudder pedal force and vastus lateralis activity were both significantly greater during engine-off maneuvers than during low-level dynamic flight (p < 0.05). Maximum aircraft control forces for all controls were within 10% of the design standards. The mean EMG activity across all muscles and maneuvers was 26% MVC, with a peak of 61% MVC in vastus lateralis during the two-engine-off approach. The median frequency of the vastus lateralis EMG signal decreased 13.0% and 16.0% for the one engine-off and two-engine-off maneuvers, respectively. The forces required to fly a Hercules aircraft during emergency maneuvers are similar to the aircraft design standards. However, the levels of vastus lateralis muscle activation observed during the engine-off maneuvers can be sustained for approximately 1 min only. Thus, if two engines fail more than 1 min before landing, pilots may have to alternate control of the aircraft to share the workload and enable the aircraft to land safely.

Free-Flight Experiments in LISA Pathfinder

NASA Technical Reports Server (NTRS)

Thorpe, J. I.; Cutler, C. J.; Hewitson, M.; Jennrich, O.; Maghami, P.; Paczkowski, S.; Russano, G.; Vitale, S.; Weber, W. J.

2014-01-01

The LISA Pathfinder mission will demonstrate the technology of drag-free test masses for use as inertial references in future space-based gravitational wave detectors. To accomplish this, the Pathfinder spacecraft will perform drag-free flight about a test mass while measuring the acceleration of this primary test mass relative to a second reference test mass. Because the reference test mass is contained within the same spacecraft, it is necessary to apply forces on it to maintain its position and attitude relative to the spacecraft. These forces are a potential source of acceleration noise in the LISA Pathfinder system that are not present in the full LISA configuration. While LISA Pathfinder has been designed to meet it's primary mission requirements in the presence of this noise, recent estimates suggest that the on-orbit performance may be limited by this 'suspension noise'. The drift-mode or free-flight experiments provide an opportunity to mitigate this noise source and further characterize the underlying disturbances that are of interest to the designers of LISA-like instruments. This article provides a high-level overview of these experiments and the methods under development to analyze the resulting data.

Design Genetic Algorithm Optimization Education Software Based Fuzzy Controller for a Tricopter Fly Path Planning

ERIC Educational Resources Information Center

Tran, Huu-Khoa; Chiou, Juing -Shian; Peng, Shou-Tao

2016-01-01

In this paper, the feasibility of a Genetic Algorithm Optimization (GAO) education software based Fuzzy Logic Controller (GAO-FLC) for simulating the flight motion control of Unmanned Aerial Vehicles (UAVs) is designed. The generated flight trajectories integrate the optimized Scaling Factors (SF) fuzzy controller gains by using GAO algorithm. The…

CATE: A Case Study of an Interdisciplinary Student-Led Microgravity Experiment

NASA Astrophysics Data System (ADS)

Colwell, J. E.; Dove, A.; Lane, S. S.; Tiller, C.; Whitaker, A.; Lai, K.; Hoover, B.; Benjamin, S.

2015-12-01

The Collisional Accretion Experiment (CATE) was designed, built, and flown on NASA's C-9 parabolic flight airplane in less than a year by an interdisciplinary team of 6 undergraduate students under the supervision of two faculty. CATE was selected in the initial NASA Undergraduate Student Instrument Project (USIP) solicitation in the Fall of 2013, and the experiment flight campaign was in July 2014. The experiment studied collisions between different particle populations at low velocities (sub-m/s) in a vacuum and microgravity to gain insight into processes in the protoplanetary disk and planetary ring systems. Faculty provided the experiment concept and key experiment design parameters, and the student team developed the detailed hardware design for all components, manufactured and tested hardware, operated the experiment in flight, and analyzed data post-flight. Students also developed and led an active social media campaign and education and public outreach campaign to engage local high school students in the project. The ability to follow an experiment through from conception to flight was a key benefit for undergraduate students whose available time for projects such as this is frequently limited to their junior and senior years. Key factors for success of the program included having an existing laboratory infrastructure and experience in developing flight payloads and an intrinsically simple experiment concept. Students were highly motivated, in part, by their sense of technical and scientific ownership of the project, and this engagement was key to the project's success.

Physics-based simulations of aerial attacks by peregrine falcons reveal that stooping at high speed maximizes catch success against agile prey.

PubMed

Mills, Robin; Hildenbrandt, Hanno; Taylor, Graham K; Hemelrijk, Charlotte K

2018-04-01

The peregrine falcon Falco peregrinus is renowned for attacking its prey from high altitude in a fast controlled dive called a stoop. Many other raptors employ a similar mode of attack, but the functional benefits of stooping remain obscure. Here we investigate whether, when, and why stooping promotes catch success, using a three-dimensional, agent-based modeling approach to simulate attacks of falcons on aerial prey. We simulate avian flapping and gliding flight using an analytical quasi-steady model of the aerodynamic forces and moments, parametrized by empirical measurements of flight morphology. The model-birds' flight control inputs are commanded by their guidance system, comprising a phenomenological model of its vision, guidance, and control. To intercept its prey, model-falcons use the same guidance law as missiles (pure proportional navigation); this assumption is corroborated by empirical data on peregrine falcons hunting lures. We parametrically vary the falcon's starting position relative to its prey, together with the feedback gain of its guidance loop, under differing assumptions regarding its errors and delay in vision and control, and for three different patterns of prey motion. We find that, when the prey maneuvers erratically, high-altitude stoops increase catch success compared to low-altitude attacks, but only if the falcon's guidance law is appropriately tuned, and only given a high degree of precision in vision and control. Remarkably, the optimal tuning of the guidance law in our simulations coincides closely with what has been observed empirically in peregrines. High-altitude stoops are shown to be beneficial because their high airspeed enables production of higher aerodynamic forces for maneuvering, and facilitates higher roll agility as the wings are tucked, each of which is essential to catching maneuvering prey at realistic response delays.

Feedback control laws for highly maneuverable aircraft

NASA Technical Reports Server (NTRS)

Garrard, William L.; Balas, Gary J.

1992-01-01

The results of a study of the application of H infinity and mu synthesis techniques to the design of feedback control laws for the longitudinal dynamics of the High Angle of Attack Research Vehicle (HARV) are presented. The objective of this study is to develop methods for the design of feedback control laws which cause the closed loop longitudinal dynamics of the HARV to meet handling quality specifications over the entire flight envelope. Control law designs are based on models of the HARV linearized at various flight conditions. The control laws are evaluated by both linear and nonlinear simulations of typical maneuvers. The fixed gain control laws resulting from both the H infinity and mu synthesis techniques result in excellent performance even when the aircraft performs maneuvers in which the system states vary significantly from their equilibrium design values. Both the H infinity and mu synthesis control laws result in performance which compares favorably with an existing baseline longitudinal control law.

Jumping in simulated and true microgravity: response to maximal efforts with three landing types

NASA Technical Reports Server (NTRS)

D'Andrea, Susan E.; Perusek, Gail P.; Rajulu, Sudhakar; Perry, Julie; Davis, Brian L.

2005-01-01

BACKGROUND: Exercise is a promising countermeasure to the physiological deconditioning experienced in microgravity, but has not proven effective in eliminating the ongoing loss of bone mineral, most likely due to the lack of high-impact forces and loading rates during in-flight activity. We wanted to determine lower-extremity response to high-impact jumping exercises in true and simulated microgravity and establish if 1-G force magnitudes can be achieved in a weightless environment. METHODS: Jumping experiments were performed in a ground-based zero-gravity simulator (ZGS) in 1 G, and during parabolic flight with a gravity-replacement system. There were 12 subjects who participated in the study, with 4 subjects common to both conditions. Force, loading rates, jump height, and kinematics were analyzed during jumps with three distinct landings: two-footed toe-heel, one-footed toe-heel, and flat-footed. Gravity replacement loads of 45%, 60%, 75%, and 100% bodyweight were used in the ZGS; because of time constraints, these loads were limited to 60% and 75% bodyweight in parabolic flight. RESULTS: Average peak ground-reaction forces during landing ranged between 1902+/-607 and 2631+/-663 N in the ZGS and between 1683+/-807 and 2683+/-1174 N in the KC-135. No significant differences were found between the simulated and true microgravity conditions, but neither condition achieved the magnitudes found in 1 G. CONCLUSION: Data support the hypothesis that jumping exercises can impart high-impact forces during weightlessness and that the custom-designed ZGS will replicate what is experienced in true microgravity.

X-15 test pilots - Engle, Rushworth, McKay, Knight, Thompson, and Dana

NASA Technical Reports Server (NTRS)

1966-01-01

The X-15 flight crew, left to right; Air Force Captain Joseph H. Engle, Air Force Major Robert A. Rushworth, NASA pilot John B. 'Jack' McKay, Air Force pilot William J. 'Pete' Knight, NASA pilot Milton O. Thompson, and NASA pilot Bill Dana. of their 125 X-15 flights, 8 were above the 50 miles that constituted the Air Force's definition of the beginning of space (Engle 3, Dana 2, Rushworth, Knight, and McKay one each). NASA used the international definition of space as beginning at 62 miles above the earth. The X-15 was a rocket-powered aircraft 50 ft long with a wingspan of 22 ft. It was a missile-shaped vehicle with an unusual wedge-shaped vertical tail, thin stubby wings, and unique fairings that extended along the side of the fuselage. The X-15 weighed about 14,000 lb empty and approximately 34,000 lb at launch. The XLR-99 rocket engine, manufactured by Thiokol Chemical Corp., was pilot controlled and was capable of developing 57,000 lb of rated thrust (actual thrust reportedly climbed to 60,000 lb). North American Aviation built three X-15 aircraft for the program. The X-15 research aircraft was developed to provide in-flight information and data on aerodynamics, structures, flight controls, and the physiological aspects of high-speed, high-altitude flight. A follow-on program used the aircraft as a testbed to carry various scientific experiments beyond the Earth's atmosphere on a repeated basis. For flight in the dense air of the usable atmosphere, the X-15 used conventional aerodynamic controls such as rudder surfaces on the vertical stabilizers to control yaw and canted horizontal surfaces on the tail to control pitch when moving in synchronization or roll when moved differentially. For flight in the thin air outside of the appreciable Earth's atmosphere, the X-15 used a reaction control system. Hydrogen peroxide thrust rockets located on the nose of the aircraft provided pitch and yaw control. Those on the wings provided roll control. Because of the large fuel consumption, the X-15 was air launched from a B-52 aircraft at 45,000 ft and a speed of about 500 mph. Depending on the mission, the rocket engine provided thrust for the first 80 to 120 sec of flight. The remainder of the normal 10 to 11 min. flight was powerless and ended with a 200-mph glide landing. Generally, one of two types of X-15 flight profiles was used: a high-altitude flight plan that called for the pilot to maintain a steep rate of climb, or a speed profile that called for the pilot to push over and maintain a level altitude. The X-15 was flown over a period of nearly 10 years--June 1959 to Oct. 1968--and set the world's unofficial speed and altitude records of 4,520 mph (Mach 6.7) and 354,200 ft (over 67 mi) in a program to investigate all aspects of piloted hypersonic flight. Information gained from the highly successful X-15 program contributed to the development of the Mercury, Gemini, and Apollo manned spaceflight programs, and also the Space Shuttle program. The X-15s made a total of 199 flights and were manufactured by North American Aviation. X-15-1, serial number 56-6670, is now located at the National Air and Space Museum, Washington DC. North American X-15A-2, serial number 56-6671, is at the United States Air Force Museum, Wright-Patterson AFB, Ohio. The X-15-3, serial number 56-6672, crashed on 15 November 1967, resulting in the death of Maj. Michael J. Adams.

An Indispensable Ingredient: Flight Research and Aircraft Design

NASA Technical Reports Server (NTRS)

Gorn, Michael H.

2003-01-01

Flight research-the art of flying actual vehicles in the atmosphere in order to collect data about their behavior-has played a historic and decisive role in the design of aircraft. Naturally, wind tunnel experiments, computational fluid dynamics, and mathematical analyses all informed the judgments of the individuals who conceived of new aircraft. But flight research has offered moments of realization found in no other method. Engineer Dale Reed and research pilot Milt Thompson experienced one such epiphany on March 1, 1963, at the National Aeronautics and Space Administration s Dryden Flight Research Center in Edwards, California. On that date, Thompson sat in the cockpit of a small, simple, gumdrop-shaped aircraft known as the M2-F1, lashed by a long towline to a late-model Pontiac Catalina. As the Pontiac raced across Rogers Dry Lake, it eventually gained enough speed to make the M2-F1 airborne. Thompson braced himself for the world s first flight in a vehicle of its kind, called a lifting body because of its high lift-to-drag ratio. Reed later recounted what he saw:

Multi-Flight-Phase GPS Navigation Filter Applications to Terrestrial Vehicle Navigation and Positioning

NASA Technical Reports Server (NTRS)

Park, Young W.; Montez, Moises N.

1994-01-01

A candidate onboard space navigation filter demonstrated excellent performance (less than 8 meter level RMS semi-major axis accuracy) in performing orbit determination of a low-Earth orbit Explorer satellite using single-frequency real GPS data. This performance is significantly better than predicted by other simulation studies using dual-frequency GPS data. The study results revealed the significance of two new modeling approaches evaluated in the work. One approach introduces a single-frequency ionospheric correction through pseudo-range and phase range averaging implementation. The other approach demonstrates a precise axis-dependent characterization of dynamic sample space uncertainty to compute a more accurate Kalman filter gain. Additionally, this navigation filter demonstrates a flexibility to accommodate both perturbational dynamic and observational biases required for multi-flight phase and inhomogeneous application environments. This paper reviews the potential application of these methods and the filter structure to terrestrial vehicle and positioning applications. Both the single-frequency ionospheric correction method and the axis-dependent state noise modeling approach offer valuable contributions in cost and accuracy improvements for terrestrial GPS receivers. With a modular design approach to either 'plug-in' or 'unplug' various force models, this multi-flight phase navigation filter design structure also provides a versatile GPS navigation software engine for both atmospheric and exo-atmospheric navigation or positioning use, thereby streamlining the flight phase or application-dependent software requirements. Thus, a standardized GPS navigation software engine that can reduce the development and maintenance cost of commercial GPS receivers is now possible.

Simulated flight path control of fighter pilots and novice subjects at +3 Gz in a human centrifuge.

PubMed

Dalecki, Marc; Bock, Otmar; Guardiera, Simon

2010-05-01

We have previously shown that subjects produce exaggerated manual forces in +3 Gz. When subjects execute discrete flight path changes in a flight simulator, their performance is less stable in +3 Gz than in +1 Gz. Here we explore whether Gz-related deficits are found with continuous flight path changes. Novice subjects and fighter pilots sat in a high-fidelity flight simulator equipped with the reproduction of the Eurofighter 2000 cockpit, including the realistic flight stick, and pursued continuous altitude changes of a target airplane in +1 Gz and +3 Gz. Subjects also produced verbal responses in a Stroop task. Pursuit and Stroop tasks were administered alone and concurrently. Flight instability increased in +3 Gz compared to +1 Gz in novices (+46%), but not in pilots (+3%), and even there only during the first minute. Flight performance improved after the first minute in both subject groups. Stroop reaction time was higher in novices (+5.27%) than in pilots (+3.77%) at +3 Gz. Dual-task costs did not differ between groups or Gz levels. Deficits of force production in high Gz are largely compensated for when subjects apply forces to produce a continuously changing flight path. This compensation seems not to require additional cognitive resources and may be achieved by using visual feedback. Force production deficits in high Gz seem to have no appreciable effects on flight performance and cognitive load of experienced pilots using a force-plus-displacement stick in +3 Gz. It remains to be shown whether this conclusion extends to purely isometric sticks and to higher Gz levels.

Overview of Orion Crew Module and Launch Abort Vehicle Dynamic Stability

NASA Technical Reports Server (NTRS)

Owens, Donald B.; Aibicjpm. Vamessa V.

2011-01-01

With the retirement of the Space Shuttle, NASA is designing a new spacecraft, called Orion, to fly astronauts to low earth orbit and beyond. Characterization of the dynamic stability of the Orion spacecraft is important for the design of the spacecraft and trajectory construction. Dynamic stability affects the stability and control of the Orion Crew Module during re-entry, especially below Mach = 2.0 and including flight under the drogues. The Launch Abort Vehicle is affected by dynamic stability as well, especially during the re-orientation and heatshield forward segments of the flight. The dynamic stability was assessed using the forced oscillation technique, free-to-oscillate, ballistic range, and sub-scale free-flight tests. All of the test techniques demonstrated that in heatshield-forward flight the Crew Module and Launch Abort Vehicle are dynamically unstable in a significant portion of their flight trajectory. This paper will provide a brief overview of the Orion dynamic aero program and a high-level summary of the dynamic stability characteristics of the Orion spacecraft.

Validation of engineering methods for predicting hypersonic vehicle controls forces and moments

NASA Technical Reports Server (NTRS)

Maughmer, M.; Straussfogel, D.; Long, L.; Ozoroski, L.

1991-01-01

This work examines the ability of the aerodynamic analysis methods contained in an industry standard conceptual design code, the Aerodynamic Preliminary Analysis System (APAS II), to estimate the forces and moments generated through control surface deflections from low subsonic to high hypersonic speeds. Predicted control forces and moments generated by various control effectors are compared with previously published wind-tunnel and flight-test data for three vehicles: the North American X-15, a hypersonic research airplane concept, and the Space Shuttle Orbiter. Qualitative summaries of the results are given for each force and moment coefficient and each control derivative in the various speed ranges. Results show that all predictions of longitudinal stability and control derivatives are acceptable for use at the conceptual design stage.

A direct application of the non-linear inverse transformation flight control system design on a STOVL aircraft

NASA Technical Reports Server (NTRS)

Chung, W. W.; Mcneill, W. E.; Stortz, M. W.

1993-01-01

The nonlinear inverse transformation flight control system design method is applied to the Lockheed Ft. Worth Company's E-7D short takeoff and vertical land (STOVL) supersonic fighter/attack aircraft design with a modified General Electric F110 engine which has augmented propulsive lift capability. The system is fully augmented to provide flight path control and velocity control, and rate command attitude hold for angular axes during the transition and hover operations. In cruise mode, the flight control system is configured to provide direct thrust command, rate command attitude hold for pitch and roll axes, and sideslip command with turn coordination. A control selector based on the nonlinear inverse transformation method is designed specifically to be compatible with the propulsion system's physical configuration which has a two dimensional convergent-divergent aft nozzle, a vectorable ventral nozzle, and a thrust augmented ejector. The nonlinear inverse transformation is used to determine the propulsive forces and nozzle deflections, which in combination with the aerodynamic forces and moments (including propulsive induced contributions), and gravitational force, are required to achieve the longitudinal and vertical acceleration commands. The longitudinal control axes are fully decoupled within the propulsion system's performance envelope. A piloted motion-base flight simulation was conducted on the Vertical Motion Simulator (VMS) at NASA Ames Research Center to examine the handling qualities of this design. Based on results of the simulation, refinements to the control system have been made and will also be covered in the report.

In-flight calibration of the high-gain antenna pointing for the Mariner Venus-Mercury 1973 spacecraft

NASA Technical Reports Server (NTRS)

Hardman, J. M.; Havens, W. F.; Ohtakay, H.

1975-01-01

The methods used to in-flight calibrate the pointing direction of the Mariner Venus-Mercury 1973 spacecraft high gain antenna and the achieved antenna pointing accuracy are described. The overall pointing calibration was accomplished by performing calibration sequences at a number of points along the spacecraft trajectory. Each of these consisted of articulating the antenna about the expected spacecraft-earth vector to determine systematic pointing errors. The high gain antenna pointing system, the error model used in the calibration, and the calibration and pointing strategy and results are discussed.

Surface tension dominates insect flight on fluid interfaces.

PubMed

Mukundarajan, Haripriya; Bardon, Thibaut C; Kim, Dong Hyun; Prakash, Manu

2016-03-01

Flight on the 2D air-water interface, with body weight supported by surface tension, is a unique locomotion strategy well adapted for the environmental niche on the surface of water. Although previously described in aquatic insects like stoneflies, the biomechanics of interfacial flight has never been analysed. Here, we report interfacial flight as an adapted behaviour in waterlily beetles (Galerucella nymphaeae) which are also dexterous airborne fliers. We present the first quantitative biomechanical model of interfacial flight in insects, uncovering an intricate interplay of capillary, aerodynamic and neuromuscular forces. We show that waterlily beetles use their tarsal claws to attach themselves to the interface, via a fluid contact line pinned at the claw. We investigate the kinematics of interfacial flight trajectories using high-speed imaging and construct a mathematical model describing the flight dynamics. Our results show that non-linear surface tension forces make interfacial flight energetically expensive compared with airborne flight at the relatively high speeds characteristic of waterlily beetles, and cause chaotic dynamics to arise naturally in these regimes. We identify the crucial roles of capillary-gravity wave drag and oscillatory surface tension forces which dominate interfacial flight, showing that the air-water interface presents a radically modified force landscape for flapping wing flight compared with air. © 2016. Published by The Company of Biologists Ltd.

Surface tension dominates insect flight on fluid interfaces

PubMed Central

Mukundarajan, Haripriya; Bardon, Thibaut C.; Kim, Dong Hyun; Prakash, Manu

2016-01-01

ABSTRACT Flight on the 2D air–water interface, with body weight supported by surface tension, is a unique locomotion strategy well adapted for the environmental niche on the surface of water. Although previously described in aquatic insects like stoneflies, the biomechanics of interfacial flight has never been analysed. Here, we report interfacial flight as an adapted behaviour in waterlily beetles (Galerucella nymphaeae) which are also dexterous airborne fliers. We present the first quantitative biomechanical model of interfacial flight in insects, uncovering an intricate interplay of capillary, aerodynamic and neuromuscular forces. We show that waterlily beetles use their tarsal claws to attach themselves to the interface, via a fluid contact line pinned at the claw. We investigate the kinematics of interfacial flight trajectories using high-speed imaging and construct a mathematical model describing the flight dynamics. Our results show that non-linear surface tension forces make interfacial flight energetically expensive compared with airborne flight at the relatively high speeds characteristic of waterlily beetles, and cause chaotic dynamics to arise naturally in these regimes. We identify the crucial roles of capillary–gravity wave drag and oscillatory surface tension forces which dominate interfacial flight, showing that the air–water interface presents a radically modified force landscape for flapping wing flight compared with air. PMID:26936640

X-37 Storable Propulsion System Design and Operations

NASA Technical Reports Server (NTRS)

Rodriguez, Henry; Popp, Chris; Rehagen, Ronald J.

2005-01-01

In a response to NASA's X-37 TA-10 Cycle-1 contract, Boeing assessed nitrogen tetroxide (N2O4) and monomethyl hydrazine (MMH) Storable Propellant Propulsion Systems to select a low risk X-37 propulsion development approach. Space Shuttle lessons learned, planetary spacecraft, and Boeing Satellite HS-601 systems were reviewed to arrive at a low risk and reliable storable propulsion system. This paper describes the requirements, trade studies, design solutions, flight and ground operational issues which drove X-37 toward the selection of a storable propulsion system. The design of storable propulsion systems offers the leveraging of hardware experience that can accelerate progress toward critical design. It also involves the experience gained from launching systems using MMH and N2O4 propellants. Leveraging of previously flight-qualified hardware may offer economic benefits and may reduce risk in cost and schedule. This paper summarizes recommendations based on experience gained from Space Shuttle and similar propulsion systems utilizing MMH and N2O4 propellants. System design insights gained from flying storable propulsion are presented and addressed in the context of the design approach of the X-37 propulsion system.

X-37 Storable Propulsion System Design and Operations

NASA Technical Reports Server (NTRS)

Rodriguez, Henry; Popp, Chris; Rehegan, Ronald J.

2006-01-01

In a response to NASA's X-37 TA-10 Cycle-1 contract, Boeing assessed nitrogen tetroxide (N2O4) and monomethyl hydrazine (MMH) Storable Propellant Propulsion Systems to select a low risk X-37 propulsion development approach. Space Shuttle lessons learned, planetary spacecraft, and Boeing Satellite HS-601 systems were reviewed to arrive at a low risk and reliable storable propulsion system. This paper describes the requirements, trade studies, design solutions, flight and ground operational issues which drove X-37 toward the selection of a storable propulsion system. The design of storable propulsion systems offers the leveraging of hardware experience that can accelerate progress toward critical design. It also involves the experience gained from launching systems using MMH and N2O4 propellants. Leveraging of previously flight-qualified hardware may offer economic benefits and may reduce risk in cost and schedule. This paper summarizes recommendations based on experience gained from Space Shuttle and similar propulsion systems utilizing MMH and N2O4 propellants. System design insights gained from flying storable propulsion are presented and addressed in the context of the design approach of the X-37 propulsion system.

23. Photographic copy of an asconstructed site plan for North ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

23. Photographic copy of an as-constructed site plan for North Base: Air Force Flight Test Center, Edwards Air Force Base, Edwards, California: North Base Site Plan, February 1970. This drawing shows the North Base building distribution substantially as it appears in 1995. Records on file at AFFTC/CE-CECC-B (Design/Construction Flight/RPMC), Edwards AFB, California. - Edwards Air Force Base, North Base, North Base Road, Boron, Kern County, CA

Baroclinic instability with variable static stability - A design study for a spherical atmospheric model experiment. [for Spacelab flight

NASA Technical Reports Server (NTRS)

Giere, A. C.; Fowlis, W. W.

1980-01-01

The effect of a radially-variable, dielectric body force, analogous to gravity on baroclinic instability for the design of a spherical, synoptic-scale, atmospheric model experiment in a Spacelab flight is investigated. Exact solutions are examined for quasi-geostrophic baroclinic instability in which the rotational Froude number is a linear function of the height. Flow in a rotating rectilinear channel with a vertically variable body force without horizontal shear of the basic state is also discussed.

Deep Space Test Bed for Radiation Studies

NASA Technical Reports Server (NTRS)

Adams, James H.; Adcock, Leonard; Apple, Jeffery; Christl, Mark; Cleveand, William; Cox, Mark; Dietz, Kurt; Ferguson, Cynthia; Fountain, Walt; Ghita, Bogdan

2006-01-01

The Deep Space Test-Bed (DSTB) Facility is designed to investigate the effects of galactic cosmic rays on crews and systems during missions to the Moon or Mars. To gain access to the interplanetary ionizing radiation environment the DSTB uses high-altitude polar balloon flights. The DSTB provides a platform for measurements to validate the radiation transport codes that are used by NASA to calculate the radiation environment within crewed space systems. It is also designed to support other Exploration related investigations such as measuring the shielding effectiveness of candidate spacecraft and habitat materials, testing new radiation monitoring instrumentation and flight avionics and investigating the biological effects of deep space radiation. We describe the work completed thus far in the development of the DSTB and its current status.

M2-F1 in flight

NASA Technical Reports Server (NTRS)

1965-01-01

The M2-F1 Lifting Body is seen here under tow, high above Rogers Dry Lake near the Flight Research Center (later redesignated the Dryden Flight Research Center), Edwards, California. R. Dale Reed effectively advocated the project with the support of NASA research pilot Milt Thompson. Together, they gained the support of Flight Research Center Director Paul Bikle. After a six-month feasibility study, Bikle gave approval in the fall of 1962 for the M2-F1 to be built. The wingless, lifting body aircraft design was initially concieved as a means of landing an aircraft horizontally after atmospheric reentry. The absence of wings would make the extreme heat of re-entry less damaging to the vehicle. In 1962, Flight Research Center management approved a program to build a lightweight, unpowered lifting body as a prototype to flight test the wingless concept. It would look like a 'flying bathtub,' and was designated the M2-F1, the 'M' referring to 'manned' and 'F' referring to 'flight' version. It featured a plywood shell placed over a tubular steel frame crafted at Dryden. Construction was completed in 1963. The first flight tests of the M2-F1 were over Rogers Dry Lake at the end of a tow rope attached to a hopped-up Pontiac convertible driven at speeds up to about 120 mph. These initial tests produced enough flight data about the M2-F1 to proceed with flights behind a NASA C-47 tow plane at greater altitudes. The C-47 took the craft to an altitude of 12,000 where free flights back to Rogers Dry Lake began. Pilot for the first series of flights of the M2-F1 was NASA research pilot Milt Thompson. Typical glide flights with the M2-F1 lasted about two minutes and reached speeds of 110 to l20 mph. More than 400 ground tows and 77 aircraft tow flights were carried out with the M2-F1. The success of Dryden's M2-F1 program led to NASA's development and construction of two heavyweight lifting bodies based on studies at NASA's Ames and Langley research centers--the M2-F2 and the HL-10, both built by the Northrop Corporation, and the U.S. Air Force's X-24 program. The Lifting Body program also heavily influenced the Space Shuttle program. The M2-F1 program demonstrated the feasibility of the lifting body concept for horizontal landings of atmospheric entry vehicles. It also demonstrated a procurement and management concept for prototype flight research vehicles that produced rapid results at very low cost (approximately $50,000, excluding salaries of government employees assigned to the project).

Development of An Intelligent Flight Propulsion Control System

NASA Technical Reports Server (NTRS)

Calise, A. J.; Rysdyk, R. T.; Leonhardt, B. K.

1999-01-01

The initial design and demonstration of an Intelligent Flight Propulsion and Control System (IFPCS) is documented. The design is based on the implementation of a nonlinear adaptive flight control architecture. This initial design of the IFPCS enhances flight safety by using propulsion sources to provide redundancy in flight control. The IFPCS enhances the conventional gain scheduled approach in significant ways: (1) The IFPCS provides a back up flight control system that results in consistent responses over a wide range of unanticipated failures. (2) The IFPCS is applicable to a variety of aircraft models without redesign and,(3) significantly reduces the laborious research and design necessary in a gain scheduled approach. The control augmentation is detailed within an approximate Input-Output Linearization setting. The availability of propulsion only provides two control inputs, symmetric and differential thrust. Earlier Propulsion Control Augmentation (PCA) work performed by NASA provided for a trajectory controller with pilot command input of glidepath and heading. This work is aimed at demonstrating the flexibility of the IFPCS in providing consistency in flying qualities under a variety of failure scenarios. This report documents the initial design phase where propulsion only is used. Results confirm that the engine dynamics and associated hard nonlineaaities result in poor handling qualities at best. However, as demonstrated in simulation, the IFPCS is capable of results similar to the gain scheduled designs of the NASA PCA work. The IFPCS design uses crude estimates of aircraft behaviour. The adaptive control architecture demonstrates robust stability and provides robust performance. In this work, robust stability means that all states, errors, and adaptive parameters remain bounded under a wide class of uncertainties and input and output disturbances. Robust performance is measured in the quality of the tracking. The results demonstrate the flexibility of the IFPCS architecture and the ability to provide robust performance under a broad range of uncertainty. Robust stability is proved using Lyapunov like analysis. Future development of the IFPCS will include integration of conventional control surfaces with the use of propulsion augmentation, and utilization of available lift and drag devices, to demonstrate adaptive control capability under a greater variety of failure scenarios. Further work will specifically address the effects of actuator saturation.

Design and evaluation of a Stochastic Optimal Feed-forward and Feedback Technology (SOFFT) flight control architecture

NASA Technical Reports Server (NTRS)

Ostroff, Aaron J.; Proffitt, Melissa S.

1994-01-01

This paper describes the design and evaluation of a stochastic optimal feed-forward and feedback technology (SOFFT) control architecture with emphasis on the feed-forward controller design. The SOFFT approach allows the designer to independently design the feed-forward and feedback controllers to meet separate objectives and then integrate the two controllers. The feed-forward controller has been integrated with an existing high-angle-of-attack (high-alpha) feedback controller. The feed-forward controller includes a variable command model with parameters selected to satisfy level 1 flying qualities with a high-alpha adjustment to achieve desired agility guidelines, a nonlinear interpolation approach that scales entire matrices for approximation of the plant model, and equations for calculating feed-forward gains developed for perfect plant-model tracking. The SOFFT design was applied to a nonlinear batch simulation model of an F/A-18 aircraft modified for thrust vectoring. Simulation results show that agility guidelines are met and that the SOFFT controller filters undesired pilot-induced frequencies more effectively during a tracking task than a flight controller that has the same feedback control law but does not have the SOFFT feed-forward control.

Gain Scheduling for the Orion Launch Abort Vehicle Controller

NASA Technical Reports Server (NTRS)

McNamara, Sara J.; Restrepo, Carolina I.; Madsen, Jennifer M.; Medina, Edgar A.; Proud, Ryan W.; Whitley, Ryan J.

2011-01-01

One of NASAs challenges for the Orion vehicle is the control system design for the Launch Abort Vehicle (LAV), which is required to abort safely at any time during the atmospheric ascent portion of ight. The focus of this paper is the gain design and scheduling process for a controller that covers the wide range of vehicle configurations and flight conditions experienced during the full envelope of potential abort trajectories from the pad to exo-atmospheric flight. Several factors are taken into account in the automation process for tuning the gains including the abort effectors, the environmental changes and the autopilot modes. Gain scheduling is accomplished using a linear quadratic regulator (LQR) approach for the decoupled, simplified linear model throughout the operational envelope in time, altitude and Mach number. The derived gains are then implemented into the full linear model for controller requirement validation. Finally, the gains are tested and evaluated in a non-linear simulation using the vehicles ight software to ensure performance requirements are met. An overview of the LAV controller design and a description of the linear plant models are presented. Examples of the most significant challenges with the automation of the gain tuning process are then discussed. In conclusion, the paper will consider the lessons learned through out the process, especially in regards to automation, and examine the usefulness of the gain scheduling tool and process developed as applicable to non-Orion vehicles.

Hyper-X Engine Design and Ground Test Program

NASA Technical Reports Server (NTRS)

Voland, R. T.; Rock, K. E.; Huebner, L. D.; Witte, D. W.; Fischer, K. E.; McClinton, C. R.

1998-01-01

The Hyper-X Program, NASA's focused hypersonic technology program jointly run by NASA Langley and Dryden, is designed to move hypersonic, air-breathing vehicle technology from the laboratory environment to the flight environment, the last stage preceding prototype development. The Hyper-X research vehicle will provide the first ever opportunity to obtain data on an airframe integrated supersonic combustion ramjet propulsion system in flight, providing the first flight validation of wind tunnel, numerical and analytical methods used for design of these vehicles. A substantial portion of the integrated vehicle/engine flowpath development, engine systems verification and validation and flight test risk reduction efforts are experimentally based, including vehicle aeropropulsive force and moment database generation for flight control law development, and integrated vehicle/engine performance validation. The Mach 7 engine flowpath development tests have been completed, and effort is now shifting to engine controls, systems and performance verification and validation tests, as well as, additional flight test risk reduction tests. The engine wind tunnel tests required for these efforts range from tests of partial width engines in both small and large scramjet test facilities, to tests of the full flight engine on a vehicle simulator and tests of a complete flight vehicle in the Langley 8-Ft. High Temperature Tunnel. These tests will begin in the summer of 1998 and continue through 1999. The first flight test is planned for early 2000.

Evaluation of a bar-code system to detect unaccompanied baggage

DOT National Transportation Integrated Search

1988-02-01

The objective of the Unaccompanied Baggage Detection System (UBDS) Project has : been to gain field experience with a system designed to identify passengers who : check baggage for a flight and subsequently fail to board that flight. In the first : p...

Quasi-steady aerodynamic model of clap-and-fling flapping MAV and validation using free-flight data.

PubMed

Armanini, S F; Caetano, J V; Croon, G C H E de; Visser, C C de; Mulder, M

2016-06-30

Flapping-wing aerodynamic models that are accurate, computationally efficient and physically meaningful, are challenging to obtain. Such models are essential to design flapping-wing micro air vehicles and to develop advanced controllers enhancing the autonomy of such vehicles. In this work, a phenomenological model is developed for the time-resolved aerodynamic forces on clap-and-fling ornithopters. The model is based on quasi-steady theory and accounts for inertial, circulatory, added mass and viscous forces. It extends existing quasi-steady approaches by: including a fling circulation factor to account for unsteady wing-wing interaction, considering real platform-specific wing kinematics and different flight regimes. The model parameters are estimated from wind tunnel measurements conducted on a real test platform. Comparison to wind tunnel data shows that the model predicts the lift forces on the test platform accurately, and accounts for wing-wing interaction effectively. Additionally, validation tests with real free-flight data show that lift forces can be predicted with considerable accuracy in different flight regimes. The complete parameter-varying model represents a wide range of flight conditions, is computationally simple, physically meaningful and requires few measurements. It is therefore potentially useful for both control design and preliminary conceptual studies for developing new platforms.

Efficient Global Aerodynamic Modeling from Flight Data

NASA Technical Reports Server (NTRS)

Morelli, Eugene A.

2012-01-01

A method for identifying global aerodynamic models from flight data in an efficient manner is explained and demonstrated. A novel experiment design technique was used to obtain dynamic flight data over a range of flight conditions with a single flight maneuver. Multivariate polynomials and polynomial splines were used with orthogonalization techniques and statistical modeling metrics to synthesize global nonlinear aerodynamic models directly and completely from flight data alone. Simulation data and flight data from a subscale twin-engine jet transport aircraft were used to demonstrate the techniques. Results showed that global multivariate nonlinear aerodynamic dependencies could be accurately identified using flight data from a single maneuver. Flight-derived global aerodynamic model structures, model parameter estimates, and associated uncertainties were provided for all six nondimensional force and moment coefficients for the test aircraft. These models were combined with a propulsion model identified from engine ground test data to produce a high-fidelity nonlinear flight simulation very efficiently. Prediction testing using a multi-axis maneuver showed that the identified global model accurately predicted aircraft responses.

Aeroservoelastic stabilization technique refinement for hypersonic flight vehicles

NASA Technical Reports Server (NTRS)

Cheng, Peter Y.; Chan, Samuel Y.; Myers, Thomas T.; Klyde, David H.; Mcruer, Duane T.

1992-01-01

Conventional gain-stabilization techniques introduce low frequency effective time delays which can be troublesome from the viewpoint of SSTOV vehicles' flying qualities. These time delays can be alleviated through a blending of gain-stabilization and phase-stabilization techniques; the resulting hybrid phase stabilization (HPS) for the low-frequency structural modes has been noted to have greater residual response than a conventional gain-stabilizer design. HPS design procedures are presently refined, and residual response metrics are developed.

Launch vehicle tracking enhancement through Global Positioning System Metric Tracking

NASA Astrophysics Data System (ADS)

Moore, T. C.; Li, Hanchu; Gray, T.; Doran, A.

United Launch Alliance (ULA) initiated operational flights of both the Atlas V and Delta IV launch vehicle families in 2002. The Atlas V and Delta IV launch vehicles were developed jointly with the US Air Force (USAF) as part of the Evolved Expendable Launch Vehicle (EELV) program. Both Launch Vehicle (LV) families have provided 100% mission success since their respective inaugural launches and demonstrated launch capability from both Vandenberg Air Force Base (VAFB) on the Western Test Range and Cape Canaveral Air Force Station (CCAFS) on the Eastern Test Range. However, the current EELV fleet communications, tracking, & control architecture & technology, which date back to the origins of the space launch business, require support by a large and high cost ground footprint. The USAF has embarked on an initiative known as Future Flight Safety System (FFSS) that will significantly reduce Test Range Operations and Maintenance (O& M) cost by closing facilities and decommissioning ground assets. In support of the FFSS, a Global Positioning System Metric Tracking (GPS MT) System based on the Global Positioning System (GPS) satellite constellation has been developed for EELV which will allow both Ranges to divest some of their radar assets. The Air Force, ULA and Space Vector have flown the first 2 Atlas Certification vehicles demonstrating the successful operation of the GPS MT System. The first Atlas V certification flight was completed in February 2012 from CCAFS, the second Atlas V certification flight from VAFB was completed in September 2012 and the third certification flight on a Delta IV was completed October 2012 from CCAFS. The GPS MT System will provide precise LV position, velocity and timing information that can replace ground radar tracking resource functionality. The GPS MT system will provide an independent position/velocity S-Band telemetry downlink to support the current man-in-the-loop ground-based commanded destruct of an anomalous flight- The system utilizes a 50 channel digital receiver capable of navigating in high dynamic environments and high altitudes fed by antennas mounted diametrically opposed on the second stage airframe skin. To enhance cost effectiveness, the GPS MT System design implemented existing commercial parts and common environmental and interface requirements for both EELVs. The EELV GPS MT System design is complete, successfully qualified and has demonstrated that the system performs as simulated. This paper summarizes the current development status, system cost comparison, and performance capabilities of the EELV GPS MT System.

Thermal Test Verification of Emission Control through Directional Baffles for the James Webb Space Telescope

NASA Technical Reports Server (NTRS)

Garrison, Matthew; Rashford, Robert; Switzer, Timothy; Shaw, David; White, Bryant; Lynch, Michael; Huber, Frank; Bachtell, Neal

2009-01-01

The thermal performance of NASA s planned James Webb Space Telescope is highly reliant on a collection of directional baffles that are part of the Integrated Science Instrument Module Electronics Compartment. In order to verify the performance of the baffle concept, two test assemblies were recently fabricated and tested at the Goddard Space Flight Center. The centerpiece of the testing was a fixture that used bolometers to measure the emission field through the baffles while the radiator panels and baffles ran a flight-like temperature. Although not all test goals were able to be met due to facility malfunctions, the test was able to prove the design viability enough to gain approval to begin manufacturing the flight article.

Fault tolerant control laws

NASA Technical Reports Server (NTRS)

Ly, U. L.; Ho, J. K.

1986-01-01

A systematic procedure for the synthesis of fault tolerant control laws to actuator failure has been presented. Two design methods were used to synthesize fault tolerant controllers: the conventional LQ design method and a direct feedback controller design method SANDY. The latter method is used primarily to streamline the full-state Q feedback design into a practical implementable output feedback controller structure. To achieve robustness to control actuator failure, the redundant surfaces are properly balanced according to their control effectiveness. A simple gain schedule based on the landing gear up/down logic involving only three gains was developed to handle three design flight conditions: Mach .25 and Mach .60 at 5000 ft and Mach .90 at 20,000 ft. The fault tolerant control law developed in this study provides good stability augmentation and performance for the relaxed static stability aircraft. The augmented aircraft responses are found to be invariant to the presence of a failure. Furthermore, single-loop stability margins of +6 dB in gain and +30 deg in phase were achieved along with -40 dB/decade rolloff at high frequency.

Hyper-X Mach 7 Scramjet Design, Ground Test and Flight Results

NASA Technical Reports Server (NTRS)

Ferlemann, Shelly M.; McClinton, Charles R.; Rock, Ken E.; Voland, Randy T.

2005-01-01

The successful Mach 7 flight test of the Hyper-X (X-43) research vehicle has provided the major, essential demonstration of the capability of the airframe integrated scramjet engine. This flight was a crucial first step toward realizing the potential for airbreathing hypersonic propulsion for application to space launch vehicles. However, it is not sufficient to have just achieved a successful flight. The more useful knowledge gained from the flight is how well the prediction methods matched the actual test results in order to have confidence that these methods can be applied to the design of other scramjet engines and powered vehicles. The propulsion predictions for the Mach 7 flight test were calculated using the computer code, SRGULL, with input from computational fluid dynamics (CFD) and wind tunnel tests. This paper will discuss the evolution of the Mach 7 Hyper-X engine, ground wind tunnel experiments, propulsion prediction methodology, flight results and validation of design methods.

32 CFR 705.32 - Aviation events and parachute demonstrations.

Code of Federal Regulations, 2012 CFR

2012-07-01

... the officially designated military flight or parachute demonstration teams, flyover by aircraft, a... of military flight and parachute demonstration teams. Armed Forces recruiting teams are available to... support of air show fund raising efforts in the form of provision of military flight and parachute...

32 CFR 705.32 - Aviation events and parachute demonstrations.

Code of Federal Regulations, 2014 CFR

2014-07-01

... the officially designated military flight or parachute demonstration teams, flyover by aircraft, a... of military flight and parachute demonstration teams. Armed Forces recruiting teams are available to... support of air show fund raising efforts in the form of provision of military flight and parachute...

32 CFR 705.32 - Aviation events and parachute demonstrations.

Code of Federal Regulations, 2013 CFR

2013-07-01

... the officially designated military flight or parachute demonstration teams, flyover by aircraft, a... of military flight and parachute demonstration teams. Armed Forces recruiting teams are available to... support of air show fund raising efforts in the form of provision of military flight and parachute...

32 CFR 705.32 - Aviation events and parachute demonstrations.

Code of Federal Regulations, 2010 CFR

2010-07-01

... the officially designated military flight or parachute demonstration teams, flyover by aircraft, a... of military flight and parachute demonstration teams. Armed Forces recruiting teams are available to... support of air show fund raising efforts in the form of provision of military flight and parachute...

32 CFR 705.32 - Aviation events and parachute demonstrations.

Code of Federal Regulations, 2011 CFR

2011-07-01

... the officially designated military flight or parachute demonstration teams, flyover by aircraft, a... of military flight and parachute demonstration teams. Armed Forces recruiting teams are available to... support of air show fund raising efforts in the form of provision of military flight and parachute...

Aggressive behavior and performance in the Tegu lizard Tupinambis merianae.

PubMed

Herrel, Anthony; Andrade, Denis V; de Carvalho, José Eduardo; Brito, Ananda; Abe, Augusto; Navas, Carlos

2009-01-01

Aggression is an important component of behavior in many animals and may be crucial to providing individuals with a competitive advantage when resources are limited. Although much is known about the effects of catecholamines and hormones on aggression, relatively few studies have examined the effects of physical performance on aggression. Here we use a large, sexually dimorphic teiid lizard to test whether individuals that show high levels of physical performance (bite force) are also more aggressive toward a potential threat (i.e., a human approaching the lizard). Our results show that independent of their sex, larger individuals with higher bite forces were indeed more aggressive. Moreover, our data show that individuals with higher bite forces tend to show decreased escape responses and are slower, providing evidence for a trade-off between fight and flight abilities. As bite force increased dramatically with body size, we suggest that large body size and bite force may reduce the threshold for an individual to engage in an aggressive encounter, allowing it to potentially gain or maintain resources and fight off predators while minimizing the risk of injury.

Navier-Stokes Simulation of UH-60A Rotor/Wake Interaction Using Adaptive Mesh Refinement

NASA Technical Reports Server (NTRS)

Chaderjian, Neal M.

2017-01-01

Time-dependent Navier-Stokes simulations have been carried out for a flexible UH-60A rotor in forward flight, where the rotor wake interacts with the rotor blades. These flow conditions involved blade vortex interaction and dynamic stall, two common conditions that occur as modern helicopter designs strive to achieve greater flight speeds and payload capacity. These numerical simulations utilized high-order spatial accuracy and delayed detached eddy simulation. Emphasis was placed on understanding how improved rotor wake resolution affects the prediction of the normal force, pitching moment, and chord force of the rotor. Adaptive mesh refinement was used to highly resolve the turbulent rotor wake in a computationally efficient manner. Moreover, blade vortex interaction was found to trigger dynamic stall. Time-dependent flow visualization was utilized to provide an improved understanding of the numerical and physical mechanisms involved with three-dimensional dynamic stall.

Propulsion Flight-Test Fixture

NASA Technical Reports Server (NTRS)

Palumbo, Nate; Vachon, M. Jake; Richwine, Dave; Moes, Tim; Creech, Gray

2003-01-01

NASA Dryden Flight Research Center s new Propulsion Flight Test Fixture (PFTF), designed in house, is an airborne engine-testing facility that enables engineers to gather flight data on small experimental engines. Without the PFTF, it would be necessary to obtain such data from traditional wind tunnels, ground test stands, or laboratory test rigs. Traditionally, flight testing is reserved for the last phase of engine development. Generally, engines that embody new propulsion concepts are not put into flight environments until their designs are mature: in such cases, either vehicles are designed around the engines or else the engines are mounted in or on missiles. However, a captive carry capability of the PFTF makes it possible to test engines that feature air-breathing designs (for example, designs based on the rocket-based combined cycle) economically in subscale experiments. The discovery of unknowns made evident through flight tests provides valuable information to engine designers early in development, before key design decisions are made, thereby potentially affording large benefits in the long term. This is especially true in the transonic region of flight (from mach 0.9 to around 1.2), where it can be difficult to obtain data from wind tunnels and computational fluid dynamics. In January 2002, flight-envelope expansion to verify the design and capabilities of the PFTF was completed. The PFTF was flown on a specially equipped supersonic F-15B research testbed airplane, mounted on the airplane at a center-line attachment fixture, as shown in Figure 1. NASA s F-15B testbed has been used for several years as a flight-research platform. Equipped with extensive research air-data, video, and other instrumentation systems, the airplane carries externally mounted test articles. Traditionally, the majority of test articles flown have been mounted at the centerline tank-attachment fixture, which is a hard-point (essentially, a standardized weapon-mounting fixture). This hard-point has large weight margins, and, because it is located near the center of gravity of the airplane, the weight of equipment mounted there exerts a minimal effect on the stability and controllability of the airplane. The PFTF (see Figure 2) includes a one-piece aluminum structure that contains space for instrumentation, propellant tanks, and feed-system components. The PFTF also houses a force balance, on which is mounted the subscale engine or other experimental apparatus that is to be the subject of a flight test. The force balance measures a combination of inertial and aerodynamic forces and moments acting on the experimental apparatus.

X-15 flight crew - Engle, Rushworth, McKay, Knight, Thompson, and Dana

NASA Technical Reports Server (NTRS)

1966-01-01

The X-15 flight crew, left to right; Air Force Captain Joseph H. Engle, Air Force Major Robert A. Rushworth, NASA pilot John B. 'Jack' McKay, Air Force Major William J. 'Pete' Knight, NASA pilot Milton O. Thompson, and NASA pilot Bill Dana. These six pilots made 125 of the 199 total flights in the X-15. Rushworth made 34 flights (the most of any X-15 pilot); McKay flew 29 times; Engle, Knight, and Dana each flew 16 times; Thompson's total was 14. The X-15 was a rocket-powered aircraft 50 ft long with a wingspan of 22 ft. It was a missile-shaped vehicle with an unusual wedge-shaped vertical tail, thin stubby wings, and unique fairings that extended along the side of the fuselage. The X-15 weighed about 14,000 lb empty and approximately 34,000 lb at launch. The XLR-99 rocket engine, manufactured by Thiokol Chemical Corp., was pilot controlled and was capable of developing 57,000 lb of rated thrust (actual thrust reportedly climbed to 60,000 lb). North American Aviation built three X-15 aircraft for the program. The X-15 research aircraft was developed to provide in-flight information and data on aerodynamics, structures, flight controls, and the physiological aspects of high-speed, high-altitude flight. A follow-on program used the aircraft as a testbed to carry various scientific experiments beyond the Earth's atmosphere on a repeated basis. For flight in the dense air of the usable atmosphere, the X-15 used conventional aerodynamic controls such as rudder surfaces on the vertical stabilizers to control yaw and canted horizontal surfaces on the tail to control pitch when moving in synchronization or roll when moved differentially. For flight in the thin air outside of the appreciable Earth's atmosphere, the X-15 used a reaction control system. Hydrogen peroxide thrust rockets located on the nose of the aircraft provided pitch and yaw control. Those on the wings provided roll control. Because of the large fuel consumption, the X-15 was air launched from a B-52 aircraft at 45,000 ft and a speed of about 500 mph. Depending on the mission, the rocket engine provided thrust for the first 80 to 120 sec of flight. The remainder of the normal 10 to 11 min. flight was powerless and ended with a 200-mph glide landing. Generally, one of two types of X-15 flight profiles was used: a high-altitude flight plan that called for the pilot to maintain a steep rate of climb, or a speed profile that called for the pilot to push over and maintain a level altitude. The X-15 was flown over a period of nearly 10 years--June 1959 to Oct. 1968--and set the world's unofficial speed and altitude records of 4,520 mph (Mach 6.7) and 354,200 ft (over 67 mi) in a program to investigate all aspects of piloted hypersonic flight. Information gained from the highly successful X-15 program contributed to the development of the Mercury, Gemini, and Apollo manned spaceflight programs, and also the Space Shuttle program. The X-15s made a total of 199 flights and were manufactured by North American Aviation. X-15-1, serial number 56-6670, is now located at the National Air and Space Museum, Washington DC. North American X-15A-2, serial number 56-6671, is at the United States Air Force Museum, Wright-Patterson AFB, Ohio. The X-15-3, serial number 56-6672, crashed on 15 November 1967, resulting in the death of Maj. Michael J. Adams.

Summary of Results from the Risk Management Program for the Mars Microrover Flight Experiment

NASA Technical Reports Server (NTRS)

Shishko, Robert; Matijevic, Jacob R.

2000-01-01

On 4 July 1997, the Mars Pathfinder landed on the surface of Mars carrying the first planetary rover, known as the Sojourner. Formally known as the Microrover Flight Experiment (MFEX), the Sojourner was a low cost, high-risk technology demonstration, in which new risk management techniques were tried. This paper summarizes the activities and results of the effort to conduct a low-cost, yet meaningful risk management program for the MFEX. The specific activities focused on cost, performance, schedule, and operations risks. Just as the systems engineering process was iterative and produced successive refinements of requirements, designs, etc., so was the risk management process. Qualitative risk assessments were performed first to gain some insights for refining the microrover design and operations concept. These then evolved into more quantitative analyses. Risk management lessons from the manager's perspective is presented for other low-cost, high-risk space missions.

Highly accurate analytic formulae for projectile motion subjected to quadratic drag

NASA Astrophysics Data System (ADS)

Turkyilmazoglu, Mustafa

2016-05-01

The classical phenomenon of motion of a projectile fired (thrown) into the horizon through resistive air charging a quadratic drag onto the object is revisited in this paper. No exact solution is known that describes the full physical event under such an exerted resistance force. Finding elegant analytical approximations for the most interesting engineering features of dynamical behavior of the projectile is the principal target. Within this purpose, some analytical explicit expressions are derived that accurately predict the maximum height, its arrival time as well as the flight range of the projectile at the highest ascent. The most significant property of the proposed formulas is that they are not restricted to the initial speed and firing angle of the object, nor to the drag coefficient of the medium. In combination with the available approximations in the literature, it is possible to gain information about the flight and complete the picture of a trajectory with high precision, without having to numerically simulate the full governing equations of motion.

Shuttle Risk Progression by Flight

NASA Technical Reports Server (NTRS)

Hamlin, Teri; Kahn, Joe; Thigpen, Eric; Zhu, Tony; Lo, Yohon

2011-01-01

Understanding the early mission risk and progression of risk as a vehicle gains insights through flight is important: . a) To the Shuttle Program to understand the impact of re-designs and operational changes on risk. . b) To new programs to understand reliability growth and first flight risk. . Estimation of Shuttle Risk Progression by flight: . a) Uses Shuttle Probabilistic Risk Assessment (SPRA) and current knowledge to calculate early vehicle risk. . b) Shows impact of major Shuttle upgrades. . c) Can be used to understand first flight risk for new programs.

NASA B737 flight test results of the total energy control system

NASA Technical Reports Server (NTRS)

Bruce, Kevin R.

1987-01-01

The Total Energy Control System (TECS) is an integrated autopilot/autothrottle developed by BCAC that was test flown on NASA Langley's Transport System Research Vehicle (i.e., a highly modified Boeing B737). This systems was developed using principles of total energy in which the total kinetic and potential energy of the airplane was controlled by the throttles, and the energy distribution controled by the elevator. TECS integrates all the control functions of a conventional pitch autopilot and autothrottle into a single generalized control concept. This integration provides decoupled flightpath and maneuver control, as well as a coordinated throttle response for all maneuvers. A mode hierarchy was established to preclude exceeding airplane safety and performance limits. The flight test of TECS took place as a series of five flights over a 33-week period during September 1985 at NASA Langley. Most of the original flight test plan was completed within the first three flights with the system not exhibiting any instabilities or design problems that required any gain adjustment during flight.

Design and test of the 172K fluidic rudder

NASA Technical Reports Server (NTRS)

Belsterling, C. A.

1978-01-01

Progress in the development of concepts for control of aircraft without moving parts or a separate source of power is described. The design and wind tunnel tests of a full scale fluidic rudder for a Cessna 172K aircraft, intended for subsequent flight tests were documented. The 172K fluidic rudder was designed to provide a control force equivalent to 3.3 degrees of deflection of the conventional rudder. In spite of an extremely thin airfoil, cascaded fluidic amplifiers were built to fit, with the capacity for generating the required level of control force. Wind tunnel tests demonstrated that the principles of lift control using ram air power are sound and reliable under all flight conditions. The tests also demonstrated that the performance of the 172K fluidic rudder is not acceptable for flight tests until the design of the scoop is modified to prevent interference with the lift control phenomenon.

Functional integration of vertical flight path and speed control using energy principles

NASA Technical Reports Server (NTRS)

Lambregts, A. A.

1984-01-01

A generalized automatic flight control system was developed which integrates all longitudinal flight path and speed control functions previously provided by a pitch autopilot and autothrottle. In this design, a net thrust command is computed based on total energy demand arising from both flight path and speed targets. The elevator command is computed based on the energy distribution error between flight path and speed. The engine control is configured to produce the commanded net thrust. The design incorporates control strategies and hierarchy to deal systematically and effectively with all aircraft operational requirements, control nonlinearities, and performance limits. Consistent decoupled maneuver control is achieved for all modes and flight conditions without outer loop gain schedules, control law submodes, or control function duplication.

X-15 #2 on lakebed after engine failure forced pilot Jack McKay to make an emergency landing at Mud

NASA Technical Reports Server (NTRS)

1962-01-01

On 9 November 1962, an engine failure forced Jack McKay, a NASA research pilot, to make an emergency landing at Mud Lake, Nevada, in the second X-15 (56-6671); its landing gear collapsed and the X-15 flipped over on its back. McKay was promptly rescued by an Air Force medical team standing by near the launch site, and eventually recovered to fly the X-15 again. But his injuries, more serious than at first thought, eventually forced his retirement from NASA. The aircraft was sent back to the manufacturer, where it underwent extensive repairs and modifications. It returned to Edwards in February 1964 as the X-15A-2, with a longer fuselage (52 ft 5 in) and external fuel tanks. The basic X-15 was a rocket-powered aircraft 50 ft long with a wingspan of 22 ft. It was a missile-shaped vehicle with an unusual wedge-shaped vertical tail, thin stubby wings, and unique side fairings that extended along the side of the fuselage. The X-15 weighed about 14,000 lb empty and approximately 34,000 lb at launch. The XLR-99 rocket engine, manufactured by Thiokol Chemical Corp., was pilot controlled and was capable of developing 57,000 lb of thrust. North American Aviation built three X-15 aircraft for the program. The X-15 research aircraft was developed to provide in-flight information and data on aerodynamics, structures, flight controls, and the physiological aspects of high-speed, high-altitude flight. A follow-on program used the aircraft as a testbed to carry various scientific experiments beyond the Earth's atmosphere on a repeated basis. For flight in the dense air of the usable atmosphere, the X-15 used conventional aerodynamic controls such as rudder surfaces on the vertical stabilizers to control yaw and movable horizontal stabilizers to control pitch when moving in synchronization or roll when moved differentially. For flight in the thin air outside of the appreciable Earth's atmosphere, the X-15 used a reaction control system. Hydrogen peroxide thrust rockets located on the nose of the aircraft provided pitch and yaw control. Those on the wings provided roll control. Because of the large fuel consumption, the X-15 was air launched from a B-52 aircraft at 45,000 ft and a speed of about 500 mph. Depending on the mission, the rocket engine provided thrust for the first 80 to 120 sec of flight. The remainder of the normal 10 to 11 min. flight was powerless and ended with a 200-mph glide landing. Generally, one of two types of X-15 flight profiles was used; a high-altitude flight plan that called for the pilot to maintain a steep rate of climb, or a speed profile that called for the pilot to push over and maintain a level altitude. The X-15 was flown over a period of nearly 10 years--June 1959 to Oct. 1968--and set the world's unofficial speed and altitude records of 4,520 mph (Mach 6.7) and 354,200 ft in a program to investigate all aspects of manned hypersonic flight. Information gained from the highly successful X-15 program contributed to the development of the Mercury, Gemini, and Apollo manned spaceflight programs, and also the Space Shuttle program. The X-15s made a total of 199 flights, and were manufactured by North American Aviation. X-15-1, serial number 56-6670, is now located at the National Air and Space Museum, Washington DC. North American X-15A-2, serial number 56-6671, is at the United States Air Force Museum, Wright-Patterson AFB, Ohio. The X-15-3, serial number 56-6672, crashed on 15 November 1967, resulting in the death of Maj. Michael J. Adams.

Coherence and interlimb force control: Effects of visual gain.

PubMed

Kang, Nyeonju; Cauraugh, James H

2018-03-06

Neural coupling across hemispheres and homologous muscles often appears during bimanual motor control. Force coupling in a specific frequency domain may indicate specific bimanual force coordination patterns. This study investigated coherence on pairs of bimanual isometric index finger force while manipulating visual gain and task asymmetry conditions. We used two visual gain conditions (low and high gain = 8 and 512 pixels/N), and created task asymmetry by manipulating coefficient ratios imposed on the left and right index finger forces (0.4:1.6; 1:1; 1.6:0.4, respectively). Unequal coefficient ratios required different contributions from each hand to the bimanual force task resulting in force asymmetry. Fourteen healthy young adults performed bimanual isometric force control at 20% of their maximal level of the summed force of both fingers. We quantified peak coherence and relative phase angle between hands at 0-4, 4-8, and 8-12 Hz, and estimated a signal-to-noise ratio of bimanual forces. The findings revealed higher peak coherence and relative phase angle at 0-4 Hz than at 4-8 and 8-12 Hz for both visual gain conditions. Further, peak coherence and relative phase angle values at 0-4 Hz were larger at the high gain than at the low gain. At the high gain, higher peak coherence at 0-4 Hz collapsed across task asymmetry conditions significantly predicted greater signal-to-noise ratio. These findings indicate that a greater level of visual information facilitates bimanual force coupling at a specific frequency range related to sensorimotor processing. Copyright © 2018 Elsevier B.V. All rights reserved.

DAST in Flight

NASA Technical Reports Server (NTRS)

1980-01-01

The modified BQM-34 Firebee II drone with Aeroelastic Research Wing (ARW-1), a supercritical airfoil, during a 1980 research flight. The remotely-piloted vehicle, which was air launched from NASA's NB-52B mothership, participated in the Drones for Aerodynamic and Structural Testing (DAST) program which ran from 1977 to 1983. The DAST 1 aircraft (Serial #72-1557), pictured, crashed on 12 June 1980 after its right wing ripped off during a test flight near Cuddeback Dry Lake, California. The crash occurred on the modified drone's third free flight. These are the image contact sheets for each image resolution of the NASA Dryden Drones for Aerodynamic and Structural Testing (DAST) Photo Gallery. From 1977 to 1983, the Dryden Flight Research Center, Edwards, California, (under two different names) conducted the DAST Program as a high-risk flight experiment using a ground-controlled, pilotless aircraft. Described by NASA engineers as a 'wind tunnel in the sky,' the DAST was a specially modified Teledyne-Ryan BQM-34E/F Firebee II supersonic target drone that was flown to validate theoretical predictions under actual flight conditions in a joint project with the Langley Research Center, Hampton, Virginia. The DAST Program merged advances in electronic remote control systems with advances in airplane design. Drones (remotely controlled, missile-like vehicles initially developed to serve as gunnery targets) had been deployed successfully during the Vietnamese conflict as reconnaissance aircraft. After the war, the energy crisis of the 1970s led NASA to seek new ways to cut fuel use and improve airplane efficiency. The DAST Program's drones provided an economical, fuel-conscious method for conducting in-flight experiments from a remote ground site. DAST explored the technology required to build wing structures with less than normal stiffness. This was done because stiffness requires structural weight but ensures freedom from flutter-an uncontrolled, divergent oscillation of the structure, driven by aerodynamic forces and resulting in structural failure. The program used refined theoretical tools to predict at what speed flutter would occur. It then designed a high-response control system to counteract the motion and permit a much lighter wing structure. The wing had, in effect, 'electronic stiffness.' Flight research with this concept was extremely hazardous because an error in either the flutter prediction or control system implementation would result in wing structural failure and the loss of the vehicle. Because of this, flight demonstration of a sub-scale vehicle made sense from the standpoint of both safety and cost. The program anticipated structural failure during the course of the flight research. The Firebee II was a supersonic drone selected as the DAST testbed because its wing could be easily replaced, it used only tail-mounted control surfaces, and it was available as surplus from the U. S. Air Force. It was capable of 5-g turns (that is, turns producing acceleration equal to 5 times that of gravity). Langley outfitted a drone with an aeroelastic, supercritical research wing suitable for a Mach 0.98 cruise transport with a predicted flutter speed of Mach 0.95 at an altitude of 25,000 feet. Dryden and Langley, in conjunction with Boeing, designed and fabricated a digital flutter suppression system (FSS). Dryden developed an RPRV (remotely piloted research vehicle) flight control system; integrated the wing, FSS, and vehicle systems; and conducted the flight program. In addition to a digital flight control system and aeroelastic wings, each DAST drone had research equipment mounted in its nose and a mid-air retrieval system in its tail. The drones were originally launched from the NASA B-52 bomber and later from a DC-130. The DAST vehicle's flight was monitored from the sky by an F-104 chase plane. When the DAST's mission ended, it deployed a parachute and then a specially equipped Air Force helicopter recovered the drone in mid-air. On the ground, a pilot controlled the DAST vehicle from a remote cockpit while researchers in another room monitored flight data transmitted via telemetry. They made decisions on the conduct of the flight while the DAST was in the air. In case of failure in any of the ground systems, the DAST vehicle could also be flown to a recovery site using a backup control system in the F-104. The DAST Program experienced numerous problems. Only eighteen flights were achieved, eight of them captive (in which the aircraft flew only while still attached to the launch aircraft). Four of the flights were aborted and two resulted in crashes--one on June 12, 1980, and the second on June 1, 1983. Meanwhile, flight experiments with higher profiles, better funded remotely piloted research vehicles took priority over DAST missions. After the 1983 crash, which was caused by a malfunction that disconnected the landing parachute from the drone, the program was disbanded. Because DAST drones were considered expendable, certain losses were anticipated. Managers and researchers involved in other high-risk flight projects gained insights from the DAST program that could be applied to their own flight research programs. The DAST aircraft had a wingspan of 14 feet, four inches and a nose-to-tail length of 28 feet, 4 inches. The fuselage had a radius of about 2.07 feet. The aircraft's maximum loaded weight was about 2,200 pounds. It derived its power from a Continental YJ69-T-406 engine.

In-Space technology experiments program. A high efficiency thermal interface (using condensation heat transfer) between a 2-phase fluid loop and heatpipe radiator: Experiment definition phase

NASA Technical Reports Server (NTRS)

Pohner, John A.; Dempsey, Brian P.; Herold, Leroy M.

1990-01-01

Space Station elements and advanced military spacecraft will require rejection of tens of kilowatts of waste heat. Large space radiators and two-phase heat transport loops will be required. To minimize radiator size and weight, it is critical to minimize the temperature drop between the heat source and sink. Under an Air Force contract, a unique, high-performance heat exchanger is developed for coupling the radiator to the transport loop. Since fluid flow through the heat exchanger is driven by capillary forces which are easily dominated by gravity forces in ground testing, it is necessary to perform microgravity thermal testing to verify the design. This contract consists of an experiment definition phase leading to a preliminary design and cost estimate for a shuttle-based flight experiment of this heat exchanger design. This program will utilize modified hardware from a ground test program for the heat exchanger.

EC95-43057-8

NASA Image and Video Library

1995-03-24

Outlined with gold stripes are the hinged nose strakes, modifications made to NASA's F-18 HARV (High Alpha Research Vehicle) at the Dryden Flight Research Center, Edwards, California. Actuated Nose Strakes for Enhanced Rolling (ANSER) were installed to fly the third and final phase in the HARV flight test project. Normally folded flush, the units -- four feet long and six inches wide -- can be opened independently to interact with the nose vortices to produce large side forces for control. Early wind tunnel tests indicated that the strakes would be as effective in yaw control at high angles of attack as rudders are at lower angles. Testing involved evaluation of the strakes by themselves as well as combined with the aircraft's Thrust Vectoring System. The strakes were designed by NASA's Langley Research Center, then installed and flight tested at Dryden.

Integrated flight/propulsion control design for a STOVL aircraft using H-infinity control design techniques

NASA Technical Reports Server (NTRS)

Garg, Sanjay; Ouzts, Peter J.

1991-01-01

Results are presented from an application of H-infinity control design methodology to a centralized integrated flight propulsion control (IFPC) system design for a supersonic Short Takeoff and Vertical Landing (STOVL) fighter aircraft in transition flight. The emphasis is on formulating the H-infinity control design problem such that the resulting controller provides robustness to modeling uncertainties and model parameter variations with flight condition. Experience gained from a preliminary H-infinity based IFPC design study performed earlier is used as the basis to formulate the robust H-infinity control design problem and improve upon the previous design. Detailed evaluation results are presented for a reduced order controller obtained from the improved H-infinity control design showing that the control design meets the specified nominal performance objectives as well as provides stability robustness for variations in plant system dynamics with changes in aircraft trim speed within the transition flight envelope. A controller scheduling technique which accounts for changes in plant control effectiveness with variation in trim conditions is developed and off design model performance results are presented.

Application of a Constant Gain Extended Kalman Filter for In-Flight Estimation of Aircraft Engine Performance Parameters

NASA Technical Reports Server (NTRS)

Kobayashi, Takahisa; Simon, Donald L.; Litt, Jonathan S.

2005-01-01

An approach based on the Constant Gain Extended Kalman Filter (CGEKF) technique is investigated for the in-flight estimation of non-measurable performance parameters of aircraft engines. Performance parameters, such as thrust and stall margins, provide crucial information for operating an aircraft engine in a safe and efficient manner, but they cannot be directly measured during flight. A technique to accurately estimate these parameters is, therefore, essential for further enhancement of engine operation. In this paper, a CGEKF is developed by combining an on-board engine model and a single Kalman gain matrix. In order to make the on-board engine model adaptive to the real engine s performance variations due to degradation or anomalies, the CGEKF is designed with the ability to adjust its performance through the adjustment of artificial parameters called tuning parameters. With this design approach, the CGEKF can maintain accurate estimation performance when it is applied to aircraft engines at offnominal conditions. The performance of the CGEKF is evaluated in a simulation environment using numerous component degradation and fault scenarios at multiple operating conditions.

Error analysis and assessment of unsteady forces acting on a flapping wing micro air vehicle: free flight versus wind-tunnel experimental methods.

PubMed

Caetano, J V; Percin, M; van Oudheusden, B W; Remes, B; de Wagter, C; de Croon, G C H E; de Visser, C C

2015-08-20

An accurate knowledge of the unsteady aerodynamic forces acting on a bio-inspired, flapping-wing micro air vehicle (FWMAV) is crucial in the design development and optimization cycle. Two different types of experimental approaches are often used: determination of forces from position data obtained from external optical tracking during free flight, or direct measurements of forces by attaching the FWMAV to a force transducer in a wind-tunnel. This study compares the quality of the forces obtained from both methods as applied to a 17.4 gram FWMAV capable of controlled flight. A comprehensive analysis of various error sources is performed. The effects of different factors, e.g., measurement errors, error propagation, numerical differentiation, filtering frequency selection, and structural eigenmode interference, are assessed. For the forces obtained from free flight experiments it is shown that a data acquisition frequency below 200 Hz and an accuracy in the position measurements lower than ± 0.2 mm may considerably hinder determination of the unsteady forces. In general, the force component parallel to the fuselage determined by the two methods compares well for identical flight conditions; however, a significant difference was observed for the forces along the stroke plane of the wings. This was found to originate from the restrictions applied by the clamp to the dynamic oscillations observed in free flight and from the structural resonance of the clamped FWMAV structure, which generates loads that cannot be distinguished from the external forces. Furthermore, the clamping position was found to have a pronounced influence on the eigenmodes of the structure, and this effect should be taken into account for accurate force measurements.

X-2 on Transportation Dolly

NASA Technical Reports Server (NTRS)

1952-01-01

This 1952 photograph shows the X-2 #2 aircraft mounted on a special transportation dolly at Edwards Air Force Base, California. The dolly was steerable and was used for transporting the X-2 around and for towing it off the lakebed at Edwards Air Force Base after a landing. This was the number 2 airplane (46-675), which was lost on May 12, 1953, on a captive flight over Lake Ontario when the airplane exploded during a liquid-oxygen topoff test, killing the pilot, Jean Ziegler, and EB-50A crewman Frank Wolko. Almost no debris was recovered from Lake Ontario, so no cause for the explosion could be determined. Later, however, investigations of similar explosions in the X-1 #3, X-1A, and X-1D traced the problem to Ulmer leather gaskets, which exuded tricresyl phosphate. This substance caused detonations in the supercold atmosphere of the airplanes' liquid oxygen tanks. As the X-2 #2 also had these gaskets, they were probably the cause of the explosion in that aircraft as well. The X-2 was a swept-wing, rocket-powered aircraft designed to fly faster than Mach 3 (three times the speed of sound). It was built for the U.S. Air Force by the Bell Aircraft Company, Buffalo, New York. The X-2 was flown to investigate the problems of aerodynamic heating as well as stability and control effectiveness at high altitudes and high speeds (in excess of Mach 3). Bell aircraft built two X-2 aircraft. These were constructed of K-monel (a copper and nickel alloy) for the fuselage and stainless steel for the swept wings and control surfaces. The aircraft had ejectable nose capsules instead of ejection seats because the development of ejection seats had not reached maturity at the time the X-2 was conceived. The X-2 ejection canopy was successfully tested using a German V-2 rocket. The X-2 used a skid-type landing gear to make room for more fuel. The airplane was air launched from a modified Boeing B-50 Superfortress Bomber. X-2 Number 1 made its first unpowered glide flight on Aug. 5, 1954, and made a total of 17 (4 glide and 13 powered) flights before it was lost Sept. 27, 1956. The pilot on Flight 17, Capt. Milburn Apt, had flown the aircraft to a record speed of Mach 3.2 (2,094 mph), thus becoming the first person to exceed Mach 3. During that last flight, inertial coupling occurred and the pilot was killed. The aircraft suffered little damage in the crash, resulting in proposals (never implemented) from the Langley Memorial Aeronautical Laboratory, Hampton, Virginia, to rebuild it for use in a hypersonic (Mach 5+) test program. In 1953, X-2 Number 2 was lost in an in-flight explosion while at the Bell Aircraft Company during captive flight trials and was jettisoned into Lake Ontario. The Air Force had previously flown the aircraft on three glide flights at Edwards Air Force Base, California, in 1952. Although the NACA's High-Speed Flight Station, Edwards, California, (predecessor of NASA's Dryden Flight Research Center) never actually flew the X-2 aircraft, the NACA did support the program primarily through Langley Memorial Aeronautical Laboratory wind-tunnel tests and Wallops Island, Virginia, rocket-model tests. The NACA High-Speed Flight Station also provided stability-and-control recording instrumentation and simulator support for the Air Force flights. In the latter regard, the NACA worked with the Air Force in using a special computer to extrapolate and predict aircraft behavior from flight data.

Two YF-12 aircraft in flight

NASA Technical Reports Server (NTRS)

1975-01-01

The YF-12A (60-6935) carries the 'coldwall' heat transfer pod on a pylon beneath the forward fuselage. The pod is seen with its insulating coating intact. In the foreground, the YF-12C flies photo chase. The coldwall project, supported by Langley Research Center, consisted of a stainless steel tube equipped with thermocouples and pressure-sensors. A special insulating coating covered the tube, which was chilled with liquid nitrogen. At Mach 3, the insulation could be pyrotechnically blown away from the tube, instantly exposing it to the thermal environment. The experiment caused many inflight difficulties, such as engine unstarts, but eventually researchers got a successful flight. The Flight Research Center's involvement with the YF-12A, an interceptor version of the Lockheed A-12, began in 1967. Ames Research Center was interested in using wind tunnel data that had been generated at Ames under extreme secrecy. Also, the Office of Advanced Research and Technology (OART) saw the YF-12A as a means to advance high-speed technology, which would help in designing the Supersonic Transport (SST). The Air Force needed technical assistance to get the latest reconnaissance version of the A-12 family, the SR-71A, fully operational. Eventually, the Air Force offered NASA the use of two YF-12A aircraft, 60-6935 and 60-6936. A joint NASA-USAF program was mapped out in June 1969. NASA and Air Force technicians spent three months readying 935 for flight. On 11 December 1969, the flight program got underway with a successful maiden flight piloted by Col. Joe Rogers and Maj. Gary Heidelbaugh of the SR-71/F-12 Test Force. During the program, the Air Force concentrated on military applications, and NASA pursued a loads research program. NASA studies included inflight heating, skin-friction cooling, 'coldwall' research (a heat transfer experiment), flowfield studies, shaker vane research, and tests in support of the Space Shuttle landing program. Ultimately, 935 became the workhorse of the program, with 146 flights between 11 December 1969 and 7 November 1979. The second YF-12A, 936, made 62 flights. It was lost in a non-fatal crash on 24 June 1971. It was replaced by the so-called YF-12C (SR-71A 61-7951, modified with YF-12A inlets and engines and a bogus tail number 06937). The Lockheed A-12 family, known as the Blackbirds, were designed by Clarence 'Kelly' Johnson. They were constructed mostly of titanium to withstand aerodynamic heating. Fueled by JP-7, the Blackbirds were capable of cruising at Mach 3.2 and attaining altitudes in excess of 80,000 feet. The first version, a CIA reconnaissance aircraft that first flew in April 1962 was called the A-12. An interceptor version was developed in 1963 under the designation YF-12A. A USAF reconnaissance variant, called the SR-71, was first flown in 1964. The A-12 and SR-71 designs included leading and trailing edges made of high-temperature fiberglass-asbestos laminates. The NASA YF-12 research program was ambitious; the aircraft flew an average of once a week unless down for extended maintenance or modification. Program expenses averaged $3.1 million per year just to run the flight tests. NASA crews for the YF-12 included pilots Fitzhugh Fulton and Donald Mallick, anf flight test engineers Victor Horton and Ray Young. Other NASA test pilots checked out in the YF-12A included John Manke, William Dana, Gary Krier, Einar Enevoldson, Tom McMurtry, Steve Ishmael, and Michael Swann. The YF-12C was only flown by Fulton, Mallick, Horton, and Ray.

YF-12A and YF-12C in flight formation at dawn

NASA Technical Reports Server (NTRS)

1975-01-01

The YF-12A (60-6935) carries the 'coldwall' heat transfer pod on a pylon beneath the forward fuselage. The pod is seen with its insulating coating intact. In the background, the YF-12C flies photo chase. The coldwall project, supported by Langley Research Center, consisted of a stainless steel tube equipped with thermocouples and pressure-sensors. A special insulating coating covered the tube, which was chilled with liquid nitrogen. At Mach 3, the insulation could be pyrotechnically blown away from the tube, instantly exposing it to the thermal environment. The experiment caused many inflight difficulties, such as engine unstarts, but eventually researchers got a successful flight. The Flight Research Center's involvement with the YF-12A, an interceptor version of the Lockheed A-12, began in 1967. Ames Research Center was interested in using wind tunnel data that had been generated at Ames under extreme secrecy. Also, the Office of Advanced Research and Technology (OART) saw the YF-12A as a means to advance high-speed technology, which would help in designing the Supersonic Transport (SST). The Air Force needed technical assistance to get the latest reconnaissance version of the A-12 family, the SR-71A, fully operational. Eventually, the Air Force offered NASA the use of two YF-12A aircraft, 60-6935 and 60-6936. A joint NASA-USAF program was mapped out in June 1969. NASA and Air Force technicians spent three months readying 935 for flight. On 11 December 1969, the flight program got underway with a successful maiden flight piloted by Col. Joe Rogers and Maj. Gary Heidelbaugh of the SR-71/F-12 Test Force. During the program, the Air Force concentrated on military applications, and NASA pursued a loads research program. NASA studies included inflight heating, skin-friction cooling, 'coldwall' research (a heat transfer experiment), flowfield studies, shaker vane research, and tests in support of the Space Shuttle landing program. Ultimately, 935 became the workhorse of the program, with 146 flights between 11 December 1969 and 7 November 1979. The second YF-12A, 936, made 62 flights. It was lost in a non-fatal crash on 24 June 1971. It was replaced by the so-called YF-12C (SR-71A 61-7951, modified with YF-12A inlets and engines and a bogus tail number 06937). The Lockheed A-12 family, known as the Blackbirds, were designed by Clarence 'Kelly' Johnson. They were constructed mostly of titanium to withstand aerodynamic heating. Fueled by JP-7, the Blackbirds were capable of cruising at Mach 3.2 and attaining altitudes in excess of 80,000 feet. The first version, a CIA reconnaissance aircraft that first flew in April 1962 was called the A-12. An interceptor version was developed in 1963 under the designation YF-12A. A USAF reconnaissance variant, called the SR-71, was first flown in 1964. The A-12 and SR-71 designs included leading and trailing edges made of high-temperature fiberglass-asbestos laminates. The NASA YF-12 research program was ambitious; the aircraft flew an average of once a week unless down for extended maintenance or modification. Program expenses averaged $3.1 million per year just to run the flight tests. NASA crews for the YF-12 included pilots Fitzhugh Fulton and Donald Mallick, anf flight test engineers Victor Horton and Ray Young. Other NASA test pilots checked out in the YF-12A included John Manke, William Dana, Gary Krier, Einar Enevoldson, Tom McMurtry, Steve Ishmael, and Michael Swann. The YF-12C was only flown by Fulton, Mallick, Horton, and Ray.

Lifting Body Flight Vehicles

NASA Technical Reports Server (NTRS)

Barret, Chris

1998-01-01

NASA has a technology program in place to build the X-33 test vehicle and then the full sized Reusable Launch Vehicle, VentureStar. VentureStar is a Lifting Body (LB) flight vehicle which will carry our future payloads into orbit, and will do so at a much reduced cost. There were three design contenders for the new Reusable Launch Vehicle: a Winged Vehicle, a Vertical Lander, and the Lifting Body(LB). The LB design won the competition. A LB vehicle has no wings and derives its lift solely from the shape of its body, and has the unique advantages of superior volumetric efficiency, better aerodynamic efficiency at high angles-of-attack and hypersonic speeds, and reduced thermal protection system weight. Classically, in a ballistic vehicle, drag has been employed to control the level of deceleration in reentry. In the LB, lift enables the vehicle to decelerate at higher altitudes for the same velocity and defines the reentry corridor which includes a greater cross range. This paper outlines our LB heritage which was utilized in the design of the new Reusable Launch Vehicle, VentureStar. NASA and the U.S. Air Force have a rich heritage of LB vehicle design and flight experience. Eight LB's were built and over 225 LB test flights were conducted through 1975 in the initial LB Program. Three LB series were most significant in the advancement of today's LB technology: the M2-F; HL-1O; and X-24 series. The M2-F series was designed by NASA Ames Research Center, the HL-10 series by NASA Langley Research Center, and the X-24 series by the Air Force. LB vehicles are alive again today.

Automated Aerial Refueling Hitches a Ride on AFF

NASA Technical Reports Server (NTRS)

Hansen, Jennifer L.; Murray, James E.; Bever, Glenn; Campos, Norma V.; Schkolnik, Gerard

2007-01-01

The recent introduction of uninhabited aerial vehicles [UAVs (basically, remotely piloted or autonomous aircraft)] has spawned new developments in autonomous operation and posed new challenges. Automated aerial refueling (AAR) is a capability that will enable UAVs to travel greater distances and loiter longer over targets. NASA Dryden Flight Research Center, in cooperation with the Defense Advanced Research Projects Agency (DARPA), the Naval Air Systems Command (NAVAIR), the Naval Air Force Pacific Fleet, and the Air Force Research Laboratory, rapidly conceived and accomplished an AAR flight research project focused on collecting a unique, high-quality database on the dynamics of the hose and drogue of an aerial refueling system. This flight-derived database would be used to validate mathematical models of the dynamics in support of design and analysis of AAR systems for future UAVs. The project involved the use of two Dryden F/A-18 airplanes and an S-3 hose-drogue refueling store on loan from the Navy. In this year-long project, which was started on October 1, 2002, 583 research maneuvers were completed during 23 flights.

Flight demonstration of a self repairing flight control system in a NASA F-15 fighter aircraft

NASA Technical Reports Server (NTRS)

Urnes, James M.; Stewart, James; Eslinger, Robert

1990-01-01

Battle damage causing loss of control capability can compromise mission objectives and even result in aircraft loss. The Self Repairing Flight Control System (SRFCS) flight development program directly addresses this issue with a flight control system design that measures the damage and immediately refines the control system commands to preserve mission potential. The system diagnostics process detects in flight the type of faults that are difficult to isolate post flight, and thus cause excessive ground maintenance time and cost. The control systems of fighter aircraft have the control power and surface displacement to maneuver the aircraft in a very large flight envelope with a wide variation in airspeed and g maneuvering conditions, with surplus force capacity available from each control surface. Digital flight control processors are designed to include built-in status of the control system components, as well as sensor information on aircraft control maneuver commands and response. In the event of failure or loss of a control surface, the SRFCS utilizes this capability to reconfigure control commands to the remaining control surfaces, thus preserving maneuvering response. Correct post-flight repair is the key to low maintainability support costs and high aircraft mission readiness. The SRFCS utilizes the large data base available with digital flight control systems to diagnose faults. Built-in-test data and sensor data are used as inputs to an Onboard Expert System process to accurately identify failed components for post-flight maintenance action. This diagnostic technique has the advantage of functioning during flight, and so is especially useful in identifying intermittent faults that are present only during maneuver g loads or high hydraulic flow requirements. A flight system was developed to test the reconfiguration and onboard maintenance diagnostics concepts on a NASA F-15 fighter aircraft.

Vista/F-16 Multi-Axis Thrust Vectoring (MATV) control law design and evaluation

NASA Technical Reports Server (NTRS)

Zwerneman, W. D.; Eller, B. G.

1994-01-01

For the Multi-Axis Thrust Vectoring (MATV) program, a new control law was developed using multi-axis thrust vectoring to augment the aircraft's aerodynamic control power to provide maneuverability above the normal F-16 angle of attack limit. The control law architecture was developed using Lockheed Fort Worth's offline and piloted simulation capabilities. The final flight control laws were used in flight test to demonstrate tactical benefits gained by using thrust vectoring in air-to-air combat. Differences between the simulator aerodynamics data base and the actual aircraft aerodynamics led to significantly different lateral-directional flying qualities during the flight test program than those identified during piloted simulation. A 'dial-a-gain' flight test control law update was performed in the middle of the flight test program. This approach allowed for inflight optimization of the aircraft's flying qualities. While this approach is not preferred over updating the simulator aerodynamic data base and then updating the control laws, the final selected gain set did provide adequate lateral-directional flying qualities over the MATV flight envelope. The resulting handling qualities and the departure resistance of the aircraft allowed the 422nd_squadron pilots to focus entirely on evaluating the aircraft's tactical utility.

Flying Drosophila stabilize their vision-based velocity controller by sensing wind with their antennae

PubMed Central

Fuller, Sawyer Buckminster; Straw, Andrew D.; Peek, Martin Y.; Murray, Richard M.; Dickinson, Michael H.

2014-01-01

Flies and other insects use vision to regulate their groundspeed in flight, enabling them to fly in varying wind conditions. Compared with mechanosensory modalities, however, vision requires a long processing delay (~100 ms) that might introduce instability if operated at high gain. Flies also sense air motion with their antennae, but how this is used in flight control is unknown. We manipulated the antennal function of fruit flies by ablating their aristae, forcing them to rely on vision alone to regulate groundspeed. Arista-ablated flies in flight exhibited significantly greater groundspeed variability than intact flies. We then subjected them to a series of controlled impulsive wind gusts delivered by an air piston and experimentally manipulated antennae and visual feedback. The results show that an antenna-mediated response alters wing motion to cause flies to accelerate in the same direction as the gust. This response opposes flying into a headwind, but flies regularly fly upwind. To resolve this discrepancy, we obtained a dynamic model of the fly’s velocity regulator by fitting parameters of candidate models to our experimental data. The model suggests that the groundspeed variability of arista-ablated flies is the result of unstable feedback oscillations caused by the delay and high gain of visual feedback. The antenna response drives active damping with a shorter delay (~20 ms) to stabilize this regulator, in exchange for increasing the effect of rapid wind disturbances. This provides insight into flies’ multimodal sensory feedback architecture and constitutes a previously unknown role for the antennae. PMID:24639532

DAST Mated to B-52 on Ramp - Close-up

NASA Technical Reports Server (NTRS)

1979-01-01

Technicians mount a BQM-43 Firebee II drone on the wing pylon of NASA's B-52B launch aircraft. The drone was test flown as part of the Drones for Aerodynamic and Structural Testing (DAST) program. Research flights of drones with modified wings for the DAST program were conducted from 1977 to 1983. After the initial flights of Firebee II 72-1564, it was fitted with the Instrumented Standard Wing (also called the 'Blue Streak' wing). The first free flight attempt on March 7, 1979, was aborted before launch due to mechanical problems with the HH-53 recovery helicopter. The next attempt, on March 9, 1979, was successful. These are the image contact sheets for each image resolution of the NASA Dryden Drones for Aerodynamic and Structural Testing (DAST) Photo Gallery. From 1977 to 1983, the Dryden Flight Research Center, Edwards, California, (under two different names) conducted the DAST Program as a high-risk flight experiment using a ground-controlled, pilotless aircraft. Described by NASA engineers as a 'wind tunnel in the sky,' the DAST was a specially modified Teledyne-Ryan BQM-34E/F Firebee II supersonic target drone that was flown to validate theoretical predictions under actual flight conditions in a joint project with the Langley Research Center, Hampton, Virginia. The DAST Program merged advances in electronic remote control systems with advances in airplane design. Drones (remotely controlled, missile-like vehicles initially developed to serve as gunnery targets) had been deployed successfully during the Vietnamese conflict as reconnaissance aircraft. After the war, the energy crisis of the 1970s led NASA to seek new ways to cut fuel use and improve airplane efficiency. The DAST Program's drones provided an economical, fuel-conscious method for conducting in-flight experiments from a remote ground site. DAST explored the technology required to build wing structures with less than normal stiffness. This was done because stiffness requires structural weight but ensures freedom from flutter-an uncontrolled, divergent oscillation of the structure, driven by aerodynamic forces and resulting in structural failure. The program used refined theoretical tools to predict at what speed flutter would occur. It then designed a high-response control system to counteract the motion and permit a much lighter wing structure. The wing had, in effect, 'electronic stiffness.' Flight research with this concept was extremely hazardous because an error in either the flutter prediction or control system implementation would result in wing structural failure and the loss of the vehicle. Because of this, flight demonstration of a sub-scale vehicle made sense from the standpoint of both safety and cost. The program anticipated structural failure during the course of the flight research. The Firebee II was a supersonic drone selected as the DAST testbed because its wing could be easily replaced, it used only tail-mounted control surfaces, and it was available as surplus from the U. S. Air Force. It was capable of 5-g turns (that is, turns producing acceleration equal to 5 times that of gravity). Langley outfitted a drone with an aeroelastic, supercritical research wing suitable for a Mach 0.98 cruise transport with a predicted flutter speed of Mach 0.95 at an altitude of 25,000 feet. Dryden and Langley, in conjunction with Boeing, designed and fabricated a digital flutter suppression system (FSS). Dryden developed an RPRV (remotely piloted research vehicle) flight control system; integrated the wing, FSS, and vehicle systems; and conducted the flight program. In addition to a digital flight control system and aeroelastic wings, each DAST drone had research equipment mounted in its nose and a mid-air retrieval system in its tail. The drones were originally launched from the NASA B-52 bomber and later from a DC-130. The DAST vehicle's flight was monitored from the sky by an F-104 chase plane. When the DAST's mission ended, it deployed a parachute and then a specially equipped Air Force helicopter recovered the drone in mid-air. On the ground, a pilot controlled the DAST vehicle from a remote cockpit while researchers in another room monitored flight data transmitted via telemetry. They made decisions on the conduct of the flight while the DAST was in the air. In case of failure in any of the ground systems, the DAST vehicle could also be flown to a recovery site using a backup control system in the F-104. The DAST Program experienced numerous problems. Only eighteen flights were achieved, eight of them captive (in which the aircraft flew only while still attached to the launch aircraft). Four of the flights were aborted and two resulted in crashes--one on June 12, 1980, and the second on June 1, 1983. Meanwhile, flight experiments with higher profiles, better funded remotely piloted research vehicles took priority over DAST missions. After the 1983 crash, which was caused by a malfunction that disconnected the landing parachute from the drone, the program was disbanded. Because DAST drones were considered expendable, certain losses were anticipated. Managers and researchers involved in other high-risk flight projects gained insights from the DAST program that could be applied to their own flight research programs. The DAST aircraft had a wingspan of 14 feet, four inches and a nose-to-tail length of 28 feet, 4 inches. The fuselage had a radius of about 2.07 feet. The aircraft's maximum loaded weight was about 2,200 pounds. It derived its power from a Continental YJ69-T-406 engine.

X-31 Loaded in C-5 Cargo Bay

NASA Technical Reports Server (NTRS)

1995-01-01

The X-31 Enhanced Fighter Maneuverability Technology Demonstrator Aircraft, based at the NASA Dryden Flight Research Center, Edwards Air Force Base, California, is secured inside the fuselage of an Air Force Reserve C-5 transport. The C-5 was used to ferry the X-31 from Europe back to Edwards, after being flown in the Paris Air Show in June 1995. The X-31's right wing, removed so the aircraft could fit inside the C-5, is in the shipping container in the foreground. At the air show, the X-31 demonstrated the value of using thrust vectoring (directing engine exhaust flow) coupled with advanced flight control systems to provide controlled flight at very high angles of attack. The X-31 Enhanced Fighter Maneuverability (EFM) demonstrator flew at the Ames- Dryden Flight Research Facility, Edwards, California (redesignated the Dryden Flight Research Center in 1994) from February 1992 until 1995 and before that at the Air Force's Plant 42 in Palmdale, California. The goal of the project was to provide design information for the next generation of highly maneuverable fighter aircraft. This program demonstrated the value of using thrust vectoring (directing engine exhaust flow) coupled with an advanced flight control system to provide controlled flight to very high angles of attack. The result was a significant advantage over most conventional fighters in close-in combat situations. The X-31 flight program focused on agile flight within the post-stall regime, producing technical data to give aircraft designers a better understanding of aerodynamics, effectiveness of flight controls and thrust vectoring, and airflow phenomena at high angles of attack. Stall is a condition of an airplane or an airfoil in which lift decreases and drag increases due to the separation of airflow. Thrust vectoring compensates for the loss of control through normal aerodynamic surfaces that occurs during a stall. Post-stall refers to flying beyond the normal stall angle of attack, which in the X-31 was at a 30-degree angle of attack. During Dryden flight testing, the X-31 aircraft established several milestones. On November 6, 1992, the X-31 achieved controlled flight at a 70-degree angle of attack. On April 29, 1993, the second X-31 successfully executed a rapid minimum-radius, 180-degree turn using a post-stall maneuver, flying well beyond the aerodynamic limits of any conventional aircraft. This revolutionary maneuver has been called the 'Herbst Maneuver' after Wolfgang Herbst, a German proponent of using post-stall flight in air-to-air combat. It is also called a 'J Turn' when flown to an arbitrary heading change. The aircraft was flown in tactical maneuvers against an F/A-18 and other tactical aircraft as part of the test flight program. During November and December 1993, the X-31 reached a supersonic speed of Mach 1.28. In 1994, the X-31 program installed software to demonstrate quasi-tailless operation. The X-31 flight test program was conducted by an international test organization (ITO) managed by the Advanced Research Projects Office (ARPA), known as the Defense Advanced Research Projects Office (DARPA) before March 1993. The ITO included the U.S. Navy and U.S. Air Force, Rockwell Aerospace, the Federal Republic of Germany, Daimler-Benz (formerly Messerschmitt-Bolkow-Blohm and Deutsche Aerospace), and NASA. Gary Trippensee was the ITO director and NASA Project Manager. Pilots came from participating organizations. The X-31 was 43.33 feet long with a wingspan of 23.83 feet. It was powered by a single General Electric P404-GE-400 turbofan engine that produced 16,000 pounds of thrust in afterburner.

View of new centrifuge at Flight Acceleration Facility

NASA Technical Reports Server (NTRS)

1966-01-01

View of the new centrifuge at the Manned Spacecraft Center (MSC), located in the Flight Acceleration Facility, bldg 29. The 50-ft. arm can swing the three man gondola to create g-forces astronauts will experience during controlled flight and during reentry. The centrifuge was designed primarily for training Apollo astronauts.

Design, analysis and control of large transports so that control of engine thrust can be used as a back-up of the primary flight controls. Ph.D. Thesis

NASA Technical Reports Server (NTRS)

Roskam, Jan; Ackers, Deane E.; Gerren, Donna S.

1995-01-01

A propulsion controlled aircraft (PCA) system has been developed at NASA Dryden Flight Research Center at Edwards Air Force Base, California, to provide safe, emergency landing capability should the primary flight control system of the aircraft fail. As a result of the successful PCA work being done at NASA Dryden, this project investigated the possibility of incorporating the PCA system as a backup flight control system in the design of a large, ultra-high capacity megatransport in such a way that flight path control using only the engines is not only possible, but meets MIL-Spec Level 1 or Level 2 handling quality requirements. An 800 passenger megatransport aircraft was designed and programmed into the NASA Dryden simulator. Many different analysis methods were used to evaluate the flying qualities of the megatransport while using engine thrust for flight path control, including: (1) Bode and root locus plot analysis to evaluate the frequency and damping ratio response of the megatransport; (2) analysis of actual simulator strip chart recordings to evaluate the time history response of the megatransport; and (3) analysis of Cooper-Harper pilot ratings by two NaSA test pilots.

Controls design with crossfeeds for hovering rotorcraft using quantitative feedback theory

NASA Technical Reports Server (NTRS)

Tischler, Mark B.; Biezad, Daniel J.; Cheng, Rendy

1996-01-01

A multi-input, multi-output controls design with dynamic crossfeed pre-compensation is presented for rotorcraft in near-hovering flight using Quantitative Feedback Theory (QFT). The resulting closed-loop control system bandwidth allows the rotorcraft to be considered for use as an inflight simulator. The use of dynamic, robust crossfeeds prior to the QFT design reduces the magnitude of required feedback gain and results in performance that meets most handling qualities specifications relative to the decoupling of off-axis responses. Handling qualities are Level 1 for both low-gain tasks and high-gain tasks in the roll, pitch, and yaw axes except for the 10 deg/sec moderate-amplitude yaw command where the rotorcraft exhibits Level 2 handling qualities in the yaw axis caused by phase lag. The combined effect of the QFT feedback design following the implementation of low-order, dynamic crossfeed compensators successfully decouples ten of twelve off-axis channels. For the other two channels it was not possible to find a single, low-order crossfeed that was effective. This is an area to be investigated in future research.

Aerodynamics power consumption for mechanical flapping wings undergoing flapping and pitching motion

NASA Astrophysics Data System (ADS)

Razak, N. A.; Dimitriadis, G.; Razaami, A. F.

2017-07-01

Lately, due to the growing interest in Micro Aerial Vehicles (MAV), interest in flapping flight has been rekindled. The reason lies in the improved performance of flapping wing flight at low Reynolds number regime. Many studies involving flapping wing flight focused on the generation of unsteady aerodynamic forces such as lift and thrust. There is one aspect of flapping wing flight that received less attention. The aspect is aerodynamic power consumption. Since most mechanical flapping wing aircraft ever designed are battery powered, power consumption is fundamental in improving flight endurance. This paper reports the results of experiments carried out on mechanical wings under going active root flapping and pitching in the wind tunnel. The objective of the work is to investigate the effect of the pitch angle oscillations and wing profile on the power consumption of flapping wings via generation of unsteady aerodynamic forces. The experiments were repeated for different airspeeds, flapping and pitching kinematics, geometric angle of attack and wing sections with symmetric and cambered airfoils. A specially designed mechanical flapper modelled on large migrating birds was used. It will be shown that, under pitch leading conditions, less power is required to overcome the unsteady aerodnamics forces. The study finds less power requirement for downstroke compared to upstroke motion. Overall results demonstrate power consumption depends directly on the unsteady lift force.

Rotorcraft flying qualities improvement using advanced control

NASA Technical Reports Server (NTRS)

Walker, D.; Postlethwaite, I.; Howitt, J.; Foster, N.

1993-01-01

We report on recent experience gained when a multivariable helicopter flight control law was tested on the Large Motion Simulator (LMS) at DRA Bedford. This was part of a study into the application of multivariable control theory to the design of full-authority flight control systems for high-performance helicopters. In this paper, we present some of the results that were obtained during the piloted simulation trial and from subsequent off-line simulation and analysis. The performance provided by the control law led to level 1 handling quality ratings for almost all of the mission task elements assessed, both during the real-time and off-line analysis.

Characterization of Custom-Designed Charge-Coupled Devices for Applications to Gas and Aerosol Monitoring Sensorcraft Instrument

NASA Technical Reports Server (NTRS)

Refaat, Tamer F.; Abedin, M. Nurul; Farnsworth, Glenn R.; Garcia, Christopher S.; Zawodny, Joseph M.

2005-01-01

Custom-designed charge-coupled devices (CCD) for Gas and Aerosols Monitoring Sensorcraft instrument were developed. These custom-designed CCD devices are linear arrays with pixel format of 512x1 elements and pixel size of 10x200 sq m. These devices were characterized at NASA Langley Research Center to achieve a full well capacity as high as 6,000,000 e-. This met the aircraft flight mission requirements in terms of signal-to-noise performance and maximum dynamic range. Characterization and analysis of the electrical and optical properties of the CCDs were carried out at room temperature. This includes measurements of photon transfer curves, gain coefficient histograms, read noise, and spectral response. Test results obtained on these devices successfully demonstrated the objectives of the aircraft flight mission. In this paper, we describe the characterization results and also discuss their applications to future mission.

Perseus B Taxi Tests in Preparation for a New Series of Flight Tests

NASA Image and Video Library

1998-04-27

The Perseus B remotely piloted aircraft taxis on the runway at Edwards Air Force Base, California, before a series of development flights at NASA's Dryden flight Research Center. The Perseus B is the latest of three versions of the Perseus design developed by Aurora Flight Sciences under NASA's Environmental Research Aircraft and Sensor Technology (ERAST) program.

Perseus B Taxi Tests in Preparation for a New Series of Flight Tests

NASA Image and Video Library

1998-04-27

The Perseus B remotely piloted aircraft on the runway at Edwards Air Force Base, California at the conclusion of a development flight at NASA's Dryden flight Research Center. The Perseus B is the latest of three versions of the Perseus design developed by Aurora Flight Sciences under NASA's Environmental Research Aircraft and Sensor Technology (ERAST) program.

Application of Strength Diagnosis.

ERIC Educational Resources Information Center

Newton, Robert U.; Dugan, Eric

2002-01-01

Discusses the various strength qualities (maximum strength, high- and low-load speed strength, reactive strength, rate of force development, and skill performance), noting why a training program design based on strength diagnosis can lead to greater efficacy and better performance gains for the athlete. Examples of tests used to assess strength…

Incorporating Manual and Autonomous Code Generation

NASA Technical Reports Server (NTRS)

McComas, David

1998-01-01

Code can be generated manually or using code-generated software tools, but how do you interpret the two? This article looks at a design methodology that combines object-oriented design with autonomic code generation for attitude control flight software. Recent improvements in space flight computers are allowing software engineers to spend more time engineering the applications software. The application developed was the attitude control flight software for an astronomical satellite called the Microwave Anisotropy Probe (MAP). The MAP flight system is being designed, developed, and integrated at NASA's Goddard Space Flight Center. The MAP controls engineers are using Integrated Systems Inc.'s MATRIXx for their controls analysis. In addition to providing a graphical analysis for an environment, MATRIXx includes an autonomic code generation facility called AutoCode. This article examines the forces that shaped the final design and describes three highlights of the design process: (1) Defining the manual to autonomic code interface; (2) Applying object-oriented design to the manual flight code; (3) Implementing the object-oriented design in C.

Accuracy of a vertical jump contact mat for determining jump height and flight time.

PubMed

Whitmer, Tyler D; Fry, Andrew C; Forsythe, Charles M; Andre, Matthew J; Lane, Michael T; Hudy, Andrea; Honnold, Darric E

2015-04-01

Several devices are available to measure vertical jump (VJ) height based on flight time, VJ reach height, or ground reaction forces. The purpose of this study was to determine the accuracy of a VJ mat for measuring flight time and VJ height compared with a VJ tester or a force plate. Seventeen men and 18 women (X ± SD; age = 20.9 ± 0.7 years, height = 176.1 ± 0.9 cm, weight = 72.6 ± 13.5 kg) served as subjects. Subjects performed counter-movement vertical jumps while standing on both a force plate (1,000 Hz) and a VJ mat. A Vertec VJ tester was used to measure jump reach. Compared with the force plate, the VJ mat reported greater VJ height (VJ mat = 0.50 ± 0.12 m, force plate = 0.34 ± 0.10 m) and flight time (VJ mat = 0.629 ± 0.078 seconds, force plate = 0.524 ± 0.077 seconds). Comparison of VJ heights from the VJ mat and the Vertec revealed no significant differences (Vertec = 0.48 ± 0.11 m). Regression analyses indicated strong relationships between testing methods and suggested that high VJ performances may be underestimated with the VJ mat. This particular VJ mat compared favorably with the Vertec but not the force plate. It seems that the different flight times derived from the VJ mat may permit the VJ mat to be in closer agreement with VJ heights from the Vertec. Also, the VJ mat may not be an appropriate tool for assessing high VJ performances (i.e., ≥0.70 m; ≈28 inches). Practitioners and researchers using similar VJ mats may not obtain accurate flight times and may underestimate high performers.

Effects of mass on aircraft sidearm controller characteristics

NASA Technical Reports Server (NTRS)

Wagner, Charles A.

1994-01-01

When designing a flight simulator, providing a set of low mass variable-characteristic pilot controls can be very difficult. Thus, a strong incentive exists to identify the highest possible mass that will not degrade the validity of a simulation. The NASA Dryden Flight Research Center has conducted a brief flight program to determine the maximum acceptable mass (system inertia) of an aircraft sidearm controller as a function of force gradient. This information is useful for control system design in aircraft as well as development of suitable flight simulator controls. A modified Learjet with a variable-characteristic sidearm controller was used to obtain data. A boundary was defined between mass considered acceptable and mass considered unacceptable to the pilot. This boundary is defined as a function of force gradient over a range of natural frequencies. This investigation is limited to a study of mass-frequency characteristics only. Results of this investigation are presented in this paper.

Ares-I-X Stability and Control Flight Test: Analysis and Plans

NASA Technical Reports Server (NTRS)

Brandon, Jay M.; Derry, Stephen D.; Heim, Eugene H.; Hueschen, Richard M.; Bacon, Barton J.

2008-01-01

The flight test of the Ares I-X vehicle provides a unique opportunity to reduce risk of the design of the Ares I vehicle and test out design, math modeling, and analysis methods. One of the key features of the Ares I design is the significant static aerodynamic instability coupled with the relatively flexible vehicle - potentially resulting in a challenging controls problem to provide adequate flight path performance while also providing adequate structural mode damping and preventing adverse control coupling to the flexible structural modes. Another challenge is to obtain enough data from the single flight to be able to conduct analysis showing the effectiveness of the controls solutions and have data to inform design decisions for Ares I. This paper will outline the modeling approaches and control system design to conduct this flight test, and also the system identification techniques developed to extract key information such as control system performance (gain/phase margins, for example), structural dynamics responses, and aerodynamic model estimations.

Design Guide for Aerodynamics Testing of Earth and Planetary Entry Vehicles in a Ballistic Range

NASA Technical Reports Server (NTRS)

Bogdanoff, David W.

2017-01-01

The purpose of this manual is to aid in the design of an aerodynamics test of an earth or planetary entry capsule in a ballistic range. In this manual, much use is made of the results and experience gained in 50 years of ballistic range aerodynamics testing at the NASA Ames Research Center, and in particular, that gained in the last 27 years, while the author was working at NASA Ames. The topics treated herein include: Data to be obtained; flight data needed to design test; Reynolds number and dynamic similarity of flight trajectory and ballistic range test; capabilities of various ballistic ranges; Calculations of swerves due to average and oscillating lift and of drag-induced velocity decreases; Model and sabot design; materials, weights and stresses; Sabot separation; Launches at angle of attack and slapping with paper to produce pitch/yaw oscillations.

Experiment Configurations for the DAST

NASA Technical Reports Server (NTRS)

1978-01-01

This image shows three vehicle configurations considered for the Drones for Aerodynamic and Structural Testing (DAST) program, conducted at NASA's Dryden Flight Research Center between 1977 and 1983. The DAST project planned for three wing configurations. These were the Instrumented Standard Wing (ISW), the Aeroelastic Research Wing-1 (ARW-1), and the ARW-2. After the DAST-1 crash, project personnel fitted a second Firebee II with a rebuilt ARW-1 wing. Due to the project's ending, it never flew the ARW-2 wing. These are the image contact sheets for each image resolution of the NASA Dryden Drones for Aerodynamic and Structural Testing (DAST) Photo Gallery. From 1977 to 1983, the Dryden Flight Research Center, Edwards, California, (under two different names) conducted the DAST Program as a high-risk flight experiment using a ground-controlled, pilotless aircraft. Described by NASA engineers as a 'wind tunnel in the sky,' the DAST was a specially modified Teledyne-Ryan BQM-34E/F Firebee II supersonic target drone that was flown to validate theoretical predictions under actual flight conditions in a joint project with the Langley Research Center, Hampton, Virginia. The DAST Program merged advances in electronic remote control systems with advances in airplane design. Drones (remotely controlled, missile-like vehicles initially developed to serve as gunnery targets) had been deployed successfully during the Vietnamese conflict as reconnaissance aircraft. After the war, the energy crisis of the 1970s led NASA to seek new ways to cut fuel use and improve airplane efficiency. The DAST Program's drones provided an economical, fuel-conscious method for conducting in-flight experiments from a remote ground site. DAST explored the technology required to build wing structures with less than normal stiffness. This was done because stiffness requires structural weight but ensures freedom from flutter-an uncontrolled, divergent oscillation of the structure, driven by aerodynamic forces and resulting in structural failure. The program used refined theoretical tools to predict at what speed flutter would occur. It then designed a high-response control system to counteract the motion and permit a much lighter wing structure. The wing had, in effect, 'electronic stiffness.' Flight research with this concept was extremely hazardous because an error in either the flutter prediction or control system implementation would result in wing structural failure and the loss of the vehicle. Because of this, flight demonstration of a sub-scale vehicle made sense from the standpoint of both safety and cost. The program anticipated structural failure during the course of the flight research. The Firebee II was a supersonic drone selected as the DAST testbed because its wing could be easily replaced, it used only tail-mounted control surfaces, and it was available as surplus from the U. S. Air Force. It was capable of 5-g turns (that is, turns producing acceleration equal to 5 times that of gravity). Langley outfitted a drone with an aeroelastic, supercritical research wing suitable for a Mach 0.98 cruise transport with a predicted flutter speed of Mach 0.95 at an altitude of 25,000 feet. Dryden and Langley, in conjunction with Boeing, designed and fabricated a digital flutter suppression system (FSS). Dryden developed an RPRV (remotely piloted research vehicle) flight control system; integrated the wing, FSS, and vehicle systems; and conducted the flight program. In addition to a digital flight control system and aeroelastic wings, each DAST drone had research equipment mounted in its nose and a mid-air retrieval system in its tail. The drones were originally launched from the NASA B-52 bomber and later from a DC-130. The DAST vehicle's flight was monitored from the sky by an F-104 chase plane. When the DAST's mission ended, it deployed a parachute and then a specially equipped Air Force helicopter recovered the drone in mid-air. On the ground, a pilot controlled the DAST vehicle from a remote cockpit while researchers in another room monitored flight data transmitted via telemetry. They made decisions on the conduct of the flight while the DAST was in the air. In case of failure in any of the ground systems, the DAST vehicle could also be flown to a recovery site using a backup control system in the F-104. The DAST Program experienced numerous problems. Only eighteen flights were achieved, eight of them captive (in which the aircraft flew only while still attached to the launch aircraft). Four of the flights were aborted and two resulted in crashes--one on June 12, 1980, and the second on June 1, 1983. Meanwhile, flight experiments with higher profiles, better funded remotely piloted research vehicles took priority over DAST missions. After the 1983 crash, which was caused by a malfunction that disconnected the landing parachute from the drone, the program was disbanded. Because DAST drones were considered expendable, certain losses were anticipated. Managers and researchers involved in other high-risk flight projects gained insights from the DAST program that could be applied to their own flight research programs. The DAST aircraft had a wingspan of 14 feet, four inches and a nose-to-tail length of 28 feet, 4 inches. The fuselage had a radius of about 2.07 feet. The aircraft's maximum loaded weight was about 2,200 pounds. It derived its power from a Continental YJ69-T-406 engine.

Extended Duration Orbiter Medical Project

NASA Technical Reports Server (NTRS)

Sawin, Charles F. (Editor); Taylor, Gerald R. (Editor); Smith, Wanda L. (Editor); Brown, J. Travis (Technical Monitor)

1999-01-01

Biomedical issues have presented a challenge to flight physicians, scientists, and engineers ever since the advent of high-speed, high-altitude airplane flight in the 1940s. In 1958, preparations began for the first manned space flights of Project Mercury. The medical data and flight experience gained through Mercury's six flights and the Gemini, Apollo, and Skylab projects, as well as subsequent space flights, comprised the knowledge base that was used to develop and implement the Extended Duration Orbiter Medical Project (EDOMP). The EDOMP yielded substantial amounts of data in six areas of space biomedical research. In addition, a significant amount of hardware was developed and tested under the EDOMP. This hardware was designed to improve data gathering capabilities and maintain crew physical fitness, while minimizing the overall impact to the microgravity environment. The biomedical findings as well as the hardware development results realized from the EDOMP have been important to the continuing success of extended Space Shuttle flights and have formed the basis for medical studies of crew members living for three to five months aboard the Russian space station, Mir. EDOMP data and hardware are also being used in preparation for the construction and habitation of International Space Station. All data sets were grouped to be non-attributable to individuals, and submitted to NASA s Life Sciences Data Archive.

Air Force loadmasters oversee unloading of the full-scale Orion abort test crew module mockup from a C-17 cargo aircraft at Edwards Air Force Base March 28.

NASA Image and Video Library

2008-03-28

A full-scale flight-test mockup of the Constellation program's Orion crew vehicle arrived at NASA's Dryden Flight Research Center in late March 2008 to undergo preparations for the first short-range flight test of the spacecraft's astronaut escape system later that year. Engineers and technicians at NASA's Langley Research Center fabricated the structure, which precisely represents the size, outer shape and mass characteristics of the Orion space capsule. The Orion crew module mockup was ferried to NASA Dryden on an Air Force C-17. After painting in the Edwards Air Force Base paint hangar, the conical capsule was taken to Dryden for installation of flight computers, instrumentation and other electronics prior to being sent to the U.S. Army's White Sands Missile Range in New Mexico for integration with the escape system and the first abort flight test in late 2008. The tests were designed to ensure a safe, reliable method of escape for astronauts in case of an emergency.

Model Development and Model-Based Control Design for High Performance Nonlinear Smart Systems

DTIC Science & Technology

2007-11-20

potentially impact a broad range of flow control problems of interest to the Air Force and Boeing. Point of contact: James Mabe , Boeing Phantom Works...rotorcraft blades. In both cases, models and control designs will be validated using data from Boeing experiments and flight tests. Point of contact: James ... Mabe , Boeing Phantom Works, Seattle, WA, 206-655-0091. 3. PZT Unimorphs – Boeing: Nonlinear structural models developed through AFOSR support are being

VUV Testing of Science Cameras at MSFC: QE Measurement of the CLASP Flight Cameras

NASA Technical Reports Server (NTRS)

Champey, Patrick; Kobayashi, Ken; Winebarger, Amy; Cirtain, Jonathan; Hyde, David; Robertson, Bryan; Beabout, Brent; Beabout, Dyana; Stewart, Mike

2015-01-01

The NASA Marshall Space Flight Center (MSFC) has developed a science camera suitable for sub-orbital missions for observations in the UV, EUV and soft X-ray. Six cameras were built and tested for the Chromospheric Lyman-Alpha Spectro-Polarimeter (CLASP), a joint National Astronomical Observatory of Japan (NAOJ) and MSFC sounding rocket mission. The CLASP camera design includes a frame-transfer e2v CCD57-10 512x512 detector, dual channel analog readout electronics and an internally mounted cold block. At the flight operating temperature of -20 C, the CLASP cameras achieved the low-noise performance requirements (less than or equal to 25 e- read noise and greater than or equal to 10 e-/sec/pix dark current), in addition to maintaining a stable gain of approximately equal to 2.0 e-/DN. The e2v CCD57-10 detectors were coated with Lumogen-E to improve quantum efficiency (QE) at the Lyman- wavelength. A vacuum ultra-violet (VUV) monochromator and a NIST calibrated photodiode were employed to measure the QE of each camera. Four flight-like cameras were tested in a high-vacuum chamber, which was configured to operate several tests intended to verify the QE, gain, read noise, dark current and residual non-linearity of the CCD. We present and discuss the QE measurements performed on the CLASP cameras. We also discuss the high-vacuum system outfitted for testing of UV and EUV science cameras at MSFC.

VUV testing of science cameras at MSFC: QE measurement of the CLASP flight cameras

NASA Astrophysics Data System (ADS)

Champey, P.; Kobayashi, K.; Winebarger, A.; Cirtain, J.; Hyde, D.; Robertson, B.; Beabout, B.; Beabout, D.; Stewart, M.

2015-08-01

The NASA Marshall Space Flight Center (MSFC) has developed a science camera suitable for sub-orbital missions for observations in the UV, EUV and soft X-ray. Six cameras were built and tested for the Chromospheric Lyman-Alpha Spectro-Polarimeter (CLASP), a joint MSFC, National Astronomical Observatory of Japan (NAOJ), Instituto de Astrofisica de Canarias (IAC) and Institut D'Astrophysique Spatiale (IAS) sounding rocket mission. The CLASP camera design includes a frame-transfer e2v CCD57-10 512 × 512 detector, dual channel analog readout and an internally mounted cold block. At the flight CCD temperature of -20C, the CLASP cameras exceeded the low-noise performance requirements (<= 25 e- read noise and <= 10 e- /sec/pixel dark current), in addition to maintaining a stable gain of ≍ 2.0 e-/DN. The e2v CCD57-10 detectors were coated with Lumogen-E to improve quantum efficiency (QE) at the Lyman- wavelength. A vacuum ultra-violet (VUV) monochromator and a NIST calibrated photodiode were employed to measure the QE of each camera. Three flight cameras and one engineering camera were tested in a high-vacuum chamber, which was configured to operate several tests intended to verify the QE, gain, read noise and dark current of the CCD. We present and discuss the QE measurements performed on the CLASP cameras. We also discuss the high-vacuum system outfitted for testing of UV, EUV and X-ray science cameras at MSFC.

Eclipse takeoff and flight

NASA Technical Reports Server (NTRS)

1998-01-01

This 25-second clip shows the QF-106 'Delta Dart' tethered to the USAF C-141A during takeoff and in flight. NASA Dryden Flight Research Center, Edwards, California, supported a Kelly Space and Technology, Inc. (KST)/U.S. Air Force project known as Eclipse, which demonstrated a reusable tow launch vehicle concept. The purpose of the project was to demonstrate a reusable tow launch vehicle concept that had been conceived and patented by KST. Kelly Space obtained a contract with the USAF Research Laboratory for the tow launch demonstration project under the Small Business Innovation Research (SBIR) program. The USAF SBIR contract included the modifications to turn the QF-106 into the Experimental Demonstrator #1 (EXD-01), and the C141A aircraft to incorporate the tow provisions to link the two aircraft, as well as conducting flight tests. The demonstration consisted of ground and flight tests. These tests included a Combined Systems Test of both airplanes joined by a tow rope, a towed taxi test, and six towed flights. The primary goal of the project was demonstrating the tow phase of the Eclipse concept using a scaled-down tow aircraft (C-141A) and a representative aerodynamically-shaped aircraft (QF-106A) as a launch vehicle. This was successfully accomplished. On December 20, 1997, NASA research pilot Mark Stucky flew a QF-106 on the first towed flight behind an Air Force C-141 in the joint Eclipse project with KST to demonstrate the reusable tow launch vehicle concept developed by KST. Kelly hoped to use the data from the tow tests to validate a tow-to-launch procedure for reusable space launch vehicles. Stucky flew six successful tow tests between December 1997 and February 6, 1998. On February 6, 1998, the sixth and final towed flight brought the project to a successful completion. Preliminary flight results determined that the handling qualities of the QF-106 on tow were very stable; actual flight measured values of tow rope tension were well within predictions made by the simulation, aerodynamic characteristics and elastic properties of the tow rope were a significant component of the towing system; and the Dryden high-fidelity simulation provided a representative model of the performance of the QF-106 and C-141A airplanes in tow configuration. Total time on tow for the entire project was 5 hours, 34 minutes, and 29 seconds. All six flights were highly productive, and all project objectives were achieved. All three of the project objectives were successfully accomplished. The objectives were: demonstration of towed takeoff, climb-out, and separation of the EXD-01 from the towing aircraft; validation of simulation models of the towed aircraft systems; and development of ground and flight procedures for towing and launching a delta-winged airplane configuration safely behind a transport-type aircraft. NASA Dryden served as the responsible test organization and had flight safety responsibility for the Eclipse project. Dryden also supplied engineering, simulation, instrumentation, range support, research pilots, and chase aircraft for the test series. Dryden personnel also performed the modifications to convert the QF-106 into the piloted EXD-01 aircraft. During the early flight phase of the project, Tracor, Inc. provided maintenance and ground support for the two QF-106 airplanes. The Air Force Flight Test Center (AFFTC), Edwards, California, provided the C-141A transport aircraft for the project, its flight and engineering support, and the aircrew. Kelly Space and Technology provided the modification design and fabrication of the hardware that was installed on the EXD-01 aircraft. Kelly Space and Technology hopes to use the data gleaned from the tow tests to develop a series of low-cost reusable launch vehicles, in particular to gain experience towing delta-wing aircraft having high wing loading, and in general to demonstrate various operational procedures such as ground processing and abort scenarios. The first successful towed flight occurred on December 20, 1997. Prior to this first tow test flight, the C-141A and EXD-01 were used to conduct a series of tethered taxi tests that would validate the tow procedures. Before these tethered taxi tests, a successful joint flight test was conducted in late October 1996, by Dryden, AFFTC, and KST, in which one of the Dryden F-18 chase aircraft flew at various ranges and locations behind the C-141A to define the wake turbulence and wingtip vortex environment. This flight test was replicated in July 1997, with an unmodified QF-106 flight proficiency aircraft.

Traveler Phase 1A Joint Review

NASA Technical Reports Server (NTRS)

St. John, Clint; Scofield, Jan; Skoog, Mark; Flock, Alex; Williams, Ethan; Guirguis, Luke; Loudon, Kevin; Sutherland, Jeffrey; Lehmann, Richard; Garland, Michael;

Stability Assessment and Tuning of an Adaptively Augmented Classical Controller for Launch Vehicle Flight Control

NASA Technical Reports Server (NTRS)

VanZwieten, Tannen; Zhu, J. Jim; Adami, Tony; Berry, Kyle; Grammar, Alex; Orr, Jeb S.; Best, Eric A.

2014-01-01

Recently, a robust and practical adaptive control scheme for launch vehicles [ [1] has been introduced. It augments a classical controller with a real-time loop-gain adaptation, and it is therefore called Adaptive Augmentation Control (AAC). The loop-gain will be increased from the nominal design when the tracking error between the (filtered) output and the (filtered) command trajectory is large; whereas it will be decreased when excitation of flex or sloshing modes are detected. There is a need to determine the range and rate of the loop-gain adaptation in order to retain (exponential) stability, which is critical in vehicle operation, and to develop some theoretically based heuristic tuning methods for the adaptive law gain parameters. The classical launch vehicle flight controller design technics are based on gain-scheduling, whereby the launch vehicle dynamics model is linearized at selected operating points along the nominal tracking command trajectory, and Linear Time-Invariant (LTI) controller design techniques are employed to ensure asymptotic stability of the tracking error dynamics, typically by meeting some prescribed Gain Margin (GM) and Phase Margin (PM) specifications. The controller gains at the design points are then scheduled, tuned and sometimes interpolated to achieve good performance and stability robustness under external disturbances (e.g. winds) and structural perturbations (e.g. vehicle modeling errors). While the GM does give a bound for loop-gain variation without losing stability, it is for constant dispersions of the loop-gain because the GM is based on frequency-domain analysis, which is applicable only for LTI systems. The real-time adaptive loop-gain variation of the AAC effectively renders the closed-loop system a time-varying system, for which it is well-known that the LTI system stability criterion is neither necessary nor sufficient when applying to a Linear Time-Varying (LTV) system in a frozen-time fashion. Therefore, a generalized stability metric for time-varying loop=gain perturbations is needed for the AAC.

In-flight total forces, moments and static aeroelastic characteristics of an oblique-wing research airplane

NASA Technical Reports Server (NTRS)

Curry, R. E.; Sim, A. G.

1984-01-01

A low-speed flight investigation has provided total force and moment coefficients and aeroelastic effects for the AD-1 oblique-wing research airplane. The results were interpreted and compared with predictions that were based on wind tunnel data. An assessment has been made of the aeroelastic wing bending design criteria. Lateral-directional trim requirements caused by asymmetry were determined. At angles of attack near stall, flow visualization indicated viscous flow separation and spanwise vortex flow. These effects were also apparent in the force and moment data.

ER-2 High Altitude Solar Cell Calibration Flights

NASA Technical Reports Server (NTRS)

Myers, Matthew; Wolford, David; Snyder, David; Piszczor, Michael

2015-01-01

Evaluation of space photovoltaics using ground-based simulators requires primary standard cells which have been characterized in a space or near-space environment. Due to the high cost inherent in testing cells in space, most primary standards are tested on high altitude fixed wing aircraft or balloons. The ER-2 test platform is the latest system developed by the Glenn Research Center (GRC) for near-space photovoltaic characterization. This system offers several improvements over GRC's current Learjet platform including higher altitude, larger testing area, onboard spectrometers, and longer flight season. The ER-2 system was developed by GRC in cooperation with NASA's Armstrong Flight Research Center (AFRC) as well as partners at the Naval Research Laboratory and Air Force Research Laboratory. The system was designed and built between June and September of 2014, with the integration and first flights taking place at AFRC's Palmdale facility in October of 2014. Three flights were made testing cells from GRC as well as commercial industry partners. Cell performance data was successfully collected on all three flights as well as solar spectra. The data was processed using a Langley extrapolation method, and performance results showed a less than half a percent variation between flights, and less than a percent variation from GRC's current Learjet test platform.

Scramjet nozzle design and analysis as applied to a highly integrated hypersonic research airplane

NASA Technical Reports Server (NTRS)

Small, W. J.; Weidner, J. P.; Johnston, P. J.

1976-01-01

Engine-nozzle airframe integration at hypersonic speeds was conducted by using a high-speed research aircraft concept as a focus. Recently developed techniques for analysis of scramjet-nozzle exhaust flows provide a realistic analysis of complex forces resulting from the engine-nozzle airframe coupling. By properly integrating the engine-nozzle propulsive system with the airframe, efficient, controlled and stable flight results over a wide speed range.

X-2 in flight after drop from B-50 mothership

NASA Technical Reports Server (NTRS)

1956-01-01

The Bell Aircraft Company X-2 (46-674) drops away from its Boeing B-50 mothership in this photo. Lt. Col. Frank 'Pete' Everest piloted 674 on its first unpowered flight on 5 August 1954. He made the first rocket-powered flight on 18 November 1955. Everest made the first supersonic X-2 flight in 674 on 25 April 1956, achieving a speed of Mach 1.40. In July, he reached Mach 2.87, just short of the Mach 3 goal. The other X-2, 675, was written off prior to making any powered flights. An explosion during a captive flight resulted in the death of Bell test pilot Jean 'Skip' Ziegler. The X-2 was jettisoned over Lake Ontario, and the launch aircraft was damaged beyond repair. The first X-2, 674, continued flying, making a total of 17 launches. On 7 September 1956, Capt. Iven Kincheloe became the first man to exceed 100,000 feet when he reached an altitude of 126,200 feet in 674. The X-2, initially an Air Force program, was scheduled to be transferred to the civilian National Advisory Committee for Aeronautics (NACA) for scientific research. The Air Force delayed turning the aircraft over to the NACA in the hope of attaining Mach 3 in the airplane. The service requested and received a two-month extension to qualify another Air Force test pilot, Capt. Miburn 'Mel' Apt, in the X-2 and attempt to exceed Mach 3. After several ground briefings in the simulator, Apt (with no previous rocket plane experience) made his flight on 27 September 1956. Apt raced away from the B-50 under full power, quickly outdistancing the F-100 chase planes. At high altitude, he nosed over, accelerating rapidly. The X-2 reached Mach 3.2 (2,094 mph) at 65,000 feet. Apt became the first man to fly more than three times the speed of sound. Still above Mach 3, he began an abrupt turn back to Edwards. This maneuver proved fatal as the X-2 began a series of diverging rolls and tumbled out of control. Apt tried to regain control of the aircraft. Unable to do so, Apt separated the escape capsule. Too late, he attempted to bail out and was killed when the capsule impacted on the Edwards bombing range. The rest of the X-2 crashed five miles away. The X-2 was a swept-wing, rocket-powered aircraft designed to fly faster than Mach 3 (three times the speed of sound). It was built for the U.S. Air Force by the Bell Aircraft Company, Buffalo, New York. The X-2 was flown to investigate the problems of aerodynamic heating as well as stability and control effectiveness at high altitudes and high speeds (in excess of Mach 3). Bell aircraft built two X-2 aircraft. These were constructed of K-monel (a copper and nickel alloy) for the fuselage and stainless steel for the swept wings and control surfaces. The aircraft had ejectable nose capsules instead of ejection seats because the development of ejection seats had not reached maturity at the time the X-2 was conceived. The X-2 ejection canopy was successfully tested using a German V-2 rocket. The X-2 used a skid-type landing gear to make room for more fuel. The airplane was air launched from a modified Boeing B-50 Superfortress Bomber. X-2 Number 1 made its first unpowered glide flight on Aug. 5, 1954, and made a total of 17 (4 glide and 13 powered) flights before it was lost Sept. 27, 1956. The pilot on Flight 17, Capt. Milburn Apt, had flown the aircraft to a record speed of Mach 3.2 (2,094 mph), thus becoming the first person to exceed Mach 3. During that last flight, inertial coupling occurred and the pilot was killed. The aircraft suffered little damage in the crash, resulting in proposals (never implemented) from the Langley Memorial Aeronautical Laboratory, Hampton, Virginia, to rebuild it for use in a hypersonic (Mach 5+) test program. In 1953, X-2 Number 2 was lost in an in-flight explosion while at the Bell Aircraft Company during captive flight trials and was jettisoned into Lake Ontario. The Air Force had previously flown the aircraft on three glide flights at Edwards Air Force Base, California, in 1952. Although the NACA's High-Speed Flight Station, Edwards, California, (predecessor of NASA's Dryden Flight Research Center) never actually flew the X-2 aircraft, the NACA did support the program primarily through Langley Memorial Aeronautical Laboratory wind-tunnel tests and Wallops Island, Virginia, rocket-model tests. The NACA High-Speed Flight Station also provided stability and control recording instrumentation and simulator support for the Air Force flights. In the latter regard, the NACA worked with the Air Force in using a special computer to extrapolate and predict aircraft behavior from flight data.

Design and implementation of a compliant robot with force feedback and strategy planning software

NASA Technical Reports Server (NTRS)

Premack, T.; Strempek, F. M.; Solis, L. A.; Brodd, S. S.; Cutler, E. P.; Purves, L. R.

1984-01-01

Force-feedback robotics techniques are being developed for automated precision assembly and servicing of NASA space flight equipment. Design and implementation of a prototype robot which provides compliance and monitors forces is in progress. Computer software to specify assembly steps and makes force feedback adjustments during assembly are coded and tested for three generically different precision mating problems. A model program demonstrates that a suitably autonomous robot can plan its own strategy.

Physics-based simulations of aerial attacks by peregrine falcons reveal that stooping at high speed maximizes catch success against agile prey

PubMed Central

Hildenbrandt, Hanno

2018-01-01

The peregrine falcon Falco peregrinus is renowned for attacking its prey from high altitude in a fast controlled dive called a stoop. Many other raptors employ a similar mode of attack, but the functional benefits of stooping remain obscure. Here we investigate whether, when, and why stooping promotes catch success, using a three-dimensional, agent-based modeling approach to simulate attacks of falcons on aerial prey. We simulate avian flapping and gliding flight using an analytical quasi-steady model of the aerodynamic forces and moments, parametrized by empirical measurements of flight morphology. The model-birds’ flight control inputs are commanded by their guidance system, comprising a phenomenological model of its vision, guidance, and control. To intercept its prey, model-falcons use the same guidance law as missiles (pure proportional navigation); this assumption is corroborated by empirical data on peregrine falcons hunting lures. We parametrically vary the falcon’s starting position relative to its prey, together with the feedback gain of its guidance loop, under differing assumptions regarding its errors and delay in vision and control, and for three different patterns of prey motion. We find that, when the prey maneuvers erratically, high-altitude stoops increase catch success compared to low-altitude attacks, but only if the falcon’s guidance law is appropriately tuned, and only given a high degree of precision in vision and control. Remarkably, the optimal tuning of the guidance law in our simulations coincides closely with what has been observed empirically in peregrines. High-altitude stoops are shown to be beneficial because their high airspeed enables production of higher aerodynamic forces for maneuvering, and facilitates higher roll agility as the wings are tucked, each of which is essential to catching maneuvering prey at realistic response delays. PMID:29649207

Force reflection with compliance control

NASA Technical Reports Server (NTRS)

Kim, Won S. (Inventor)

1993-01-01

Two types of systems for force-reflecting control, which enables high force-reflection gain, are presented: position-error-based force reflection and low-pass-filtered force reflection. Both of the systems are combined with shared compliance control. In the position-error-based class, the position error between the commanded and the actual position of a compliantly controlled robot is used to provide force reflection. In the low-pass-filtered force reflection class, the low-pass-filtered output of the compliance control is used to provide force reflection. The increase in force reflection gain can be more than 10-fold as compared to a conventional high-bandwidth pure force reflection system, when high compliance values are used for the compliance control.

Stability of simulated flight path control at +3 Gz in a human centrifuge.

PubMed

Guardiera, Simon; Dalecki, Marc; Bock, Otmar

2010-04-01

Earlier studies have shown that naïve subjects and experienced jet pilots produce exaggerated manual forces when exposed to increased acceleration (+Gz). This study was designed to evaluate whether this exaggeration affects the stability of simulated flight path control. We evaluated naïve subjects' performance in a flight simulator which either remained stationary (+1 Gz), or rotated to induce an acceleration in accordance to the simulated flight path with a mean acceleration of about +3 Gz. In either case, subjects were requested to produce a series of altitude changes in pursuit of a visual target airplane. Resulting flight paths were analyzed to determine the largest oscillation after an altitude change (Oscillation) and the mean deviation between subject and target flight path (Tracking Error). Flight stability after an altitude change was degraded in +3 Gz compared to +1 Gz, as evidenced by larger Oscillations (+11%) and increased Tracking Errors (+80%). These deficits correlated significantly with subjects' +3 Gz deficits in a manual-force production task. We conclude that force exaggeration in +3 Gz may impair flight stability during simulated jet maneuvers in naïve subjects, most likely as a consequence of vestibular stimulation.

Electric Propulsion Space Experiment (ESEX): Spacecraft design issues for high-power electric propulsion

NASA Astrophysics Data System (ADS)

Kriebel, Mary M.; Sanks, Terry M.

1992-02-01

Electric propulsion provides high specific impulses, and low thrust when compared to chemical propulsion systems. Therefore, electric propulsion offers improvements over chemical systems such as increased station-keeping time, prolonged on-orbit maneuverability, low acceleration of large structures, and increased launch vehicle flexibility. The anticipated near-term operational electric propulsion system for an electric orbit transfer vehicle is an arcjet propulsion system. Towards this end, the USAF's Phillips Laboratory (PL) has awarded a prime contract to TRW Space & Technology Group to design, build, and space qualify a 30-kWe class arcjet as well as develop and demonstrate, on the ground, a flight-qualified arcjet propulsion flight unit. The name of this effort is the 30 kWe Class Arcjet Advanced Technology Transition Demonstration (Arcjet ATTD) program. Once the flight unit has completed its ground qualification test, it will be given to the Space Test and Transportation Program Office of the Air Force's Space Systems Division (ST/T) for launch vehicle integration and space test. The flight unit's space test is known as the Electric Propulsion Space Experiment (ESEX). ESEX's mission scenario is 10 firings of 15 minutes each. The objectives of the ESEX flight are to measure arcjet plume deposition, electromagnetic interference, thermal radiation, and acceleration in space. Plume deposition, electromagnetic interference, and thermal radiation are operational issues that are primarily being answered for operational use. This paper describes the Arcjet ATTD flight unit design and identifies specifically how the diagnostic data will be collected as part of the ESEX program.

DAST Mated to B-52 in Flight - Close-up from Below

NASA Technical Reports Server (NTRS)

1977-01-01

This photo shows a BQM-34 Firebee II drone being carried aloft under the wing of NASA's B-52 mothership during a 1977 research flight. The Firebee/DAST research program ran from 1977 to 1983 at the NASA Dryden Flight Research Center, Edwards, California. This is the original Firebee II wing. Firebee 72-1564 made three captive flights--on November 25, 1975; May 17, 1976; and June 22, 1977--in preparation for the DAST project with modified wings. These were for checkout of the Firebee's systems and the prelaunch procedures. The first two used a DC-130A aircraft as the launch vehicle, while the third used the B-52. A single free flight using this drone occurred on July 28, 1977. The remote (ground) pilot was NASA research pilot Bill Dana. The launch and flight were successful, and the drone was caught in midair by an HH-53 helicopter. These are the image contact sheets for each image resolution of the NASA Dryden Drones for Aerodynamic and Structural Testing (DAST) Photo Gallery. From 1977 to 1983, the Dryden Flight Research Center, Edwards, California, (under two different names) conducted the DAST Program as a high-risk flight experiment using a ground-controlled, pilotless aircraft. Described by NASA engineers as a 'wind tunnel in the sky,' the DAST was a specially modified Teledyne-Ryan BQM-34E/F Firebee II supersonic target drone that was flown to validate theoretical predictions under actual flight conditions in a joint project with the Langley Research Center, Hampton, Virginia. The DAST Program merged advances in electronic remote control systems with advances in airplane design. Drones (remotely controlled, missile-like vehicles initially developed to serve as gunnery targets) had been deployed successfully during the Vietnamese conflict as reconnaissance aircraft. After the war, the energy crisis of the 1970s led NASA to seek new ways to cut fuel use and improve airplane efficiency. The DAST Program's drones provided an economical, fuel-conscious method for conducting in-flight experiments from a remote ground site. DAST explored the technology required to build wing structures with less than normal stiffness. This was done because stiffness requires structural weight but ensures freedom from flutter-an uncontrolled, divergent oscillation of the structure, driven by aerodynamic forces and resulting in structural failure. The program used refined theoretical tools to predict at what speed flutter would occur. It then designed a high-response control system to counteract the motion and permit a much lighter wing structure. The wing had, in effect, 'electronic stiffness.' Flight research with this concept was extremely hazardous because an error in either the flutter prediction or control system implementation would result in wing structural failure and the loss of the vehicle. Because of this, flight demonstration of a sub-scale vehicle made sense from the standpoint of both safety and cost. The program anticipated structural failure during the course of the flight research. The Firebee II was a supersonic drone selected as the DAST testbed because its wing could be easily replaced, it used only tail-mounted control surfaces, and it was available as surplus from the U. S. Air Force. It was capable of 5-g turns (that is, turns producing acceleration equal to 5 times that of gravity). Langley outfitted a drone with an aeroelastic, supercritical research wing suitable for a Mach 0.98 cruise transport with a predicted flutter speed of Mach 0.95 at an altitude of 25,000 feet. Dryden and Langley, in conjunction with Boeing, designed and fabricated a digital flutter suppression system (FSS). Dryden developed an RPRV (remotely piloted research vehicle) flight control system; integrated the wing, FSS, and vehicle systems; and conducted the flight program. In addition to a digital flight control system and aeroelastic wings, each DAST drone had research equipment mounted in its nose and a mid-air retrieval system in its tail. The drones were originally launched from the NASA B-52 bomber and later from a DC-130. The DAST vehicle's flight was monitored from the sky by an F-104 chase plane. When the DAST's mission ended, it deployed a parachute and then a specially equipped Air Force helicopter recovered the drone in mid-air. On the ground, a pilot controlled the DAST vehicle from a remote cockpit while researchers in another room monitored flight data transmitted via telemetry. They made decisions on the conduct of the flight while the DAST was in the air. In case of failure in any of the ground systems, the DAST vehicle could also be flown to a recovery site using a backup control system in the F-104. The DAST Program experienced numerous problems. Only eighteen flights were achieved, eight of them captive (in which the aircraft flew only while still attached to the launch aircraft). Four of the flights were aborted and two resulted in crashes--one on June 12, 1980, and the second on June 1, 1983. Meanwhile, flight experiments with higher profiles, better funded remotely piloted research vehicles took priority over DAST missions. After the 1983 crash, which was caused by a malfunction that disconnected the landing parachute from the drone, the program was disbanded. Because DAST drones were considered expendable, certain losses were anticipated. Managers and researchers involved in other high-risk flight projects gained insights from the DAST program that could be applied to their own flight research programs. The DAST aircraft had a wingspan of 14 feet, four inches and a nose-to-tail length of 28 feet, 4 inches. The fuselage had a radius of about 2.07 feet. The aircraft's maximum loaded weight was about 2,200 pounds. It derived its power from a Continental YJ69-T-406 engine.

Development of a Model Following Control Law for Inflight Simulation and Flight Controls Research

NASA Technical Reports Server (NTRS)

Takahashi, Mark; Fletcher, Jay; Aiken, Edwin W. (Technical Monitor)

1994-01-01

The U.S. Army and NASA are currently developing the Rotorcraft Aircrew Systems Concepts Airborne Laboratory (RASCAL) at the Ames Research Center. RASCAL, shown in Figure 1, is a UH-60, which is being modified in a phased development program to have a research fly-by-wire flight control system, and an advanced navigation research platform. An important part of the flight controls and handling qualities research on RASCAL will be an FCS design for the aircraft to achieve high bandwidth control responses and disturbance rejection characteristics. Initially, body states will be used as feedbacks, but research into the use of rotor states will also be considered in later stages to maximize agility and maneuverability. In addition to supporting flight controls research, this FCS design will serve as the inflight simulation control law to support basic handling qualities, guidance, and displays research. Research in high bandwidth controls laws is motivated by the desire to improve the handling qualities in aggressive maneuvering and in severely degraded weather conditions. Naturally, these advantages will also improve the quality of the model following, thereby improving the inflight simulation capabilities of the research vehicle. High bandwidth in the control laws provides tighter tracking allowing for higher response bandwidths which can meet handling qualities requirements for aggressive maneuvering. System sensitivity is also reduced preventing variations in the response from the vehicle due to changing flight conditions. In addition, improved gust rejection will result from this reduced sensitivity. The gust rejection coupled with a highly stable system will make more precise maneuvering and pointing possible in severely degraded weather conditions. The difficulty in achieving higher bandwidths from the control laws in the feedback and in the responses arises from the complexity of the models that are needed to produce a satisfactory design. In this case, high quality models that include rotor dynamics in a physically meaningful context must be available. A non-physical accounting of the rotor, such as lumping the effect as a time delay, is not likely to produce the desired results. High order simulation models based on first principals are satisfactory for the initial design phase in order to work out the control law design concept and get an initial set of gains. These models, however, have known deficiencies, which must be resolved in the final control law design. The error in the pitch-roll cross coupling is one notable deficiency that even sophisticated rotorcraft models including complex wake aerodynamics have yet to capture successfully. This error must be accounted for to achieve the desired decoupling. The approach to design the proposed inflight simulation control law is based on using a combination of simulation and identified models. The linear and nonlinear higher order models were used to develop an explicit model following control structure. This structure was developed to accommodate the design of control laws compliant to many of the quantitative requirements in ADS-33C. Furthermore, it also allows for control law research using rotor-state feedback and other design methodologies such as Quantitative Feedback and H-Infinity. Final gain selection will be based on higher order identified models which include rotor degrees of freedom.

A program for calculating load coefficient matrices utilizing the force summation method, L218 (LOADS). Volume 2: Supplemental system design and maintenance document

NASA Technical Reports Server (NTRS)

Anderson, L. R.; Miller, R. D.

1979-01-01

The LOADS computer program L218 which calculates dynamic load coefficient matrices utilizing the force summation method is described. The load equations are derived for a flight vehicle in straight and level flight and excited by gusts and/or control motions. In addition, sensor equations are calculated for use with an active control system. The load coefficient matrices are calculated for the following types of loads: (1) translational and rotational accelerations, velocities, and displacements; (2) panel aerodynamic forces; (3) net panel forces; and (4) shears, bending moments, and torsions.

Perseus B Heads for Landing on Edwards AFB Runway

NASA Image and Video Library

1997-04-30

The Perseus B remotely piloted aircraft nears touchdown at Edwards Air Force Base, Calif. at the conclusion of a development flight at NASA's Dryden Flight Research Center. The Perseus B is the latest of three versions of the Perseus design developed by Aurora Flight Sciences under NASA's Environmental Research Aircraft and Sensor Technology (ERAST) program.

"Miniature Aces" NASA's Dale Reed Flight Research Laboratory

NASA Image and Video Library

2015-02-23

This video uncovers the workings, tools, and rationale of the scaled aircraft lab at NASA’s Armstrong Flight Research Center on Edwards Air Force Base. Watch commercial-off-the-shelf aircraft, one-of-a-kind designs, powered aircraft, and gliders take-off, fly, and land. The chief pilot and designer of the lab explains how and why they do what they do.

X-31 impact site

NASA Technical Reports Server (NTRS)

1995-01-01

This photos shows the remains of the number one X-31 following a mishap on 19 January 1995. Two X-31A demonstartor aircraft were built to explore Enhanced Fighter Manueverability (EFM). The aircraft demonstrated techniques to allow close-in aerial combat beyond normal flight envelope parameters. The X-31 EFM demonstrator program was the first international effort to buid an X-plane. The aircraft was designed and funded jointly by the U.S. Defense Advanced Research Projects Agency (DARPA) and the German Ministry of Defense. Design and construction of components and systems was split between Rockwell International in the United States and Messerschmitt-Bolkow-Blohm (MBB) in the Federal Republic of Germany. Other participants included the U.S. Air Force, U.S. Navy, and NASA. Fianl assembly took place at Rockwell's North American Aircraft Division. The flight test progarm included an international team of pilots from Rockwell, MBB, and the U.S. and German armed forces. The demonstrator aircraft were designed for subsonic flight only. Overall program costs were reduced by using as amny off-the-shelf components as possible. The weight of the airframe was reduced through extensive use of graphite/epoxy thermosplastics in construction of aerodynamic surfaces and the forward fuselage. the design included fully movable canards, thrust-vectoring paddles, and a cranked composite delta wing. The first X-31A (Bu. No. 164584) was rolled out from Rockwell's Palmdale, Californai facility on 1 March 1990. Rockwell chief test pilot Norman 'Ken' Dyson made the first flight on 11 October 1990. Aircraft 584 completed a total of 292 flights during the EFM test program. Aircraft 584 was lost on 19 January 1995 while returning to Edwards Air Force Base on its third flight of the day. German test pilot Karl-Heinz Langwas nearing the end of a one-hour mission to measure aircraft parameters. The aircraft was flying at 21,000 feet and 141 knots in straight and level flight with the thrust-vectoring and high angle-of-attack (AOA) flight control mode turned off. Lang noticed erroneous airspeed readings (207 knots indicated airspeed at 20 degrees AOA). About a minute later, while executing the pre-landing checklist, the X-31 began a rapidly increasing series of pitch oscillations, ending up at about 20 degrees past vertical. This was followed by a sharp roll. Lang ejected at 18,050 feet, and parachuted to safety with some injuries. The X-31 crashed in a sparsely-populated area a mile and a half north of the base boundary. Investigators later determined that ice in the air data probe caused the erroneous readings that led to the accident.

HIFiRE-5 Flight Vehicle Design

NASA Technical Reports Server (NTRS)

Kimmel, Roger L.; Adamczak, David; Berger, Karen; Choudhari, Meelan

2010-01-01

The Hypersonic International Flight Research Experimentation (HIFiRE) program is a hypersonic flight test program executed by the Air Force Research Laboratories (AFRL) and Australian Defence Science and Technology Organization (DSTO). HIFiRE flight 5 is devoted to measuring transition on a three-dimensional body. This paper summarizes payload configuration, trajectory, vehicle stability limits and roughness tolerances. Results show that the proposed configuration is suitable for testing transition on a three-dimensional body. Transition is predicted to occur within the test window, and a design has been developed that will allow the vehicle to be manufactured within prescribed roughness tolerances

Investigation of aerodynamic design issues with regions of separated flow

NASA Technical Reports Server (NTRS)

Gally, Tom

1993-01-01

Existing aerodynamic design methods have generally concentrated on the optimization of airfoil or wing shapes to produce a minimum drag while satisfying some basic constraints such as lift, pitching moment, or thickness. Since the minimization of drag almost always precludes the existence of separated flow, the evaluation and validation of these design methods for their robustness and accuracy when separated flow is present has not been aggressively pursued. However, two new applications for these design tools may be expected to include separated flow and the issues of aerodynamic design with this feature must be addressed. The first application of the aerodynamic design tools is the design of airfoils or wings to provide an optimal performance over a wide range of flight conditions (multipoint design). While the definition of 'optimal performance' in the multipoint setting is currently being hashed out, it is recognized that given a wide range of flight conditions, it will not be possible to ensure a minimum drag constraint at all conditions, and in fact some amount of separated flow (presumably small) may have to be allowed at the more demanding flight conditions. Thus a multipoint design method must be tolerant of the existence of separated flow and may include some controls upon its extent. The second application is in the design of wings with extended high speed buffet boundaries of their flight envelopes. Buffet occurs on a wing when regions of flow separation have grown to the extent that their time varying pressures induce possible destructive effects upon the wing structure or adversely effect either the aircraft controllability or passenger comfort. A conservative approach to the expansion of the buffet flight boundary is to simply expand the flight envelope of nonseparated flow under the assumption that buffet will also thus be alleviated. However, having the ability to design a wing with separated flow and thus to control the location, extent and severity of the separated flow regions may allow aircraft manufacturers to gain an advantage in the early design stages of an aircraft, when configuration changes are relatively inexpensive to make. The goal of the summer research at NASA Langley Research Center (LaRC) was twofold: first, to investigate a particular airfoil design problem observed under conditions of strong shock induced flow separation on the upper surface of an airfoil at transonic conditions; and second, to suggest and investigate design methodologies for the prediction (or detection) and control of flow separation. The context of both investigations was to use an existing two dimensional Navier-Stokes flow solver and the constrained direct/iterative surface curvature (CDISC) design algorithm developed at LaRC. As a lead in to the primary task, it was necessary to gain a familiarity with both the design method and the computational analysis and to perform the FORTRAN coding needed to couple them together.

Adaptive wing and flow control technology

NASA Astrophysics Data System (ADS)

Stanewsky, E.

2001-10-01

The development of the boundary layer and the interaction of the boundary layer with the outer “inviscid” flow field, exacerbated at high speed by the occurrence of shock waves, essentially determine the performance boundaries of high-speed flight. Furthermore, flight and freestream conditions may change considerably during an aircraft mission while the aircraft itself is only designed for multiple but fixed design points thus impairing overall performance. Consequently, flow and boundary layer control and adaptive wing technology may have revolutionary new benefits for take-off, landing and cruise operating conditions for many aircraft by enabling real-time effective geometry optimization relative to the flight conditions. In this paper we will consider various conventional and novel means of boundary layer and flow control applied to moderate-to-large aspect ratio wings, delta wings and bodies with the specific objectives of drag reduction, lift enhancement, separation suppression and the improvement of air-vehicle control effectiveness. In addition, adaptive wing concepts of varying complexity and corresponding aerodynamic performance gains will be discussed, also giving some examples of possible structural realizations. Furthermore, penalties associated with the implementation of control and adaptation mechanisms into actual aircraft will be addressed. Note that the present contribution is rather application oriented.

Aerodynamic surface distension system for high angle of attack forebody vortex control

NASA Technical Reports Server (NTRS)

Zell, Peter T. (Inventor)

1994-01-01

A deployable system is introduced for assisting flight control under certain flight conditions, such as at high angles of attack, whereby two inflatable membranes are located on the forebody portion of an aircraft on opposite sides thereof. The members form control surfaces for effecting lateral control forces if one is inflated and longitudinal control forces if both are inflated.

Design of Quiet Rotorcraft Approach Trajectories: Verification Phase

NASA Technical Reports Server (NTRS)

Padula, Sharon L.

2010-01-01

Flight testing that is planned for October 2010 will provide an opportunity to evaluate rotorcraft trajectory optimization techniques. The flight test will involve a fully instrumented MD-902 helicopter, which will be flown over an array of microphones. In this work, the helicopter approach trajectory is optimized via a multiobjective genetic algorithm to improve community noise, passenger comfort, and pilot acceptance. Previously developed optimization strategies are modified to accommodate new helicopter data and to increase pilot acceptance. This paper describes the MD-902 trajectory optimization plus general optimization strategies and modifications that are needed to reduce the uncertainty in noise predictions. The constraints that are imposed by the flight test conditions and characteristics of the MD-902 helicopter limit the testing possibilities. However, the insights that will be gained through this research will prove highly valuable.

DAST Being Calibrated for Flight in Hangar

NASA Technical Reports Server (NTRS)

1982-01-01

DAST-2, a modified BQM-34 Firebee II drone, undergoes calibration in a hangar at the NASA Dryden Flight Research Center. After the crash of the first DAST vehicle, project personnel fitted a second Firebee II (serial # 72-1558) with the rebuilt ARW-1 (ARW-1R) wing. The DAST-2 made a captive flight aboard the B-52 on October 29, 1982, followed by a free flight on November 3, 1982. During January and February of 1983, three launch attempts from the B-52 had to be aborted due to various problems. Following this, the project changed the launch aircraft to a DC-130A. Two captive flights occurred in May 1983. The first launch attempt from the DC-130 took place on June 1, 1983. The mothership released the DAST-2, but the recovery system immediately fired without being commanded. The parachute then disconnected from the vehicle, and the DAST-2 crashed into a farm field near Harper Dry Lake. Wags called this the 'Alfalfa Field Impact Test.' These are the image contact sheets for each image resolution of the NASA Dryden Drones for Aerodynamic and Structural Testing (DAST) Photo Gallery. From 1977 to 1983, the Dryden Flight Research Center, Edwards, California, (under two different names) conducted the DAST Program as a high-risk flight experiment using a ground-controlled, pilotless aircraft. Described by NASA engineers as a 'wind tunnel in the sky,' the DAST was a specially modified Teledyne-Ryan BQM-34E/F Firebee II supersonic target drone that was flown to validate theoretical predictions under actual flight conditions in a joint project with the Langley Research Center, Hampton, Virginia. The DAST Program merged advances in electronic remote control systems with advances in airplane design. Drones (remotely controlled, missile-like vehicles initially developed to serve as gunnery targets) had been deployed successfully during the Vietnamese conflict as reconnaissance aircraft. After the war, the energy crisis of the 1970s led NASA to seek new ways to cut fuel use and improve airplane efficiency. The DAST Program's drones provided an economical, fuel-conscious method for conducting in-flight experiments from a remote ground site. DAST explored the technology required to build wing structures with less than normal stiffness. This was done because stiffness requires structural weight but ensures freedom from flutter-an uncontrolled, divergent oscillation of the structure, driven by aerodynamic forces and resulting in structural failure. The program used refined theoretical tools to predict at what speed flutter would occur. It then designed a high-response control system to counteract the motion and permit a much lighter wing structure. The wing had, in effect, 'electronic stiffness.' Flight research with this concept was extremely hazardous because an error in either the flutter prediction or control system implementation would result in wing structural failure and the loss of the vehicle. Because of this, flight demonstration of a sub-scale vehicle made sense from the standpoint of both safety and cost. The program anticipated structural failure during the course of the flight research. The Firebee II was a supersonic drone selected as the DAST testbed because its wing could be easily replaced, it used only tail-mounted control surfaces, and it was available as surplus from the U. S. Air Force. It was capable of 5-g turns (that is, turns producing acceleration equal to 5 times that of gravity). Langley outfitted a drone with an aeroelastic, supercritical research wing suitable for a Mach 0.98 cruise transport with a predicted flutter speed of Mach 0.95 at an altitude of 25,000 feet. Dryden and Langley, in conjunction with Boeing, designed and fabricated a digital flutter suppression system (FSS). Dryden developed an RPRV (remotely piloted research vehicle) flight control system; integrated the wing, FSS, and vehicle systems; and conducted the flight program. In addition to a digital flight control system and aeroelastic wings, each DAST drone had research equipment mounted in its nose and a mid-air retrieval system in its tail. The drones were originally launched from the NASA B-52 bomber and later from a DC-130. The DAST vehicle's flight was monitored from the sky by an F-104 chase plane. When the DAST's mission ended, it deployed a parachute and then a specially equipped Air Force helicopter recovered the drone in mid-air. On the ground, a pilot controlled the DAST vehicle from a remote cockpit while researchers in another room monitored flight data transmitted via telemetry. They made decisions on the conduct of the flight while the DAST was in the air. In case of failure in any of the ground systems, the DAST vehicle could also be flown to a recovery site using a backup control system in the F-104. The DAST Program experienced numerous problems. Only eighteen flights were achieved, eight of them captive (in which the aircraft flew only while still attached to the launch aircraft). Four of the flights were aborted and two resulted in crashes--one on June 12, 1980, and the second on June 1, 1983. Meanwhile, flight experiments with higher profiles, better funded remotely piloted research vehicles took priority over DAST missions. After the 1983 crash, which was caused by a malfunction that disconnected the landing parachute from the drone, the program was disbanded. Because DAST drones were considered expendable, certain losses were anticipated. Managers and researchers involved in other high-risk flight projects gained insights from the DAST program that could be applied to their own flight research programs. The DAST aircraft had a wingspan of 14 feet, four inches and a nose-to-tail length of 28 feet, 4 inches. The fuselage had a radius of about 2.07 feet. The aircraft's maximum loaded weight was about 2,200 pounds. It derived its power from a Continental YJ69-T-406 engine.

DAST in Flight Showing Diverging Wingtip Oscillations

NASA Technical Reports Server (NTRS)

1980-01-01

Two BQM-34 Firebee II drones were modified with supercritical airfoils, called the Aeroelastic Research Wing (ARW), for the Drones for Aerodynamic and Structural Testing (DAST) program, which ran from 1977 to 1983. In this view of DAST-1 (Serial # 72-1557), taken on June 12, 1980, severe wingtip flutter is visible. Moments later, the right wing failed catastrophically and the vehicle crashed near Cuddeback Dry Lake. Before the drone was lost, it had made two captive and two free flights. Its first free flight, on October 2, 1979, was cut short by an uplink receiver failure. The drone was caught in midair by an HH-3 helicopter. The second free flight, on March 12, 1980, was successful, ending in a midair recovery. The third free flight, made on June 12, was to expand the flutter envelope. All of these missions launched from the NASA B-52. From 1977 to 1983, the Dryden Flight Research Center, Edwards, California, (under two different names) conducted the DAST Program as a high-risk flight experiment using a ground-controlled, pilotless aircraft. Described by NASA engineers as a 'wind tunnel in the sky,' the DAST was a specially modified Teledyne-Ryan BQM-34E/F Firebee II supersonic target drone that was flown to validate theoretical predictions under actual flight conditions in a joint project with the Langley Research Center, Hampton, Virginia. The DAST Program merged advances in electronic remote control systems with advances in airplane design. Drones (remotely controlled, missile-like vehicles initially developed to serve as gunnery targets) had been deployed successfully during the Vietnamese conflict as reconnaissance aircraft. After the war, the energy crisis of the 1970s led NASA to seek new ways to cut fuel use and improve airplane efficiency. The DAST Program's drones provided an economical, fuel-conscious method for conducting in-flight experiments from a remote ground site. DAST explored the technology required to build wing structures with less than normal stiffness. This was done because stiffness requires structural weight but ensures freedom from flutter-an uncontrolled, divergent oscillation of the structure, driven by aerodynamic forces and resulting in structural failure. The program used refined theoretical tools to predict at what speed flutter would occur. It then designed a high-response control system to counteract the motion and permit a much lighter wing structure. The wing had, in effect, 'electronic stiffness.' Flight research with this concept was extremely hazardous because an error in either the flutter prediction or control system implementation would result in wing structural failure and the loss of the vehicle. Because of this, flight demonstration of a sub-scale vehicle made sense from the standpoint of both safety and cost. The program anticipated structural failure during the course of the flight research. The Firebee II was a supersonic drone selected as the DAST testbed because its wing could be easily replaced, it used only tail-mounted control surfaces, and it was available as surplus from the U. S. Air Force. It was capable of 5-g turns (that is, turns producing acceleration equal to 5 times that of gravity). Langley outfitted a drone with an aeroelastic, supercritical research wing suitable for a Mach 0.98 cruise transport with a predicted flutter speed of Mach 0.95 at an altitude of 25,000 feet. Dryden and Langley, in conjunction with Boeing, designed and fabricated a digital flutter suppression system (FSS). Dryden developed an RPRV (remotely piloted research vehicle) flight control system; integrated the wing, FSS, and vehicle systems; and conducted the flight program. In addition to a digital flight control system and aeroelastic wings, each DAST drone had research equipment mounted in its nose and a mid-air retrieval system in its tail. The drones were originally launched from the NASA B-52 bomber and later from a DC-130. The DAST vehicle's flight was monitored from the sky by an F-104 chase plane. When the DAST's mission ended, it deployed a parachute and then a specially equipped Air Force helicopter recovered the drone in mid-air. On the ground, a pilot controlled the DAST vehicle from a remote cockpit while researchers in another room monitored flight data transmitted via telemetry. They made decisions on the conduct of the flight while the DAST was in the air. In case of failure in any of the ground systems, the DAST vehicle could also be flown to a recovery site using a backup control system in the F-104. The DAST Program experienced numerous problems. Only eighteen flights were achieved, eight of them captive (in which the aircraft flew only while still attached to the launch aircraft). Four of the flights were aborted and two resulted in crashes--one on June 12, 1980, and the second on June 1, 1983. Meanwhile, flight experiments with higher profiles, better funded remotely piloted research vehicles took priority over DAST missions. After the 1983 crash, which was caused by a malfunction that disconnected the landing parachute from the drone, the program was disbanded. Because DAST drones were considered expendable, certain losses were anticipated. Managers and researchers involved in other high-risk flight projects gained insights from the DAST program that could be applied to their own flight research programs. The DAST aircraft had a wingspan of 14 feet, four inches and a nose-to-tail length of 28 feet, 4 inches. The fuselage had a radius of about 2.07 feet. The aircraft's maximum loaded weight was about 2,200 pounds. It derived its power from a Continental YJ69-T-406 engine.

Why and Whither Hypersonics Research in the US Air Force

DTIC Science & Technology

2000-12-01

of strong interest in high supersonic and hypersonic flight. Actual flight achievements—notably the first supersonic flight by the XS - 1 on 14...Air Force is sponsoring Future Strike Aircraft studies with Boeing, Lockheed Martin, and Northrop Grumman . Under these studies, subsonic, supersonic...SMC/XRD) MSE Technology Applications, Inc. Northrop Grumman Air Combat Systems Orbital Sciences Corporation Boeing Phantom Works 18 May 2000

Needs, Effectiveness, and Gap Assessment for Key A-10C Missions: An Overview of Findings

DTIC Science & Technology

2016-01-01

weapons, further decreasing capacity. The GPS-guided SDB I ( GBU - 39 ) and the multimode SDB II ( GBU -53) begin to be highly useful in these circumstances...fairly close to friendly forces.8 It is specifically designed for use against moving targets; a Link 16 datalink permits target updates in flight, if...not jammed, and a laser-guidance mode allows it to be guided to specific targets if a JTAC is available to provide laser designation . It also has

Training monitoring skills in helicopter pilots.

PubMed

Potter, Brian A; Blickensderfer, Elizabeth L; Boquet, Albert J

2014-05-01

Prior research has indicated that ineffective pilot monitoring has been associated with aircraft accidents. Despite this finding, empirical research concerning pilot monitoring skill training programs is nearly nonexistent. E-learning may prove to be an effective method to foster nontechnical flight skills, including monitoring. This study examined the effect of using e-learning to enhance helicopter aircrew monitoring skill performance. The design was a posttest only field study. Forty-four helicopter pilots completed either an e-learning training module or a control activity and then flew two scenarios in a high-fidelity flight simulator. Learner reactions and knowledge gained were assessed immediately following the e-learning module. Two observer raters assessed behaviors and performance outcomes using recordings of the simulation flights. Subjects who completed the e-learning training module scored almost twice as high as did the control group on the administered knowledge test (experimental group, mean = 92.8%; control group, mean = 47.7%) and demonstrated up to 150% more monitoring behaviors during the simulated flights than the control subjects. In addition, the participating pilots rated the course highly. The results supported the hypothesis that a relatively inexpensive and brief training course implemented through e-learning can foster monitoring skill development among helicopter pilots.

A NASA painter applies the first primer coat to NASA's Orion full-scale abort flight test crew module in the Edwards Air Force Base paint hangar.

NASA Image and Video Library

2008-03-29

A full-scale flight-test mockup of the Constellation program's Orion crew vehicle arrived at NASA's Dryden Flight Research Center in late March 2008 to undergo preparations for the first short-range flight test of the spacecraft's astronaut escape system later that year. Engineers and technicians at NASA's Langley Research Center fabricated the structure, which precisely represents the size, outer shape and mass characteristics of the Orion space capsule. The Orion crew module mockup was ferried to NASA Dryden on an Air Force C-17. After painting in the Edwards Air Force Base paint hangar, the conical capsule was taken to Dryden for installation of flight computers, instrumentation and other electronics prior to being sent to the U.S. Army's White Sands Missile Range in New Mexico for integration with the escape system and the first abort flight test in late 2008. The tests were designed to ensure a safe, reliable method of escape for astronauts in case of an emergency.

Paint shop technicians carefully apply masking prior to painting the Orion full-scale abort flight test crew module in the Edwards Air Force Base paint hangar.

NASA Image and Video Library

2008-03-29

A full-scale flight-test mockup of the Constellation program's Orion crew vehicle arrived at NASA's Dryden Flight Research Center in late March 2008 to undergo preparations for the first short-range flight test of the spacecraft's astronaut escape system later that year. Engineers and technicians at NASA's Langley Research Center fabricated the structure, which precisely represents the size, outer shape and mass characteristics of the Orion space capsule. The Orion crew module mockup was ferried to NASA Dryden on an Air Force C-17. After painting in the Edwards Air Force Base paint hangar, the conical capsule was taken to Dryden for installation of flight computers, instrumentation and other electronics prior to being sent to the U.S. Army's White Sands Missile Range in New Mexico for integration with the escape system and the first abort flight test in late 2008. The tests were designed to ensure a safe, reliable method of escape for astronauts in case of an emergency.

NASA paint shop technicians prepare the Orion full-scale flight test crew module for painting in the Edwards Air Force Base paint hangar.

NASA Image and Video Library

2008-03-29

A full-scale flight-test mockup of the Constellation program's Orion crew vehicle arrived at NASA's Dryden Flight Research Center in late March 2008 to undergo preparations for the first short-range flight test of the spacecraft's astronaut escape system later that year. Engineers and technicians at NASA's Langley Research Center fabricated the structure, which precisely represents the size, outer shape and mass characteristics of the Orion space capsule. The Orion crew module mockup was ferried to NASA Dryden on an Air Force C-17. After painting in the Edwards Air Force Base paint hangar, the conical capsule was taken to Dryden for installation of flight computers, instrumentation and other electronics prior to being sent to the U.S. Army's White Sands Missile Range in New Mexico for integration with the escape system and the first abort flight test in late 2008. The tests were designed to ensure a safe, reliable method of escape for astronauts in case of an emergency.

Optimized merging of search coil and fluxgate data for MMS

NASA Astrophysics Data System (ADS)

Fischer, David; Magnes, Werner; Hagen, Christian; Dors, Ivan; Chutter, Mark W.; Needell, Jerry; Torbert, Roy B.; Le Contel, Olivier; Strangeway, Robert J.; Kubin, Gernot; Valavanoglou, Aris; Plaschke, Ferdinand; Nakamura, Rumi; Mirioni, Laurent; Russell, Christopher T.; Leinweber, Hannes K.; Bromund, Kenneth R.; Le, Guan; Kepko, Lawrence; Anderson, Brian J.; Slavin, James A.; Baumjohann, Wolfgang

2016-11-01

The Magnetospheric Multiscale mission (MMS) targets the characterization of fine-scale current structures in the Earth's tail and magnetopause. The high speed of these structures, when traversing one of the MMS spacecraft, creates magnetic field signatures that cross the sensitive frequency bands of both search coil and fluxgate magnetometers. Higher data quality for analysis of these events can be achieved by combining data from both instrument types and using the frequency bands with best sensitivity and signal-to-noise ratio from both sensors. This can be achieved by a model-based frequency compensation approach which requires the precise knowledge of instrument gain and phase properties. We discuss relevant aspects of the instrument design and the ground calibration activities, describe the model development and explain the application on in-flight data. Finally, we show the precision of this method by comparison of in-flight data. It confirms unity gain and a time difference of less than 100 µs between the different magnetometer instruments.

How Insects Initiate Flight: Computational Analysis of a Damselfly in Takeoff Flight

NASA Astrophysics Data System (ADS)

Bode-Oke, Ayodeji; Zeyghami, Samane; Dong, Haibo; Flow Simulation Research Group Team

2017-11-01

Flight initiation is essential for survival in biological fliers and can be classified into jumping and non-jumping takeoffs. During jumping takeoffs, the legs generate most of the initial impulse. Whereas the wings generate most of the forces in non-jumping takeoffs, which are usually voluntary, slow, and stable. It is of interest to understand how non-jumping takeoffs occur and what strategies insects use to generate the required forces. Using a high fidelity computational fluid dynamics simulation, we identify the flow features and compute the wing aerodynamic forces to elucidate how flight forces are generated by a damselfly performing a non-jumping takeoff. Our results show that a damselfly generates about three times its bodyweight during the first half-stroke for liftoff while flapping through a steeply inclined stroke plane and slicing the air at high angles of attack. Consequently, a Leading Edge Vortex (LEV) is formed during both the downstroke and upstroke on all the four wings. The formation of the LEV, however, is inhibited in the subsequent upstrokes following takeoff. Accordingly, we observe a drastic reduction in the magnitude of the aerodynamic force, signifying the importance of LEV in augmenting force production. This work was supported by National Science Foundation [CBET-1313217] and Air Force Research Laboratory [FA9550-12-1-007].

Determination of the STIS CCD Gain

NASA Astrophysics Data System (ADS)

Riley, Allyssa; Monroe, TalaWanda; Lockwood, Sean

2016-09-01

This report summarizes the analysis and absolute gain results of the STIS Cycle 23 special calibration program 14424 that was designed to measure the gain of amplifiers A, C and D at nominal gain settings of 1 and 4 e-/DN. We used the mean-variance technique and the results indicate a <3.5% change in the gain for amplifier D from when it was originally calculated pre-flight. We compared these values to previous measurements from Cycles 17 through 23. This report outlines the observations, methodology, and results of the mean-variance technique.

The servicing aid tool: A teleoperated robotics system for space applications

NASA Technical Reports Server (NTRS)

Dorman, Keith W.; Pullen, John L.; Keksz, William O.; Eismann, Paul H.; Kowalski, Keith A.; Karlen, James P.

1994-01-01

The Servicing Aid Tool (SAT) is a teleoperated, force-reflecting manipulation system designed for use on the Space Shuttle. The system will assist Extravehicular Activity (EVA) servicing of spacecraft such as the Hubble Space Telescope. The SAT stands out from other robotics development programs in that special attention was given to provide a low-cost, space-qualified design which can easily and inexpensively be reconfigured and/or enhanced through the addition of existing NASA funded technology as that technology matures. SAT components are spaceflight adaptations of existing ground-based designs from Robotics Research Corporation (RRC), the leading supplier of robotics systems to the NASA and university research community in the United States. Fairchild Space is the prime contractor and provides the control electronics, safety system, system integration, and qualification testing. The manipulator consists of a 6-DOF Slave Arm mounted on a 1-DOF Positioning Link in the shuttle payload bay. The Slave Arm is controlled via a highly similar, 6-DOF, force-reflecting Master Arm from Schilling Development, Inc. This work is being performed under contract to the Goddard Space Flight Center Code, Code 442, Hubble Space Telescope Flight Systems and Servicing Project.

Multi-objective/loading optimization for rotating composite flexbeams

NASA Technical Reports Server (NTRS)

Hamilton, Brian K.; Peters, James R.

1989-01-01

With the evolution of advanced composites, the feasibility of designing bearingless rotor systems for high speed, demanding maneuver envelopes, and high aircraft gross weights has become a reality. These systems eliminate the need for hinges and heavily loaded bearings by incorporating a composite flexbeam structure which accommodates flapping, lead-lag, and feathering motions by bending and twisting while reacting full blade centrifugal force. The flight characteristics of a bearingless rotor system are largely dependent on hub design, and the principal element in this type of system is the composite flexbeam. As in any hub design, trade off studies must be performed in order to optimize performance, dynamics (stability), handling qualities, and stresses. However, since the flexbeam structure is the primary component which will determine the balance of these characteristics, its design and fabrication are not straightforward. It was concluded that: pitchcase and snubber damper representations are required in the flexbeam model for proper sizing resulting from dynamic requirements; optimization is necessary for flexbeam design, since it reduces the design iteration time and results in an improved design; and inclusion of multiple flight conditions and their corresponding fatigue allowables is necessary for the optimization procedure.

A Gain-Programmable Transit-Time-Stable and Temperature-Stable PMT Voltage Divider

NASA Astrophysics Data System (ADS)

Liu, Yaqiang; Li, Hongdi; Wang, Yu; Xing, Tao; Xie, Shuping; Uribe, J.; Baghaei, H.; Ramirez, R.; Kim, Soonseok; Wong, Wai-Hoi

2004-10-01

A gain-programmable, transit-time-stable, temperature-stable photomultiplier (PMT) voltage divider design is described in this paper. The signal-to-noise ratio can be increased by changing a PMT gain directly instead of adjusting the gain of the preamplifier. PMT gain can be changed only by adjusting the voltages for the dynodes instead of changing the total high voltage between the anode and the photocathode, which can cause a significant signal transit-time variation that cannot be accepted by an application with a critical timing requirement, such as positron emission tomography (PET) or time-of-flight (TOF) detection/PET. The dynode voltage can be controlled by a digital analog converter isolated with a linear optocoupler. The optocoupler consists of an infrared light emission diode (LED) optically coupled with two phototransistors, and one is used in a servo feedback circuit to control the LED drive current for compensating temperature characteristics. The results showed that a six times gain range could be achieved; the gain drift was <0.5% over a 20/spl deg/C temperature range; 250 ps transit-time variation was measured over the entire gain range. A compact print circuit board (PCB) for the voltage divider integrated with a fixed-gain preamplifier has been designed and constructed. It can save about $30 per PMT channel compared with a commercial PMT voltage divider along with a variable gain amplifier. The preamplifier can be totally disabled, therefore in a system with a large amount of PMTs, only one channel can be enabled for calibrating the PMT gain. This new PMT voltage divider design is being applied to our animal PET camera and TOF/PET research.

USAF bioenvironmental noise data handbook. Volume 157: KC-10A in-flight crew noise

NASA Astrophysics Data System (ADS)

Hille, H. K.

1982-09-01

The KC-10A is a standard USAF tanker-transport aircraft with high-speed, high altitude refueling and long range transport capability. This report provides measured data defining the bioacoustic environments at flight crew/passenger locations inside this helicopter during normal flight operations. Data are reported for 24 locations in a wide variety of physical and psychoacoustic measures: overall and band sound pressure levels, C-weighted and A-weighted sound levels, preferred speech interference level, perceived noise level, and limiting times for total daily exposure of personnel with and without standard Air Force ear protectors. Refer to Volume 1 of this handbook, USAF Bioenvironmental Noise Data Handbook, Vol. 1: Organization, Content and Application, AMRL-TR-75-50(1) 1975, for discussion of the objective and design of the handbook, the types of data presented, measurement procedures, instrumentation, data processing, definitions of quantities, symbols, equations, applications, limitations, etc.

Robust, Practical Adaptive Control for Launch Vehicles

NASA Technical Reports Server (NTRS)

Orr, Jeb. S.; VanZwieten, Tannen S.

2012-01-01

A modern mechanization of a classical adaptive control concept is presented with an application to launch vehicle attitude control systems. Due to a rigorous flight certification environment, many adaptive control concepts are infeasible when applied to high-risk aerospace systems; methods of stability analysis are either intractable for high complexity models or cannot be reconciled in light of classical requirements. Furthermore, many adaptive techniques appearing in the literature are not suitable for application to conditionally stable systems with complex flexible-body dynamics, as is often the case with launch vehicles. The present technique is a multiplicative forward loop gain adaptive law similar to that used for the NASA X-15 flight research vehicle. In digital implementation with several novel features, it is well-suited to application on aerodynamically unstable launch vehicles with thrust vector control via augmentation of the baseline attitude/attitude-rate feedback control scheme. The approach is compatible with standard design features of autopilots for launch vehicles, including phase stabilization of lateral bending and slosh via linear filters. In addition, the method of assessing flight control stability via classical gain and phase margins is not affected under reasonable assumptions. The algorithm s ability to recover from certain unstable operating regimes can in fact be understood in terms of frequency-domain criteria. Finally, simulation results are presented that confirm the ability of the algorithm to improve performance and robustness in realistic failure scenarios.

POST II Trajectory Animation Tool Using MATLAB, V1.0

NASA Technical Reports Server (NTRS)

Raiszadeh, Behzad

2005-01-01

A trajectory animation tool has been developed for accurately depicting position and the attitude of the bodies in flight. The movies generated from This MATLAB based tool serve as an engineering analysis aid to gain further understanding into the dynamic behavior of bodies in flight. This tool has been designed to interface with the output generated from POST II simulations, and is able to animate a single as well as multiple vehicles in flight.

Flight Testing an Iced Business Jet for Flight Simulation Model Validation

NASA Technical Reports Server (NTRS)

Ratvasky, Thomas P.; Barnhart, Billy P.; Lee, Sam; Cooper, Jon

2007-01-01

A flight test of a business jet aircraft with various ice accretions was performed to obtain data to validate flight simulation models developed through wind tunnel tests. Three types of ice accretions were tested: pre-activation roughness, runback shapes that form downstream of the thermal wing ice protection system, and a wing ice protection system failure shape. The high fidelity flight simulation models of this business jet aircraft were validated using a software tool called "Overdrive." Through comparisons of flight-extracted aerodynamic forces and moments to simulation-predicted forces and moments, the simulation models were successfully validated. Only minor adjustments in the simulation database were required to obtain adequate match, signifying the process used to develop the simulation models was successful. The simulation models were implemented in the NASA Ice Contamination Effects Flight Training Device (ICEFTD) to enable company pilots to evaluate flight characteristics of the simulation models. By and large, the pilots confirmed good similarities in the flight characteristics when compared to the real airplane. However, pilots noted pitch up tendencies at stall with the flaps extended that were not representative of the airplane and identified some differences in pilot forces. The elevator hinge moment model and implementation of the control forces on the ICEFTD were identified as a driver in the pitch ups and control force issues, and will be an area for future work.

Perseus B Heads for Landing on Edwards AFB Runway

NASA Image and Video Library

1998-04-30

The Perseus B remotely piloted aircraft approaches the runway at Edwards Air Force Base, Calif. at the conclusion of a development flight at NASA's Dryden flight Research Center in April 1998. The Perseus B is the latest of three versions of the Perseus design developed by Aurora Flight Sciences under NASA's Environmental Research Aircraft and Sensor Technology (ERAST) program.

Balloon flight test of a Compton telescope based on scintillators with silicon photomultiplier readouts

NASA Astrophysics Data System (ADS)

Bloser, P. F.; Legere, J. S.; Bancroft, C. M.; Ryan, J. M.; McConnell, M. L.

2016-03-01

We present the results of the first high-altitude balloon flight test of a concept for an advanced Compton telescope making use of modern scintillator materials with silicon photomultiplier (SiPM) readouts. There is a need in the fields of high-energy astronomy and solar physics for new medium-energy gamma-ray ( 0.4-10 MeV) detectors capable of making sensitive observations of both line and continuum sources over a wide dynamic range. A fast scintillator-based Compton telescope with SiPM readouts is a promising solution to this instrumentation challenge, since the fast response of the scintillators permits both the rejection of background via time-of-flight (ToF) discrimination and the ability to operate at high count rates. The Solar Compton Telescope (SolCompT) prototype presented here was designed to demonstrate stable performance of this technology under balloon-flight conditions. The SolCompT instrument was a simple two-element Compton telescope, consisting of an approximately one-inch cylindrical stilbene crystal for a scattering detector and a one-inch cubic LaBr3:Ce crystal for a calorimeter detector. Both scintillator detectors were read out by 2×2 arrays of Hamamatsu S11828-3344 MPPC devices. Custom front-end electronics provided optimum signal rise time and linearity, and custom power supplies automatically adjusted the SiPM bias voltage to compensate for temperature-induced gain variations. A tagged calibration source, consisting of 240 nCi of 60Co embedded in plastic scintillator, was placed in the field of view and provided a known source of gamma rays to measure in flight. The SolCompT balloon payload was launched on 24 August 2014 from Fort Sumner, NM, and spent 3.75 h at a float altitude of 123,000 ft. The instrument performed well throughout the flight. After correcting for small ( 10%) residual gain variations, we measured an in-flight ToF resolution of 760 ps (FWHM). Advanced scintillators with SiPM readouts continue to show great promise for future gamma-ray instruments.

F-18 HARV yaw rate expansion flight #125 with Inverted Recovery

NASA Technical Reports Server (NTRS)

1991-01-01

NASA's Dryden Flight Research Center, Edwards, CA, used an F-18 Hornet fighter aircraft as its High Angle-of-Attack (Alpha) Research Vehicle (HARV) in a three-phased flight research program lasting from April 1987 until September 1996. The aircraft completed 385 research flights and demonstrated stabilized flight at angles of attack between 65 and 70 degrees using thrust vectoring vanes, a research flight control system, and (eventually) forebody strakes (hinged structures on the forward side of the fuselage to provide control by interacting with vortices, generated at high angles of attack, to create side forces). This combination of technologies provided carefree handling of a fighter aircraft in a part of the flight regime that was otherwise very dangerous. Flight research with the HARV increased our understanding of flight at high angles of attack (angle of the wings with respect to the direction in which the aircraft was heading), enabling designers of U.S. fighter aircraft to design airplanes that will fly safely in portions of the flight envelope that pilots previously had to avoid. Flight 125 with the HARV involved yaw rate expansion up to 50 degrees per second (moving the nose to the left or right at that rate). NASA research pilot Ed Schneider was the pilot, and the purpose of the flight was to look at the spin characteristics of the HARV. The sequence in this particular video clip includes the first and second maneuvers in the flight. On the first maneuver, the pilot attempted to achieve a yaw rate of 40 degrees per second and actually went to 47 degrees. The spin was oscillatory in pitch (up and down) and roll (rotating around the longitudinal axis). Recovery was normal. On the second maneuver of the flight in which Schneider tried to achieve a yaw rate of 40 degrees per second, the aircraft overshot to 54 degrees per second during an oscillatory spin. In the course of the recovery, the aircraft rolled after a large sideslip buildup. Moderate aft stick application to attain a positive angle of attack resulted in an easy recovery.

X-36 in Flight near Edge of Rogers Dry Lake during 5th Flight

NASA Technical Reports Server (NTRS)

1997-01-01

This photo shows the X-36 Tailless Fighter Agility Research Aircraft passing over the edge of Rogers Dry Lake as the remotely-piloted aircraft flies over Edwards Air Force Base on a June 1997 research flight. The NASA/Boeing X-36 Tailless Fighter Agility Research Aircraft program successfully demonstrated the tailless fighter design using advanced technologies to improve the maneuverability and survivability of possible future fighter aircraft. The program met or exceeded all project goals. For 31 flights during 1997 at the Dryden Flight Research Center, Edwards, California, the project team examined the aircraft's agility at low speed / high angles of attack and at high speed / low angles of attack. The aircraft's speed envelope reached up to 206 knots (234 mph). This aircraft was very stable and maneuverable. It handled very well. The X-36 vehicle was designed to fly without the traditional tail surfaces common on most aircraft. Instead, a canard forward of the wing was used as well as split ailerons and an advanced thrust-vectoring nozzle for directional control. The X-36 was unstable in both pitch and yaw axes, so an advanced, single-channel digital fly-by-wire control system (developed with some commercially available components) was put in place to stabilize the aircraft. Using a video camera mounted in the nose of the aircraft and an onboard microphone, the X-36 was remotely controlled by a pilot in a ground station virtual cockpit. A standard fighter-type head-up display (HUD) and a moving-map representation of the vehicle's position within the range in which it flew provided excellent situational awareness for the pilot. This pilot-in-the-loop approach eliminated the need for expensive and complex autonomous flight control systems and the risks associated with their inability to deal with unknown or unforeseen phenomena in flight. Fully fueled the X-36 prototype weighed approximately 1,250 pounds. It was 19 feet long and three feet high with a wingspan of just over 10 feet. A Williams International F112 turbofan engine provided close to 700 pounds of thrust. A typical research flight lasted 35 to 45 minutes from takeoff to touchdown. A total of 31 successful research flights were flown from May 17, 1997, to November 12, 1997, amassing 15 hours and 38 minutes of flight time. The aircraft reached an altitude of 20,200 feet and a maximum angle of attack of 40 degrees. In a follow-on effort, the Air Force Research Laboratory (AFRL), Wright-Patterson Air Force Base, Ohio, contracted with Boeing to fly AFRL's Reconfigurable Control for Tailless Fighter Aircraft (RESTORE) software as a demonstration of the adaptability of the neural-net algorithm to compensate for in-flight damage or malfunction of effectors, such as flaps, ailerons and rudders. Two RESTORE research flights were flown in December 1998, proving the viability of the software approach. The X-36 aircraft flown at the Dryden Flight Research Center in 1997 was a 28-percent scale representation of a theoretical advanced fighter aircraft. The Boeing Phantom Works (formerly McDonnell Douglas) in St. Louis, Missouri, built two of the vehicles in a cooperative agreement with the Ames Research Center, Moffett Field, California.

Development of a prototype two-phase thermal bus system for Space Station

NASA Technical Reports Server (NTRS)

Myron, D. L.; Parish, R. C.

1987-01-01

This paper describes the basic elements of a pumped two-phase ammonia thermal control system designed for microgravity environments, the development of the concept into a Space Station flight design, and design details of the prototype to be ground-tested in the Johnson Space Center (JSC) Thermal Test Bed. The basic system concept is one of forced-flow heat transport through interface heat exchangers with anhydrous ammonia being pumped by a device expressly designed for two-phase fluid management in reduced gravity. Control of saturation conditions, and thus system interface temperatures, is accomplished with a single central pressure regulating valve. Flow control and liquid inventory are controlled by passive, nonelectromechanical devices. Use of these simple control elements results in minimal computer controls and high system reliability. Building on the basic system concept, a brief overview of a potential Space Station flight design is given. Primary verification of the system concept will involve testing at JSC of a 25-kW ground test article currently in fabrication.

The Role of Formal Experiment Design in Hypersonic Flight System Technology Development

NASA Technical Reports Server (NTRS)

McClinton, Charles R.; Ferlemann, Shelly M.; Rock, Ken E.; Ferlemann, Paul G.

2002-01-01

Hypersonic airbreathing engine (scramjet) powered vehicles are being considered to replace conventional rocket-powered launch systems. Effective utilization of scramjet engines requires careful integration with the air vehicle. This integration synergistically combines aerodynamic forces with propulsive cycle functions of the engine. Due to the highly integrated nature of the hypersonic vehicle design problem, the large flight envelope, and the large number of design variables, the use of a statistical design approach in design is effective. Modern Design-of-Experiments (MDOE) has been used throughout the Hyper-X program, for both systems analysis and experimental testing. Application of MDOE fall into four categories: (1) experimental testing; (2) studies of unit phenomena; (3) refining engine design; and (4) full vehicle system optimization. The MDOE process also provides analytical models, which are also used to document lessons learned, supplement low-level design tools, and accelerate future studies. This paper will discuss the design considerations for scramjet-powered vehicles, specifics of MDOE utilized for Hyper-X, and present highlights from the use of these MDOE methods within the Hyper-X Program.

Eclipse program C-141A aircraft

NASA Technical Reports Server (NTRS)

1997-01-01

This photograph shows the Air Force C-141A that was used in the Eclipse project as a tow vehicle. The project used a QF-106 interceptor aircraft to simulate a future orbiter, which would be towed to a high altitude and released to fire its own engines and carry a payload into space. In 1997 and 1998, the Dryden Flight Research Center at Edwards, California, supported and hosted a Kelly Space & Technology, Inc. project called Eclipse, which sought to demonstrate the feasibility of a reusable tow-launch vehicle concept. The project goal was to successfully tow, inflight, a modified QF-106 delta-wing aircraft with an Air Force C-141A transport aircraft. This would demonstrate the possibility of towing and launching an actual launch vehicle from behind a tow plane. Dryden was the responsible test organization and had flight safety responsibility for the Eclipse project. Dryden provided engineering, instrumentation, simulation, modification, maintenance, range support, and research pilots for the test program. The Air Force Flight Test Center (AFFTC), Edwards, California, supplied the C-141A transport aircraft and crew and configured the aircraft as needed for the tests. The AFFTC also provided the concept and detail design and analysis as well as hardware for the tow system and QF-106 modifications. Dryden performed the modifications to convert the QF-106 drone into the piloted EXD-01 (Eclipse eXperimental Demonstrator-01) experimental aircraft. Kelly Space & Technology hoped to use the results gleaned from the tow test in developing a series of low-cost, reusable launch vehicles. These tests demonstrated the validity of towing a delta-wing aircraft having high wing loading, validated the tow simulation model, and demonstrated various operational procedures, such as ground processing of in-flight maneuvers and emergency abort scenarios.

The Lifting Body Legacy...X-33

NASA Technical Reports Server (NTRS)

Barret, Chris

1999-01-01

NASA has a technology program in place to enable the development of a next generation Reusable Launch Vehicle that will carry our future payloads into orbit at a much-reduced cost. The VentureStar, Lifting Body (LB) flight vehicle, is one of the potential reusable launch vehicle configurations being studied. A LB vehicle has no wings and derives its lift solely from the shape of its body, and has the unique advantages of superior volumetric efficiency, better aerodynamic efficiency at high angles-of-attack and hypersonic speeds, and reduced thermal protection system weight. Classically, in a ballistic vehicle, drag has been employed to control the level of deceleration in reentry. In the LB, lift enables the vehicle to decelerate at higher altitudes for the same velocity and defines the reentry corridor which includes a greater cross range. This paper outlines the flight stability and control aspects of our LB heritage which was utilized in the design of the VentureStar LB and its test version, the X-33. NASA and the U.S. Air Force have a rich heritage of LB vehicle design and flight experience. In the initial LB Program, eight LB's were built and over 225 LB test flights were conducted through 1975. Three LB series were most significant in the advancement of today's LB technolocy: the M2-F; the HL-10; and the X-24 series. The M2-F series was designed by NASA Ames Research Center, the HL-10 series by NASA Langley Research Center, and the X-24 series by the U. S. Air Force. LB vehicles are alive again today with the X- 33, X-38, and VentureStar.

Comparison of Ares I-X Wind-Tunnel Derived Buffet Environment with Flight Data

NASA Technical Reports Server (NTRS)

Piatak, David J.; Sekula, Martin K.; Rausch, Russ D.

2011-01-01

The Ares I-X Flight Test Vehicle (FTV), launched in October 2009, carried with it over 243 buffet verification pressure sensors and was one of the most heavily instrumented launch vehicle flight tests. This flight test represented a unique opportunity for NASA and its partners to compare the wind-tunnel derived buffet environment with that measured during the flight of Ares I-X. It is necessary to define the launch vehicle buffet loads to ensure that structural components and vehicle subsystems possess adequate strength, stress, and fatigue margins when the vehicle structural dynamic response to buffet forcing functions are considered. Ares I-X buffet forcing functions were obtained via wind-tunnel testing of a rigid buffet model (RBM) instrumented with hundreds of unsteady pressure transducers designed to measure the buffet environment across the desired frequency range. This paper discusses the comparison of RBM and FTV buffet environments, including fluctuating pressure coefficient and normalized sectional buffet forcing function root-mean-square magnitudes, frequency content of power-spectral density functions, and force magnitudes of an alternating flow phenomena. Comparison of wind-tunnel model and flight test vehicle buffet environments show very good agreement with root-mean-square magnitudes of buffet forcing functions at the majority of vehicle stations. Spectra proved a challenge to compare because of different wind-tunnel and flight test conditions and data acquisition rates. However, meaningful and promising comparisons of buffet spectra are presented. Lastly, the buffet loads resulting from the transition of subsonic separated flow to supersonic attached flow were significantly over-predicted by wind-tunnel results.

X-15 and XB-70 parked on NASA ramp

NASA Technical Reports Server (NTRS)

1967-01-01

The X-15A-2 with drop tanks and ablative coating is shown parked on the NASA ramp in front of the XB-70. These aircraft represent two different approaches to flight research. The X-15 was a research airplane in the purest sense, whereas the XB-70 was an experimental bomber intended for production but diverted to research when production was cancelled by changes in the Department of Defense's offensive doctrine. The X-15A-2 had been modified from its original configuration with a longer fuselage and drop tanks. To protect it against aerodynamic heating, researchers had coated it with an ablative coating covered by a layer of white paint. These changes allowed the X-15A-2 to reach a maximum speed of Mach 6.7, although it could be sustained for only a brief period. The XB-70, by contrast, was designed for prolonged high-altitude cruise flight at Mach 3. The aircraft's striking shape--with a long forward fuselage, canards, a large delta wing, twin fins, and a box-like engine bay--allowed it to ride its own Mach 3 shockwave, so to speak. A joint NASA-Air Force program used the aircraft to collect data in support of the U.S supersonic transport (SST) program, which never came to fruition because of environmental concerns. X-15: The X-15 was a rocket-powered aircraft. The original three aircraft were about 50 ft long with a wingspan of 22 ft. The modified #2 aircraft (X-15A-2 was longer.) They were a missile-shaped vehicles with unusual wedge-shaped vertical tails, thin stubby wings, and unique side fairings that extended along the side of the fuselage. The X-15 weighed about 14,000 lb empty and approximately 34,000 lb at launch. The XLR-99 rocket engine, manufactured by Thiokol Chemical Corp., was pilot controlled and was rated at 57,000 lb of thrust, although there are indications that it actually achieved up to 60,000 lb. North American Aviation built three X-15 aircraft for the program. The X-15 research aircraft was developed to provide in-flight information and data on aerodynamics, structures, flight controls, and the physiological aspects of high-speed, high-altitude flight. A follow-on program used the aircraft as testbeds to carry various scientific experiments beyond the Earth's atmosphere on a repeated basis. For flight in the dense air of the usable atmosphere, the X-15 used conventional aerodynamic controls such as rudder surfaces on the vertical stabilizers to control yaw and movable horizontal stabilizers to control pitch when moving in synchronization or roll when moved differentially. For flight in the thin air outside of the appreciable Earth's atmosphere, the X-15 used a reaction control system. Hydrogen peroxide thrust rockets located on the nose of the aircraft provided pitch and yaw control. Those on the wings provided roll control. Because of the large fuel consumption, the X-15 was air launched from a B-52 aircraft at approximately 45,000 ft and a speed of about 500 mph. Depending on the mission, the rocket engine provided thrust for the first 80 to 120 sec of flight. The remainder of the normal 10 to 11 min. flight was powerless and ended with a 200-mph glide landing. Generally, one of two types of X-15 flight profiles was used; a high-altitude flight plan that called for the pilot to maintain a steep rate of climb, or a speed profile that called for the pilot to push over and maintain a level altitude. The X-15 was flown over a period of nearly 10 years -- June 1959 to Oct. 1968 -- and set the world's unofficial speed and altitude records of 4,520 mph (Mach 6.7) and 354,200 ft in a program to investigate all aspects of manned hypersonic flight. Information gained from the highly successful X-15 program contributed to the development of the Mercury, Gemini, and Apollo manned spaceflight programs, and also the Space Shuttle program. The X-15s made a total of 199 flights, and were manufactured by North American Aviation. X-15-1, serial number 56-6670, is now located at the National Air and Space Museum, Washington DC. North American X-15A-2, serial number 56-6671, is at the United States Air Force Museum, Wright-Patterson AFB, Ohio. X-15-3, serial number 56-6672, crashed on 15 November 1967, resulting in the death of Maj. Michael J. Adams. XB-70: The XB-70 was the world's largest experimental aircraft. It was capable of flight at speeds of three times the speed of sound (roughly 2,000 miles per hour) at altitudes of 70,000 feet. It was used to collect in-flight information for use in the design of future supersonic aircraft, military and civilian. The major objectives of the XB-70 flight research program were to study the airplane's stability and handling characteristics, to evaluate its response to atmospheric turbulence, and to determine the aerodynamic and propulsion performance. In addition there were secondary objectives to measure the noise and friction associated with airflow over the airplane and to determine the levels and extent of the engine noise during takeoff, landing, and ground operations. The XB-70 was about 186 feet long, 33 feet high, with a wingspan of 105 feet. Originally conceived as an advanced bomber for the United States Air Force, the XB-70 was limited to production of two aircraft when it was decided to limit the aircraft's mission to flight research. The first flight of the XB-70 was made on Sept. 21, 1964. The number two XB-70 was destroyed in a mid-air collision on June 8, 1966. Program management of the NASA-USAF research effort was assigned to NASA in March 1967. The final flight was flown on Feb. 4, 1969. Designed by North American Aviation (later North American Rockwell and still later, a division of Boeing) the XB-70 had a long fuselage with a canard or horizontal stabilizer mounted just behind the crew compartment. It had a sharply swept 65.6-percent delta wing. The outer portion of the wing could be folded down in flight to provide greater lateral-directional stability. The airplane had two windshields. A moveable outer windshield was raised for high-speed flight to reduce drag and lowered for greater visibility during takeoff and landing. The forward fuselage was constructed of riveted titanium frames and skin. The remainder of the airplane was constructed almost entirely of stainless steel. The skin was a brazed stainless-steel honeycomb material. Six General Electric YJ93-3 turbojet engines, each in the 30,000-pound-thrust class, powered the XB-70. Internal geometry of the inlets was controllable to maintain the most efficient airflow to the engines.

Advanced Smart Structures Flight Experiments for Precision Spacecraft

NASA Astrophysics Data System (ADS)

Denoyer, Keith K.; Erwin, R. Scott; Ninneman, R. Rory

2000-07-01

This paper presents an overview as well as data from four smart structures flight experiments directed by the U.S. Air Force Research Laboratory's Space Vehicles Directorate in Albuquerque, New Mexico. The Middeck Active Control Experiment $¯Flight II (MACE II) is a space shuttle flight experiment designed to investigate modeling and control issues for achieving high precision pointing and vibration control of future spacecraft. The Advanced Controls Technology Experiment (ACTEX-I) is an experiment that has demonstrated active vibration suppression using smart composite structures with embedded piezoelectric sensors and actuators. The Satellite Ultraquiet Isolation Technology Experiment (SUITE) is an isolation platform that uses active piezoelectric actuators as well as damped mechanical flexures to achieve hybrid passive/active isolation. The Vibration Isolation, Suppression, and Steering Experiment (VISS) is another isolation platform that uses viscous dampers in conjunction with electromagnetic voice coil actuators to achieve isolation as well as a steering capability for an infra-red telescope.

Control Design for an Advanced Geared Turbofan Engine

NASA Technical Reports Server (NTRS)

Chapman, Jeffryes W.; Litt, Jonathan S.

2017-01-01

This paper describes the design process for the control system of an advanced geared turbofan engine. This process is applied to a simulation that is representative of a 30,000 pound-force thrust class concept engine with two main spools, ultra-high bypass ratio, and a variable area fan nozzle. Control system requirements constrain the non-linear engine model as it operates throughout its flight envelope of sea level to 40,000 feet and from 0 to 0.8 Mach. The purpose of this paper is to review the engine control design process for an advanced turbofan engine configuration. The control architecture selected for this project was developed from literature and reflects a configuration that utilizes a proportional integral controller with sets of limiters that enable the engine to operate safely throughout its flight envelope. Simulation results show the overall system meets performance requirements without exceeding operational limits.

An Overview of Flight Test Results for a Formation Flight Autopilot

NASA Technical Reports Server (NTRS)

Hanson, Curtis E.; Ryan, Jack; Allen, Michael J.; Jacobson, Steven R.

2002-01-01

The first flight test phase of the NASA Dryden Flight Research Center Autonomous Formation Flight project has successfully demonstrated precision autonomous station-keeping of an F/A-18 research airplane with a second F/A-18 airplane. Blended inertial navigation system (INS) and global positioning system (GPS) measurements have been communicated across an air-to-air telemetry link and used to compute relative-position estimates. A precision research formation autopilot onboard the trailing airplane controls lateral and vertical spacing while the leading airplane operates under production autopilot control. Four research autopilot gain sets have been designed and flight-tested, and each exceeds the project design requirement of steady-state tracking accuracy within 1 standard deviation of 10 ft. Performance also has been demonstrated using single- and multiple-axis inputs such as step commands and frequency sweeps. This report briefly describes the experimental formation flight systems employed and discusses the navigation, guidance, and control algorithms that have been flight-tested. An overview of the flight test results of the formation autopilot during steady-state tracking and maneuvering flight is presented.

Effects of side-stick controllers on rotorcraft handling qualities for terrain flight

NASA Technical Reports Server (NTRS)

Aiken, E. W.

1985-01-01

Pertinent fixed and rotary-wing feasibility studies and handling-qualities research programs are reviewed and the effects of certain controller characteristics on handling qualities for specific rotorcraft flight tasks are summarized. The effects of the controller force-deflection relationship and the number of controlled axes that are integrated in a single controller are examined. Simulation studies were conducted which provide a significant part of the available handling qualities data. The studies demonstrate the feasibility of using a single, properly designed, limited-displacement, multiaxis controller for certain relatively routine flight tasks in a two-crew rotorcraft with nominal levels of stability and control augmentation with a high degree of reliability are incorporated, separated three or two-axis controller configurations are required for acceptable handling qualities.

Load measurement system with load cell lock-out mechanism

NASA Technical Reports Server (NTRS)

Le, Thang; Carroll, Monty; Liu, Jonathan

1995-01-01

In the frame work of the project Shuttle Plume Impingement Flight Experiment (SPIFEX), a Load Measurement System was developed and fabricated to measure the impingement force of Shuttle Reaction Control System (RCS) jets. The Load Measurement System is a force sensing system that measures any combination of normal and shear forces up to 40 N (9 lbf) in the normal direction and 22 N (5 lbf) in the shear direction with an accuracy of +/- 0.04 N (+/- 0.01 lbf) Since high resolution is required for the force measurement, the Load Measurement System is built with highly sensitive load cells. To protect these fragile load cells in the non-operational mode from being damaged due to flight loads such as launch and landing loads of the Shuttle vehicle, a motor driven device known as the Load Cell Lock-Out Mechanism was built. This Lock-Out Mechanism isolates the load cells from flight loads and re-engages the load cells for the force measurement experiment once in space. With this highly effective protection system, the SPIFEX load measurement experiment was successfully conducted on STS-44 in September 1994 with all load cells operating properly and reading impingement forces as expected.

Aerodynamic effects of flexibility in flapping wings.

PubMed

Zhao, Liang; Huang, Qingfeng; Deng, Xinyan; Sane, Sanjay P

2010-03-06

Recent work on the aerodynamics of flapping flight reveals fundamental differences in the mechanisms of aerodynamic force generation between fixed and flapping wings. When fixed wings translate at high angles of attack, they periodically generate and shed leading and trailing edge vortices as reflected in their fluctuating aerodynamic force traces and associated flow visualization. In contrast, wings flapping at high angles of attack generate stable leading edge vorticity, which persists throughout the duration of the stroke and enhances mean aerodynamic forces. Here, we show that aerodynamic forces can be controlled by altering the trailing edge flexibility of a flapping wing. We used a dynamically scaled mechanical model of flapping flight (Re approximately 2000) to measure the aerodynamic forces on flapping wings of variable flexural stiffness (EI). For low to medium angles of attack, as flexibility of the wing increases, its ability to generate aerodynamic forces decreases monotonically but its lift-to-drag ratios remain approximately constant. The instantaneous force traces reveal no major differences in the underlying modes of force generation for flexible and rigid wings, but the magnitude of force, the angle of net force vector and centre of pressure all vary systematically with wing flexibility. Even a rudimentary framework of wing veins is sufficient to restore the ability of flexible wings to generate forces at near-rigid values. Thus, the magnitude of force generation can be controlled by modulating the trailing edge flexibility and thereby controlling the magnitude of the leading edge vorticity. To characterize this, we have generated a detailed database of aerodynamic forces as a function of several variables including material properties, kinematics, aerodynamic forces and centre of pressure, which can also be used to help validate computational models of aeroelastic flapping wings. These experiments will also be useful for wing design for small robotic insects and, to a limited extent, in understanding the aerodynamics of flapping insect wings.

Aerodynamic effects of flexibility in flapping wings

PubMed Central

Zhao, Liang; Huang, Qingfeng; Deng, Xinyan; Sane, Sanjay P.

2010-01-01

Recent work on the aerodynamics of flapping flight reveals fundamental differences in the mechanisms of aerodynamic force generation between fixed and flapping wings. When fixed wings translate at high angles of attack, they periodically generate and shed leading and trailing edge vortices as reflected in their fluctuating aerodynamic force traces and associated flow visualization. In contrast, wings flapping at high angles of attack generate stable leading edge vorticity, which persists throughout the duration of the stroke and enhances mean aerodynamic forces. Here, we show that aerodynamic forces can be controlled by altering the trailing edge flexibility of a flapping wing. We used a dynamically scaled mechanical model of flapping flight (Re ≈ 2000) to measure the aerodynamic forces on flapping wings of variable flexural stiffness (EI). For low to medium angles of attack, as flexibility of the wing increases, its ability to generate aerodynamic forces decreases monotonically but its lift-to-drag ratios remain approximately constant. The instantaneous force traces reveal no major differences in the underlying modes of force generation for flexible and rigid wings, but the magnitude of force, the angle of net force vector and centre of pressure all vary systematically with wing flexibility. Even a rudimentary framework of wing veins is sufficient to restore the ability of flexible wings to generate forces at near-rigid values. Thus, the magnitude of force generation can be controlled by modulating the trailing edge flexibility and thereby controlling the magnitude of the leading edge vorticity. To characterize this, we have generated a detailed database of aerodynamic forces as a function of several variables including material properties, kinematics, aerodynamic forces and centre of pressure, which can also be used to help validate computational models of aeroelastic flapping wings. These experiments will also be useful for wing design for small robotic insects and, to a limited extent, in understanding the aerodynamics of flapping insect wings. PMID:19692394

Sensitivity analysis of the space shuttle to ascent wind profiles

NASA Technical Reports Server (NTRS)

Smith, O. E.; Austin, L. D., Jr.

1982-01-01

A parametric sensitivity analysis of the space shuttle ascent flight to the wind profile is presented. Engineering systems parameters are obtained by flight simulations using wind profile models and samples of detailed (Jimsphere) wind profile measurements. The wind models used are the synthetic vector wind model, with and without the design gust, and a model of the vector wind change with respect to time. From these comparison analyses an insight is gained on the contribution of winds to ascent subsystems flight parameters.

Morpheus Lander Roll Control System and Wind Modeling

NASA Technical Reports Server (NTRS)

Gambone, Elisabeth A.

2014-01-01

The Morpheus prototype lander is a testbed capable of vertical takeoff and landing developed by NASA Johnson Space Center to assess advanced space technologies. Morpheus completed a series of flight tests at Kennedy Space Center to demonstrate autonomous landing and hazard avoidance for future exploration missions. As a prototype vehicle being tested in Earth's atmosphere, Morpheus requires a robust roll control system to counteract aerodynamic forces. This paper describes the control algorithm designed that commands jet firing and delay times based on roll orientation. Design, analysis, and testing are supported using a high fidelity, 6 degree-of-freedom simulation of vehicle dynamics. This paper also details the wind profiles generated using historical wind data, which are necessary to validate the roll control system in the simulation environment. In preparation for Morpheus testing, the wind model was expanded to create day-of-flight wind profiles based on data delivered by Kennedy Space Center. After the test campaign, a comparison of flight and simulation performance was completed to provide additional model validation.

Integration and Test Flight Validation Plans for the Pulsed Plasma Thruster Experiment on EO- 1

NASA Technical Reports Server (NTRS)

Zakrzwski, Charles; Benson, Scott; Sanneman, Paul; Hoskins, Andy; Bauer, Frank H. (Technical Monitor)

2002-01-01

The Pulsed Plasma Thruster (PPT) Experiment on the Earth Observing One (EO-1) spacecraft has been designed to demonstrate the capability of a new generation PPT to perform spacecraft attitude control. The PPT is a small, self-contained pulsed electromagnetic propulsion system capable of delivering high specific impulse (900-1200 s), very small impulse bits (10-1000 uN-s) at low average power (less than 1 to 100 W). Teflon fuel is ablated and slightly ionized by means of a capacitative discharge. The discharge also generates electromagnetic fields that accelerate the plasma by means of the Lorentz Force. EO-1 has a single PPT that can produce thrust in either the positive or negative pitch direction. The flight validation has been designed to demonstrate of the ability of the PPT to provide precision pointing accuracy, response and stability, and confirmation of benign plume and EMI effects. This paper will document the success of the flight validation.

Effects of Stable and Unstable Resistance Training in an Altered-G Environment on Muscle Power.

PubMed

Zemková, E; Oddsson, L

2016-04-01

The study evaluated the effect of 4 weeks of combined resistance-balance training and resistance training alone in a 90° tilted environment on muscle power. Two groups of healthy young subjects performed leg extensions while in a supine position, either on a firm surface along a linear track or on an unstable surface requiring mediolateral balancing movements. Power and force during squats were measured at isokinetic velocities of 10 and 35 deg/s. Results showed significantly greater gains in peak force (44.1%; F(1,21)=8.876, p=0.026), mean force (58.6%; F(1,21)=16.136, p=0.013), peak power (58.7%; F(1,21)=18.754, p=0.009), and mean power (59.2%; F(1,21)=23.114, p=0.007) at the velocity of 35 deg/s after stable than unstable resistance training. However, there were no significant between-groups differences in pre-post training gains in peak force (10.4%; F(1,21)=1.965, p=0.74), mean force (10.3%; F(1,21)=1.889, p=0.80), peak power (12.9%; F(1,21)=2.980, p=0.49), and mean power (19.1%; F(1,21)=3.454, p=0.36) during squats at the velocity of 10 deg/s. Resistance exercises under stable conditions performed in a 90° tilted environment are more effective in the improvement of high velocity muscle power than their use in combination with balance exercises. Such training may be applicable in pre- and in-flight exercise regimens for astronauts and in functional rehabilitation of bed-ridden patients. © Georg Thieme Verlag KG Stuttgart · New York.

Handling Quality Requirements for Advanced Aircraft Design: Longitudinal Mode

DTIC Science & Technology

1979-08-01

phases of air -to- air combat, for example). This is far simpler than the general problem of control law definition. How- ever, the results of such...unlimited. Ali FORCE FUGHT DYNAMICS LABORATORYAIR FORCE WRIGHT AERONAUTICALLABORATORIES AIR FORCE SYSTEMS COMMANDI * WRIGHT-PATITERSON AIR FORCE BASE...not necessarily shared by the Air Force. Brian. W. VauVliet Project Engineer S Rorad0. Anderson, Chief Control Dynamics Branch Flight Control Division

Simulation of nap-of-the-Earth flight in helicopters

NASA Technical Reports Server (NTRS)

Condon, Gregory W.

1991-01-01

NASA-Ames along with the U.S. Army has conducted extensive simulation studies of rotorcraft in the nap-of-the-Earth (NOE) environment and has developed facility capabilities specifically designed for this flight regime. The experience gained to date in applying these facilities to the NOE flight regime are reported along with the results of specific experimental studies conducted to understand the influence of both motion and visual scene on the fidelity of NOE simulation. Included are comparisons of results from concurrent piloted simulation and flight research studies. The results of a recent simulation experiment to study simulator sickness in this flight regime is also discussed.

Ducted fan inlet/exit and rotor tip flow improvements for vertical lift systems

NASA Astrophysics Data System (ADS)

Akturk, Ali

The current research utilized experimental and computational techniques in 5" and 22" diameter ducted fan test systems that have been custom designed and manufactured. Qualitative investigation of flow around the ducted fan was also performed using smoke flow visualizations. Quantitative measurements consisted of 2D and 3D velocity measurements using planar and Stereoscopic Particle Image Velocimetry (PIV and SPIV), high resolution total pressure measurements using Kiel total pressure probes and real time six-component force and torque measurements. The computational techniques used in this thesis included a recently developed radial equilibrium based rotor model(REBRM) and a three dimensional Reynolds-Averaged Navier Stokes (RANS) based CFD model. A radial equilibrium based rotor model (REBRM) developed by the author was effectively integrated into a three-dimensional RANS based computational system. The PIV measurements and computational flow predictions using (REBRM) near the fan inlet plane were in a good agreement at hover and forward flight conditions. The aerodynamic modifications resulting from the fan inlet flow distortions in forward flight regime were clearly captured in 2D PIV results. High resolution total pressure measurements at the downstream of the fan rotor showed that tip leakage, rotor hub separation, and passage flow related total pressure losses were dominant in hover condition. However, the losses were dramatically increased in forward flight because of inlet lip separation and distortion. A novel ducted fan inlet flow conditioning concept named "Double Ducted Fan" (DDF) was developed. The (DDF) concept has a potential to significantly improve the performance and controllability of VTOL UAVs and many other ducted fan based vertical lift systems. The new concept that will significantly reduce the inlet lip separation related performance penalties used a secondary stationary duct system to control "inlet lip separation" occurring especially at elevated forward flight velocities. The (DDF) is self-adjusting in a wide forward flight velocity range. DDFs corrective aerodynamic in influence becomes more pronounced with increasing flight velocity due to its inherent design properties. RANS simulations of the flow around rotor blades and duct geometry in the rotating frame of reference provided a comprehensive description of the tip leakage and passage flow in the flow environment of the two ducted fan research facilities developed throughout this thesis. The aerodynamic measurements and results of the RANS simulation showed good agreement especially near the tip region. A number of novel tip treatments based on custom designed pressure side extensions were introduced. Various tip leakage mitigation schemes were introduced by varying the chordwise location and the width of the extension in the circumferential direction. The current study showed that a proper selection of the pressure side bump location and width were the two critical parameters in influencing the success of the tip leakage mitigation approach. Significant gains in axial mean velocity component were observed when a proper pressure side tip extension was used. It is also observed that an effective tip leakage mitigation scheme significantly reduced the tangential velocity component near the tip of the axial fan blade. Reduced tip clearance related flow interactions were essential in improving the energy efficiency and range of ducted fan based vehicle. Full and inclined pressure side tip squealers were designed. Squealer tips were effective in changing the overall trajectory of the tip vortex to a higher path in radial direction. The interaction of rotor blades and tip vortex was effectively reduced and aerodynamic performance of the rotor blades was improved. The overall aerodynamic gain was a measurable reduction in leakage mass flow rate. This leakage reduction increased thrust significantly. Full and inclined pressure side tip squealers increased thrust obtained in hover condition by 9.1 % and 9.6 % respectively. A reduction or elimination of the momentum deficit in tip vortices is essential to reduce the adverse performance effects originating from the unsteady and highly turbulent tip leakage flows rotating against a stationary casing. The novel tip treatments developed throughout this thesis research are highly effective in reducing the adverse performance effects of ducted fan systems developed for VTOL vehicles. (Abstract shortened by UMI.)

Temperature control of the Mariner class spacecraft - A seven mission summary.

NASA Technical Reports Server (NTRS)

Dumas, L. N.

1973-01-01

Mariner spacecraft have completed five missions of scientific investigation of the planets. Two additional missions are planned. A description of the thermal design of these seven spacecraft is given herein. The factors which have influenced the thermal design include the mission requirements and constraints, the flight environment, certain programmatic considerations and the experience gained as each mission is completed. These factors are reviewed and the impact of each on thermal design and developmental techniques is assessed. It is concluded that the flight success of these spacecraft indicates that adequate temperature control has been obtained, but that improvements in design data, hardware performance and analytical techniques are needed.

High-temperature acoustic test facilities and methods

NASA Astrophysics Data System (ADS)

Pearson, Jerome

1994-09-01

The Wright Laboratory is the Air Force center for air vehicles, responsible for developing advanced technology and incorporating it into new flight vehicles and for continuous technological improvement of operational air vehicles. Part of that responsibility is the problem of acoustic fatigue. With the advent of jet aircraft in the 1950's, acoustic fatigue of aircraft structure became a significant problem. In the 1960's the Wright Laboratory constructed the first large acoustic fatigue test facilities in the United States, and the laboratory has been a dominant factor in high-intensity acoustic testing since that time. This paper discusses some of the intense environments encountered by new and planned Air Force flight vehicles, and describes three new acoustic test facilities of the Wright Laboratory designed for testing structures in these dynamic environments. These new test facilities represent the state of the art in high-temperature, high-intensity acoustic testing and random fatigue testing. They will allow the laboratory scientists and engineers to test the new structures and materials required to withstand the severe environments of captive-carry missiles, augmented lift wings and flaps, exhaust structures of stealth aircraft, and hypersonic vehicle structures well into the twenty-first century.

X-36 Tailless Fighter Agility Research Aircraft in flight

NASA Technical Reports Server (NTRS)

1997-01-01

The lack of a vertical tail on the X-36 technology demonstrator is evident as the remotely piloted aircraft flies a low-altitude research flight above Rogers Dry Lake at Edwards Air Force Base in the California desert on October 30, 1997. The NASA/Boeing X-36 Tailless Fighter Agility Research Aircraft program successfully demonstrated the tailless fighter design using advanced technologies to improve the maneuverability and survivability of possible future fighter aircraft. The program met or exceeded all project goals. For 31 flights during 1997 at the Dryden Flight Research Center, Edwards, California, the project team examined the aircraft's agility at low speed / high angles of attack and at high speed / low angles of attack. The aircraft's speed envelope reached up to 206 knots (234 mph). This aircraft was very stable and maneuverable. It handled very well. The X-36 vehicle was designed to fly without the traditional tail surfaces common on most aircraft. Instead, a canard forward of the wing was used as well as split ailerons and an advanced thrust-vectoring nozzle for directional control. The X-36 was unstable in both pitch and yaw axes, so an advanced, single-channel digital fly-by-wire control system (developed with some commercially available components) was put in place to stabilize the aircraft. Using a video camera mounted in the nose of the aircraft and an onboard microphone, the X-36 was remotely controlled by a pilot in a ground station virtual cockpit. A standard fighter-type head-up display (HUD) and a moving-map representation of the vehicle's position within the range in which it flew provided excellent situational awareness for the pilot. This pilot-in-the-loop approach eliminated the need for expensive and complex autonomous flight control systems and the risks associated with their inability to deal with unknown or unforeseen phenomena in flight. Fully fueled the X-36 prototype weighed approximately 1,250 pounds. It was 19 feet long and three feet high with a wingspan of just over 10 feet. A Williams International F112 turbofan engine provided close to 700 pounds of thrust. A typical research flight lasted 35 to 45 minutes from takeoff to touchdown. A total of 31 successful research flights were flown from May 17, 1997, to November 12, 1997, amassing 15 hours and 38 minutes of flight time. The aircraft reached an altitude of 20,200 feet and a maximum angle of attack of 40 degrees. In a follow-on effort, the Air Force Research Laboratory (AFRL), Wright-Patterson Air Force Base, Ohio, contracted with Boeing to fly AFRL's Reconfigurable Control for Tailless Fighter Aircraft (RESTORE) software as a demonstration of the adaptability of the neural-net algorithm to compensate for in-flight damage or malfunction of effectors, such as flaps, ailerons and rudders. Two RESTORE research flights were flown in December 1998, proving the viability of the software approach. The X-36 aircraft flown at the Dryden Flight Research Center in 1997 was a 28-percent scale representation of a theoretical advanced fighter aircraft. The Boeing Phantom Works (formerly McDonnell Douglas) in St. Louis, Missouri, built two of the vehicles in a cooperative agreement with the Ames Research Center, Moffett Field, California.

Assessment of Comprehensive Analysis Calculation of Airloads on Helicopter Rotors

NASA Technical Reports Server (NTRS)

Yeo, Hyeonsoo; Johnson, Wayne

2004-01-01

Blade section normal force and pitching moment were investigated for six rotors operating at transition and high speeds: H-34 in flight and wind tunnel, SA 330 (research Puma), SA 349/2, UH-60A full-scale and BO-105 model (HART-I). The measured data from flight and wind tunnel tests were compared with calculations obtained using the comprehensive analysis CAMRAD II. The calculations were made using two free wake models: rolled-up and multiple-trailer with consolidation models. At transition speed, there is fair to good agreement for the blade section normal force between the test data and analysis for the H-34, research Puma, and SA 349/2 with the rolled-up wake. The calculated airloads differ significantly from the measurements for the UH-60A and BO-105. Better correlation is obtained for the UH-60A and BO-105 by using the multiple-trailer with consolidation wake model. In the high speed condition, the analysis shows generally good agreement with the research Puma flight data in both magnitude and phase. However, poor agreement is obtained for the other rotors examined. The analysis shows that the aerodynamic tip design (chord length and quarter chord location) of the Puma has an important influence on the phase correlation.

DAST in Flight just after Structural Failure of Right Wing

NASA Technical Reports Server (NTRS)

1980-01-01

Two BQM-34 Firebee II drones were modified with supercritical airfoils, called the Aeroelastic Research Wing (ARW), for the Drones for Aerodynamic and Structural Testing (DAST) program, which ran from 1977 to 1983. This photo, taken 12 June 1980, shows the DAST-1 (Serial #72-1557) immediately after it lost its right wing after suffering severe wing flutter. The vehicle crashed near Cuddeback Dry Lake. The Firebee II was selected for the DAST program because its standard wing could be removed and replaced by a supercritical wing. The project's digital flutter suppression system was intended to allow lighter wing structures, which would translate into better fuel economy for airliners. Because the DAST vehicles were flown intentionally at speeds and altitudes that would cause flutter, the program anticipated that crashes might occur. These are the image contact sheets for each image resolution of the NASA Dryden Drones for Aerodynamic and Structural Testing (DAST) Photo Gallery. From 1977 to 1983, the Dryden Flight Research Center, Edwards, California, (under two different names) conducted the DAST Program as a high-risk flight experiment using a ground-controlled, pilotless aircraft. Described by NASA engineers as a 'wind tunnel in the sky,' the DAST was a specially modified Teledyne-Ryan BQM-34E/F Firebee II supersonic target drone that was flown to validate theoretical predictions under actual flight conditions in a joint project with the Langley Research Center, Hampton, Virginia. The DAST Program merged advances in electronic remote control systems with advances in airplane design. Drones (remotely controlled, missile-like vehicles initially developed to serve as gunnery targets) had been deployed successfully during the Vietnamese conflict as reconnaissance aircraft. After the war, the energy crisis of the 1970s led NASA to seek new ways to cut fuel use and improve airplane efficiency. The DAST Program's drones provided an economical, fuel-conscious method for conducting in-flight experiments from a remote ground site. DAST explored the technology required to build wing structures with less than normal stiffness. This was done because stiffness requires structural weight but ensures freedom from flutter-an uncontrolled, divergent oscillation of the structure, driven by aerodynamic forces and resulting in structural failure. The program used refined theoretical tools to predict at what speed flutter would occur. It then designed a high-response control system to counteract the motion and permit a much lighter wing structure. The wing had, in effect, 'electronic stiffness.' Flight research with this concept was extremely hazardous because an error in either the flutter prediction or control system implementation would result in wing structural failure and the loss of the vehicle. Because of this, flight demonstration of a sub-scale vehicle made sense from the standpoint of both safety and cost. The program anticipated structural failure during the course of the flight research. The Firebee II was a supersonic drone selected as the DAST testbed because its wing could be easily replaced, it used only tail-mounted control surfaces, and it was available as surplus from the U. S. Air Force. It was capable of 5-g turns (that is, turns producing acceleration equal to 5 times that of gravity). Langley outfitted a drone with an aeroelastic, supercritical research wing suitable for a Mach 0.98 cruise transport with a predicted flutter speed of Mach 0.95 at an altitude of 25,000 feet. Dryden and Langley, in conjunction with Boeing, designed and fabricated a digital flutter suppression system (FSS). Dryden developed an RPRV (remotely piloted research vehicle) flight control system; integrated the wing, FSS, and vehicle systems; and conducted the flight program. In addition to a digital flight control system and aeroelastic wings, each DAST drone had research equipment mounted in its nose and a mid-air retrieval system in its tail. The drones were originally launched from the NASA B-52 bomber and later from a DC-130. The DAST vehicle's flight was monitored from the sky by an F-104 chase plane. When the DAST's mission ended, it deployed a parachute and then a specially equipped Air Force helicopter recovered the drone in mid-air. On the ground, a pilot controlled the DAST vehicle from a remote cockpit while researchers in another room monitored flight data transmitted via telemetry. They made decisions on the conduct of the flight while the DAST was in the air. In case of failure in any of the ground systems, the DAST vehicle could also be flown to a recovery site using a backup control system in the F-104. The DAST Program experienced numerous problems. Only eighteen flights were achieved, eight of them captive (in which the aircraft flew only while still attached to the launch aircraft). Four of the flights were aborted and two resulted in crashes--one on June 12, 1980, and the second on June 1, 1983. Meanwhile, flight experiments with higher profiles, better funded remotely piloted research vehicles took priority over DAST missions. After the 1983 crash, which was caused by a malfunction that disconnected the landing parachute from the drone, the program was disbanded. Because DAST drones were considered expendable, certain losses were anticipated. Managers and researchers involved in other high-risk flight projects gained insights from the DAST program that could be applied to their own flight research programs. The DAST aircraft had a wingspan of 14 feet, four inches and a nose-to-tail length of 28 feet, 4 inches. The fuselage had a radius of about 2.07 feet. The aircraft's maximum loaded weight was about 2,200 pounds. It derived its power from a Continental YJ69-T-406 engine.

VANDENBERG AIR FORCE BASE, CALIF. - The L-1011 carrier aircraft is in flight with its cargo underneath of the Pegasus launch vehicle and SciSat-1 spacecraft. The SciSat-1 weighs approximately 330 pounds and after launch will be placed in a 400-mile-high polar orbit to investigate processes that control the distribution of ozone in the upper atmosphere. The data from the satellite will provide Canadian and international scientists with improved measurements relating to global ozone processes and help policymakers assess existing environmental policy and develop protective measures for improving the health of our atmosphere, preventing further ozone depletion. The mission is designed to last two years.

NASA Image and Video Library

2003-08-12

VANDENBERG AIR FORCE BASE, CALIF. - The L-1011 carrier aircraft is in flight with its cargo underneath of the Pegasus launch vehicle and SciSat-1 spacecraft. The SciSat-1 weighs approximately 330 pounds and after launch will be placed in a 400-mile-high polar orbit to investigate processes that control the distribution of ozone in the upper atmosphere. The data from the satellite will provide Canadian and international scientists with improved measurements relating to global ozone processes and help policymakers assess existing environmental policy and develop protective measures for improving the health of our atmosphere, preventing further ozone depletion. The mission is designed to last two years.

VANDENBERG AIR FORCE BASE, CALIF. - The L-1011 carrier aircraft is in flight with its cargo of the Pegasus launch vehicle and SciSat-1 spacecraft underneath. The SciSat-1 weighs approximately 330 pounds and after launch will be placed in a 400-mile-high polar orbit to investigate processes that control the distribution of ozone in the upper atmosphere. The data from the satellite will provide Canadian and international scientists with improved measurements relating to global ozone processes and help policymakers assess existing environmental policy and develop protective measures for improving the health of our atmosphere, preventing further ozone depletion. The mission is designed to last two years.

NASA Image and Video Library

2003-08-12

VANDENBERG AIR FORCE BASE, CALIF. - The L-1011 carrier aircraft is in flight with its cargo of the Pegasus launch vehicle and SciSat-1 spacecraft underneath. The SciSat-1 weighs approximately 330 pounds and after launch will be placed in a 400-mile-high polar orbit to investigate processes that control the distribution of ozone in the upper atmosphere. The data from the satellite will provide Canadian and international scientists with improved measurements relating to global ozone processes and help policymakers assess existing environmental policy and develop protective measures for improving the health of our atmosphere, preventing further ozone depletion. The mission is designed to last two years.

Comparative Analysis of the Tour Jete and Aerial with Detailed Analysis of Aerial Takeoff Mechanics

NASA Astrophysics Data System (ADS)

Pierson, Mimi; Coplin, Kim

2006-10-01

Whether internally as muscle tension or from external sources, forces are necessary for all motion. This research focused on athletic rotations where conditions of flight are established during takeoff. By studying reaction forces that produce torques, moments of inertia, and linear and angular differences between distinct rotations around different principle axes of the body (tour jete in ballet - longitudinal axis; aerial in gymnastics - anteroposterior axis), and by looking at the values of angular momentum in the specific mechanics of aerial takeoff, we can gain insight into possible causes of injury, flaws in technique and limitations of athletes. Results showed significant differences in the horizontal and vertical components of takeoff between the tour jete and the aerial, and a realization that torque was produced in different biomechanical planes. Both rotations showed braking forces before takeoff to counteract forward momentum and increase vertical lift, but the angle of applied force varied, and the horizontal components of velocity and force and vertical velocity as well as moment of inertia throughout flight were consistently greater for the aerial. Breakdown of aerial takeoff highlighted the relative importance of the takeoff phases, showing that completion depends fundamentally upon the rotation of the rear foot and torso twisting during takeoff rather than the last foot in contact with the ground.

Shuttle's 160 hour ground turnaround - A design driver

NASA Technical Reports Server (NTRS)

Widick, F.

1977-01-01

Turnaround analysis added a new dimension to the Space Program with the advent of the Space Shuttle. The requirement to turn the flight hardware around in 160 working hours from landing to launch was a significant design driver and a useful tool in forcing the integration of flight and ground systems design to permit an efficient ground operation. Although there was concern that time constraints might increase program costs, the result of the analysis was to minimize facility requirements and simplify operations with resultant cost savings.

Eclipse program F-106 aircraft in flight, front view

NASA Technical Reports Server (NTRS)

1997-01-01

Shot of the QF-106 aircraft in flight with the landing gear deployed. In 1997 and 1998, the Dryden Flight Research Center at Edwards, California, supported and hosted a Kelly Space & Technology, Inc. project called Eclipse, which sought to demonstrate the feasibility of a reusable tow-launch vehicle concept. The project goal was to successfully tow, inflight, a modified QF-106 delta-wing aircraft with an Air Force C-141A transport aircraft. This would demonstrate the possibility of towing and launching an actual launch vehicle from behind a tow plane. Dryden was the responsible test organization and had flight safety responsibility for the Eclipse project. Dryden provided engineering, instrumentation, simulation, modification, maintenance, range support, and research pilots for the test program. The Air Force Flight Test Center (AFFTC), Edwards, California, supplied the C-141A transport aircraft and crew and configured the aircraft as needed for the tests. The AFFTC also provided the concept and detail design and analysis as well as hardware for the tow system and QF-106 modifications. Dryden performed the modifications to convert the QF-106 drone into the piloted EXD-01 (Eclipse eXperimental Demonstrator-01) experimental aircraft. Kelly Space & Technology hoped to use the results gleaned from the tow test in developing a series of low-cost, reusable launch vehicles. These tests demonstrated the validity of towing a delta-wing aircraft having high wing loading, validated the tow simulation model, and demonstrated various operational procedures, such as ground processing of in-flight maneuvers and emergency abort scenarios.

The MANTA: An RPV design to investigate forces and moments on a lifting surface

NASA Technical Reports Server (NTRS)

Bryan, Kevin; Soutar, John; Witty, Peter; Mediate, Bruno; Quast, Thomas; Combs, Dan; Schubert, Martin; Condron, David; Taylor, Scott; Garino, ED

1989-01-01

The overall goal was to investigate and exploit the advantages of using remotely powered vehicles (RPV's) for in-flight data collection at low Reynold's numbers. The data to be collected is on actual flight loads for any type of rectangular or tapered airfoil section, including vertical and horizontal stabilizers. The data will be on a test specimen using a force-balance system which is located forward of the aircraft to insure an undisturbed air flow over the test section. The collected data of the lift, drag and moment of the test specimen is to be radioed to a grand receiver, thus providing real-time data acquisition. The design of the mission profile and the selection of the instrumentation to satisfy aerodynamic requirements are studied and tested. A half-size demonstrator was constructed and flown to test the flight worthiness of the system.

Active member vibration control experiment in a KC-135 reduced gravity environment

NASA Technical Reports Server (NTRS)

Lawrence, C. R.; Lurie, B. J.; Chen, G.-S.; Swanson, A. D.

1991-01-01

An active member vibration control experiment in a KC-135 reduced gravity environment was carried out by the Air Force Flight Dynamics Laboratory and the Jet Propulsion Laboratory. Two active members, consisting of piezoelectric actuators, displacement sensors, and load cells, were incorporated into a 12-meter, 104 kg box-type test structure. The active member control design involved the use of bridge (compound) feedback concept, in which the collocated force and velocity signals are feedback locally. An impact-type test was designed to accommodate the extremely short duration of the reduced gravity testing window in each parabolic flight. The moving block analysis technique was used to estimate the modal frequencies and dampings from the free-decay responses. A broadband damping performance was demonstrated up to the ninth mode of 40 Hz. The best damping performance achieved in the flight test was about 5 percent in the fourth mode of the test structure.

Design and preparation of a particle dynamics space flight experiment, SHIVA.

PubMed

Trolinger, James D; L'Esperance, Drew; Rangel, Roger H; Coimbra, Carlos F M; Witherow, William K

2004-11-01

This paper describes the flight experiment, supporting ground science, and the design rationale for a project on spaceflight holography investigation in a virtual apparatus (SHIVA). SHIVA is a fundamental study of particle dynamics in fluids in microgravity. Gravitation effects and steady Stokes drag often dominate the equations of motion of a particle in a fluid and consequently microgravity provides an ideal environment in which to study the other forces, such as the pressure and viscous drag and especially the Basset history force. We have developed diagnostic recording methods using holography to save all of the particle field optical characteristics, essentially allowing the experiment to be transferred from space back to Earth in what we call the "virtual apparatus" for microgravity experiments on Earth. We can quantify precisely the three-dimensional motion of sets of particles, allowing us to test and apply new analytic solutions developed by members of the team. In addition to employing microgravity to augment the fundamental study of these forces, the resulting data will allow us to quantify and understand the ISS environment with great accuracy. This paper shows how we used both experiment and theory to identify and resolve critical issues and to produce an optimal experimental design that exploits microgravity for the study. We examined the response of particles of specific gravity from 0.1 to 20, with radii from 0.2 to 2 mm, to fluid oscillation at frequencies up to 80 Hz with amplitudes up to 200 microns. To observe some of the interesting effects predicted by the new solutions requires the precise location of the position of a particle in three dimensions. To this end we have developed digital holography algorithms that enable particle position location to a small fraction of a pixel in a CCD array. The spaceflight system will record holograms both on film and electronically. The electronic holograms can be downlinked providing real-time data, essentially acting like a remote window into the ISS experimental chamber. Ground experiments have provided input to a flight system design that can meet the requirements for a successful experiment on ISS. Moreover the ground experiments have provided a definitive, quantitative observation of the Basset history force over a wide range of conditions. Results of the ground experiments, the flight experiment design, preliminary flight hardware design, and data analysis procedures are reported.

Space shuttle guidance, navigation, and control design equations. Volume 3: Guidance

NASA Technical Reports Server (NTRS)

1973-01-01

Space shuttle guidance, navigation, and control design equations are presented. The space-shuttle mission includes three relatively distinct guidance phases which are discussed; atmospheric boost, which is characterized by an adaptive guidance law; extra-atmospheric activities; and re-entry activities, where aerodynamic surfaces are the principal effectors. Guidance tasks include pre-maneuver targeting and powered flight guidance, where powered flight is defined to include the application of aerodynamic forces as well as thruster forces. A flow chart which follows guidance activities throughout the mission from the pre-launch phase through touchdown is presented. The main guidance programs and subroutines used in each phase of a typical rendezvous mission are listed. Detailed software requirements are also presented.

Changes in Jump-Down Performance After Space Flight: Short- and Long-Term Adaptation

NASA Technical Reports Server (NTRS)

Kofman, I. S.; Reschke, M. F.; Cerisano, J. M.; Fisher, E. A.; Lawrence, E. L.; Peters, B. T.; Bloomberg, J. J.

2010-01-01

INTRODUCTION Successful jump performance requires functional coordination of visual, vestibular, and somatosensory systems, which are affected by prolonged exposure to microgravity. Astronauts returning from space flight exhibit impaired ability to coordinate effective landing strategies when jumping from a platform to the ground. This study compares the jump strategies used by astronauts before and after flight, the changes to those strategies within a test session, and the recoveries in jump-down performance parameters across several postflight test sessions. These data were obtained as part of an ongoing interdisciplinary study (Functional Task Test, FTT) designed to evaluate both astronaut postflight functional performance and related physiological changes. METHODS Six astronauts from short-duration (Shuttle) and three from long-duration (International Space Station) flights performed 3 two-footed jumps from a platform 30 cm high. A force plate measured the ground reaction forces and center-of-pressure displacement from the landings. Muscle activation data were collected from the medial gastrocnemius and anterior tibialis of both legs using surface electromyography electrodes. Two load cells in the platform measured the load exerted by each foot during the takeoff phase of the jump. Data were collected in 2 preflight sessions, on landing day (Shuttle only), and 1, 6, and 30 days after flight. RESULTS AND CONCLUSION Many of the astronauts tested were unable to maintain balance on their first postflight jump landing but recovered by the third jump, showing a learning progression in which the performance improvement could be attributed to adjustments of strategy on takeoff, landing, or both. Takeoff strategy changes were evident in air time (time between takeoff and landing), which was significantly reduced after flight, and also in increased asymmetry in foot latencies on takeoff. Landing modifications were seen in changes in ground reaction force curves. The results demonstrate astronauts adaptive capabilities and full performance recovery within days after flight.

An Assessment of CFD Effectiveness for Vortex Flow Simulation to Meet Preliminary Design Needs

NASA Technical Reports Server (NTRS)

Raj, P.; Ghaffari, F.; Finley, D. B.

2003-01-01

The low-speed flight and transonic maneuvering characteristics of combat air vehicles designed for efficient supersonic flight are significantly affected by the presence of free vortices. At moderate-to-high angles of attack, the flow invariably separates from the leading edges of the swept slender wings, as well as from the forebodies of the air vehicles, and rolls up to form free vortices. The design of military vehicles is heavily driven by the need to simultaneously improve performance and affordability.1 In order to meet this need, increasing emphasis is being placed on using Modeling & Simulation environments employing the Integrated Product & Process Development (IPPD) concept. The primary focus is on expeditiously providing design teams with high-fidelity data needed to make more informed decisions in the preliminary design stage. Extensive aerodynamic data are needed to support combat air vehicle design. Force and moment data are used to evaluate performance and handling qualities; surface pressures provide inputs for structural design; and flow-field data facilitate system integration. Continuing advances in computational fluid dynamics (CFD) provide an attractive means of generating the desired data in a manner that is responsive to the needs of the preliminary design efforts. The responsiveness is readily characterized as timely delivery of quality data at low cost.

NACA Aircraft in hangar 1953 - clockwise from front center: YF-84A, D-558-1, D-558-2, B-47, X-1 ship

NASA Technical Reports Server (NTRS)

1953-01-01

In the center foreground of this 1953 hanger photo is the YF-84A (NACA 134/Air Force 45-59490) used for vortex generator research. It arrived on November 28, 1949, and departed on April 21, 1954. Beside it is the third D-558-1 aircraft (NACA 142/Navy 37972). This aircraft was used for a total of 78 transonic research flights from April 1949 to June 1954. It replaced the second D-558-1, lost in the crash which killed Howard Lilly. Just visible on the left edge is the nose of the first D-558-2 (NACA 143/Navy 37973). Douglas turned the aircraft over to NACA on August 31, 1951, after the contractor had completed its initial test flights. NACA only made a single flight with the aircraft, on September 17, 1956, before the program was cancelled. In the center of the photo is the B-47A (NACA 150/Air Force 49-1900). The B-47 jet bomber, with its thin, swept-back wings, and six podded engines, represented the state of the art in aircraft design in the early 1950s. The aircraft undertook a number of research activities between May 1953 and its 78th and final research flight on November 22, 1957. The tests showed that the aircraft had a buffeting problem at speeds above Mach 0.8. Among the pilots who flew the B-47 were later X-15 pilots Joe Walker, A. Scott Crossfield, John B. McKay, and Neil A. Armstrong. On the right side of the B-47 is NACA's X-1 (Air Force 46-063). The second XS-1 aircraft built, it was fitted with a thicker wing than that on the first aircraft, which had exceeded Mach 1 on October 14, 1947. Flight research by NACA pilots indicated that this thicker wing produced 30 percent more drag at transonic speeds compared to the thinner wing on the first X-1. After a final flight on October 23, 1951, the aircraft was grounded due to the possibility of fatigue failure of the nitrogen spheres used to pressurize the fuel tanks. At the time of this photo, in 1953, the aircraft was in storage. In 1955, the aircraft was extensively modified, becoming the X-1E. In front of the X-1 is the XF-92A (Air Force 46-682). Unlike the X-1 and D-558 aircraft, the XF-92A was not designed as a research aircraft, but as the prototype of a delta-wing fighter. While the effort was unsuccessful, the XF-92A offered the chance to test a delta wing aircraft. A brief series of 25 flights were made using the aircraft in 1953. These showed the aircraft had violent pitch-up tendencies during turns. Despite the problems, the XF-92A contributed to later delta wing aircraft, like the F-102, F-106, and B-58. Behind the B-47, in the back of the hangar, are four other aircraft. From left to right, they are the second X-4 (Air Force 46-677) research aircraft. It was operated by the NACA from May 8, 1950, to March 22, 1954, when it left the High-Speed Flight Research Station for the U.S. Air Force Museum. It was designed to test the use of swept wings but no horizontal stabilizers. This proved to have poor transonic stability. Next to it is the ETF-51D Mustang (NACA 148/Air Force 44-84958) used for low-speed chase missions, as well as support and liaison flights. On the right side of the B-47 is the first D-558-1. Originally given the Navy number 37970, it was flown as part of the Douglas contractor program. When this was completed, the aircraft was turned over to the NACA on April 11, 1949. Although the aircraft was designation 'NACA 140,' it was never flown again. Instead, it was used to provide spare parts to keep the third D-558-1 in operation. In this photo the aircraft is partially disassembled. The final aircraft is the first X-5 (Air Force 50-1838). This was a research aircraft used to test the concept of pivoting wings which could change their sweep angle in flight. The results were mixed; the X-5 had vicious stall behavior due to the poor position of the tail and stabilizers. The mechanism used by the 'variable-sweep wing' was also complex, which limited its usefulness. Despite these problems, the X-5's primary advantage was that it was equivalent to a whole family of research aircraft. It could provide transonic data at sweep angles up to 60 degrees--the same as the delta wing XF-92A. The Dryden Flight Research Center, NASA's premier installation for aeronautical flight research, celebrated its 50th anniversary in 1996. Dryden is the 'Center of Excellence' for atmospheric flight operations. The Center's charter is to research, develop, verify, and transfer advanced aeronautics, space, and related technologies. It is located at Edwards, Calif., on the western edge of the Mojave Desert, 80 miles north of Los Angeles. Dryden's history dates back to the early fall of 1946, when a group of five aeronautical engineers arrived at what is now Edwards from the NACA's Langley Memorial Aeronautical Laboratory, Hampton, Va. Their goal was to prepare for the X-l supersonic research flights in a joint NACA-U.S. Army Air Forces-Bell Aircraft Corp. program. NACA--the National Advisory Committee for Aeronautics--was the predecessor of today's NASA. Since the days of the X-l, the first aircraft to fly faster than the speed of sound, the installation has grown in size and significance and is associated with many important developments in aviation -- supersonic and hypersonic flight, wingless lifting bodies, digital fly-by-wire, supercritical and forward-swept wings, and the space shuttles. Its name has changed many times over the years. From 14 November 1949 to 1 July 1954 it bore the name NACA High-Speed Flight Research Station.

A variable-gain output feedback control design methodology

NASA Technical Reports Server (NTRS)

Halyo, Nesim; Moerder, Daniel D.; Broussard, John R.; Taylor, Deborah B.

1989-01-01

A digital control system design technique is developed in which the control system gain matrix varies with the plant operating point parameters. The design technique is obtained by formulating the problem as an optimal stochastic output feedback control law with variable gains. This approach provides a control theory framework within which the operating range of a control law can be significantly extended. Furthermore, the approach avoids the major shortcomings of the conventional gain-scheduling techniques. The optimal variable gain output feedback control problem is solved by embedding the Multi-Configuration Control (MCC) problem, previously solved at ICS. An algorithm to compute the optimal variable gain output feedback control gain matrices is developed. The algorithm is a modified version of the MCC algorithm improved so as to handle the large dimensionality which arises particularly in variable-gain control problems. The design methodology developed is applied to a reconfigurable aircraft control problem. A variable-gain output feedback control problem was formulated to design a flight control law for an AFTI F-16 aircraft which can automatically reconfigure its control strategy to accommodate failures in the horizontal tail control surface. Simulations of the closed-loop reconfigurable system show that the approach produces a control design which can accommodate such failures with relative ease. The technique can be applied to many other problems including sensor failure accommodation, mode switching control laws and super agility.

Biomechanics and biomimetics in insect-inspired flight systems

PubMed Central

Liu, Hao; Ravi, Sridhar; Kolomenskiy, Dmitry; Tanaka, Hiroto

2016-01-01

Insect- and bird-size drones—micro air vehicles (MAV) that can perform autonomous flight in natural and man-made environments are now an active and well-integrated research area. MAVs normally operate at a low speed in a Reynolds number regime of 104–105 or lower, in which most flying animals of insects, birds and bats fly, and encounter unconventional challenges in generating sufficient aerodynamic forces to stay airborne and in controlling flight autonomy to achieve complex manoeuvres. Flying insects that power and control flight by flapping wings are capable of sophisticated aerodynamic force production and precise, agile manoeuvring, through an integrated system consisting of wings to generate aerodynamic force, muscles to move the wings and a control system to modulate power output from the muscles. In this article, we give a selective review on the state of the art of biomechanics in bioinspired flight systems in terms of flapping and flexible wing aerodynamics, flight dynamics and stability, passive and active mechanisms in stabilization and control, as well as flapping flight in unsteady environments. We further highlight recent advances in biomimetics of flapping-wing MAVs with a specific focus on insect-inspired wing design and fabrication, as well as sensing systems. This article is part of the themed issue ‘Moving in a moving medium: new perspectives on flight’. PMID:27528780

YF-22 in flight (US AF photo)

NASA Technical Reports Server (NTRS)

1990-01-01

The YF-22, prototype aircraft for the Air Force's F-22 fighter, cruises over the desert on a flight for the Air Force. It was never involved in any programs with Dryden. The United States Air Force announced the demonstration/validation phase contractors selection for the Advanced Tactical Fighter (ATF) program October 31, 1986. These contractor programs were the Lockheed YF-22 and the Northrop YF-23; each produced two prototypes and ground-based avionics testbeds. First flights of all four prototypes occured in 1990. The YF-22 was first flown on Sept. 29, 1990. The YF-22 was powered by two General Electric YF120-GE-100 engines. The final design, the F-22, was flown sometime in May 1997. The F-22 is capable of efficient supersonic operation without afterburner use (supercruise). Lockheed teamed with General Dynamics (Fort Worth) and Boeing Military Airplanes to produce two YF-22 prototypes, civil registrations N22YF (with GE YF120) and N22YX (P&W YF119). N22YF rolled out at Palmdale August 29, 1990; first flight/ferry to Edwards AFB September 29, 1990; first air refuelling (11th sortie) October 26, 1990; thrust vectoring in flight November 15, 1990; achieved Mach 1.8 December 26, 1990. Flight test demonstrations included `supercruise' flight in excess of Mach 1.58 without afterburner.

External force/velocity control for an autonomous rehabilitation robot

NASA Astrophysics Data System (ADS)

Saekow, Peerayuth; Neranon, Paramin; Smithmaitrie, Pruittikorn

2018-01-01

Stroke is a primary cause of death and the leading cause of permanent disability in adults. There are many stroke survivors, who live with a variety of levels of disability and always need rehabilitation activities on daily basis. Several studies have reported that usage of rehabilitation robotic devices shows the better improvement outcomes in upper-limb stroke patients than the conventional therapy-nurses or therapists actively help patients with exercise-based rehabilitation. This research focuses on the development of an autonomous robotic trainer designed to guide a stroke patient through an upper-limb rehabilitation task. The robotic device was designed and developed to automate the reaching exercise as mentioned. The designed robotic system is made up of a four-wheel omni-directional mobile robot, an ATI Gamma multi-axis force/torque sensor used to measure contact force and a microcontroller real-time operating system. Proportional plus Integral control was adapted to control the overall performance and stability of the autonomous assistive robot. External force control was successfully implemented to establish the behavioral control strategy for the robot force and velocity control scheme. In summary, the experimental results indicated satisfactorily stable performance of the robot force and velocity control can be considered acceptable. The gain tuning for proportional integral (PI) velocity control algorithms was suitably estimated using the Ziegler-Nichols method in which the optimized proportional and integral gains are 0.45 and 0.11, respectively. Additionally, the PI external force control gains were experimentally tuned using the trial and error method based on a set of experiments which allow a human participant moves the robot along the constrained circular path whilst attempting to minimize the radial force. The performance was analyzed based on the root mean square error (E_RMS) of the radial forces, in which the lower the variation in radial forces, the better the performance of the system. The outstanding performance of the tests as specified by the E_RMS of the radial force was observed with proportional and integral gains of Kp = 0.7 and Ki = 0.75, respectively.

NASA Dryden's new in-house designed Propulsion Flight Test Fixture (PFTF), carried on an F-15B's cen

NASA Technical Reports Server (NTRS)

2001-01-01

NASA Dryden Flight Research Center's new in-house designed Propulsion Flight Test Fixture (PFTF) is an airborne engine test facility that allows engineers to glean actual flight data on small experimental engines that would otherwise have to be gathered from traditional wind tunnels, ground test stands or laboratory setups. Now, with the 'captive carry' capability of the PFTF, new air-breathing propulsion schemes, such as Rocket Based Combined Cycle engines, can be economically flight-tested using sub-scale experiments. The PFTF flew mated to NASA Dryden's specially-equipped supersonic F-15B research aircraft during December 2001 and January 2002. The PFTF, carried on the F-15B's centerline attachment point, underwent in-flight checkout, known as flight envelope expansion, in order to verify its design and capabilities. Envelope expansion for the PFTF included envelope clearance, which involves maximum performance testing. Top speed of the F-15B with the PFTF is Mach 2.0. Other elements of envelope clearance are flying qualities assessment and flutter analysis. Airflow visualization of the PFTF and a 'stand-in' test engine was accomplished by attaching small tufts of nylon on them and videotaping the flow patterns revealed during flight. A surrogate experimental engine shape, called the cone tube, was flown attached to the force balance on the PFTF. The cone tube emulated the dimensional and mass properties of the maximum design load the PFTF can carry. As the F-15B put the PFTF and the attached cone tube through its paces, accurate data was garnered, allowing engineers to fully verify PFTF and force balance capabilities in real flight conditions. When the first actual experimental engine is ready to fly on the F-15B/PFTF, engineers will have full confidence and knowledge of what they can accomplish with this 'flying engine test stand.'

NASA Dryden's new in-house designed Propulsion Flight Test Fixture (PFTF) flew mated to a specially-

NASA Technical Reports Server (NTRS)

2001-01-01

NASA Dryden Flight Research Center's new in-house designed Propulsion Flight Test Fixture (PFTF) is an airborne engine test facility that allows engineers to glean actual flight data on small experimental engines that would otherwise have to be gathered from traditional wind tunnels, ground test stands or laboratory setups. Now, with the 'captive carry' capability of the PFTF, new air-breathing propulsion schemes, such as Rocket Based Combined Cycle engines, can be economically flight-tested using sub-scale experiments. The PFTF flew mated to NASA Dryden's specially-equipped supersonic F-15B research aircraft during December 2001 and January 2002. The PFTF, carried on the F-15B's centerline attachment point, underwent in-flight checkout, known as flight envelope expansion, in order to verify its design and capabilities. Envelope expansion for the PFTF included envelope clearance, which involves maximum performance testing. Top speed of the F-15B with the PFTF is Mach 2.0. Other elements of envelope clearance are flying qualities assessment and flutter analysis. Airflow visualization of the PFTF and a 'stand-in' test engine was accomplished by attaching small tufts of nylon on them and videotaping the flow patterns revealed during flight. A surrogate experimental engine shape, called the cone tube, was flown attached to the force balance on the PFTF. The cone tube emulated the dimensional and mass properties of the maximum design load the PFTF can carry. As the F-15B put the PFTF and the attached cone tube through its paces, accurate data was garnered, allowing engineers to fully verify PFTF and force balance capabilities in real flight conditions. When the first actual experimental engine is ready to fly on the F-15B/PFTF, engineers will have full confidence and knowledge of what they can accomplish with this 'flying engine test stand.'

Flexible strategies for flight control: an active role for the abdomen.

PubMed

Dyhr, Jonathan P; Morgansen, Kristi A; Daniel, Thomas L; Cowan, Noah J

2013-05-01

Moving animals orchestrate myriad motor systems in response to multimodal sensory inputs. Coordinating movement is particularly challenging in flight control, where animals deal with potential instability and multiple degrees of freedom of movement. Prior studies have focused on wings as the primary flight control structures, for which changes in angle of attack or shape are used to modulate lift and drag forces. However, other actuators that may impact flight performance are reflexively activated during flight. We investigated the visual-abdominal reflex displayed by the hawkmoth Manduca sexta to determine its role in flight control. We measured the open-loop stimulus-response characteristics (measured as a transfer function) between the visual stimulus and abdominal response in tethered moths. The transfer function reveals a 41 ms delay and a high-pass filter behavior with a pass band starting at ~0.5 Hz. We also developed a simplified mathematical model of hovering flight wherein articulation of the thoracic-abdominal joint redirects an average lift force provided by the wings. We show that control of the joint, subject to a high-pass filter, is sufficient to maintain stable hovering, but with a slim stability margin. Our experiments and models suggest a novel mechanism by which articulation of the body or 'airframe' of an animal can be used to redirect lift forces for effective flight control. Furthermore, the small stability margin may increase flight agility by easing the transition from stable flight to a more maneuverable, unstable regime.

Design of an all-attitude flight control system to execute commanded bank angles and angles of attack

NASA Technical Reports Server (NTRS)

Burgin, G. H.; Eggleston, D. M.

1976-01-01

A flight control system for use in air-to-air combat simulation was designed. The input to the flight control system are commanded bank angle and angle of attack, the output are commands to the control surface actuators such that the commanded values will be achieved in near minimum time and sideslip is controlled to remain small. For the longitudinal direction, a conventional linear control system with gains scheduled as a function of dynamic pressure is employed. For the lateral direction, a novel control system, consisting of a linear portion for small bank angle errors and a bang-bang control system for large errors and error rates is employed.

Flight data processing with the F-8 adaptive algorithm

NASA Technical Reports Server (NTRS)

Hartmann, G.; Stein, G.; Petersen, K.

1977-01-01

An explicit adaptive control algorithm based on maximum likelihood estimation of parameters has been designed for NASA's DFBW F-8 aircraft. To avoid iterative calculations, the algorithm uses parallel channels of Kalman filters operating at fixed locations in parameter space. This algorithm has been implemented in NASA/DFRC's Remotely Augmented Vehicle (RAV) facility. Real-time sensor outputs (rate gyro, accelerometer and surface position) are telemetered to a ground computer which sends new gain values to an on-board system. Ground test data and flight records were used to establish design values of noise statistics and to verify the ground-based adaptive software. The software and its performance evaluation based on flight data are described

Emulating avian orographic soaring with a small autonomous glider.

PubMed

Fisher, Alex; Marino, Matthew; Clothier, Reece; Watkins, Simon; Peters, Liam; Palmer, Jennifer L

2015-12-17

This paper explores a method by which an unpowered, fixed-wing micro air vehicle (MAV) may autonomously gain height by utilising orographic updrafts in urban environments. These updrafts are created when wind impinges on both man-made and natural obstacles, and are often highly turbulent and very localised. Thus in contrast to most previous autonomous soaring research, which have focused on large thermals and ridges, we use a technique inspired by kestrels known as 'wind-hovering', in order to maintain unpowered flight within small updrafts. A six-degree-of-freedom model of a MAV was developed based on wind-tunnel tests and vortex-lattice calculations, and the model was used to develop and test a simple cascaded control system designed to hold the aircraft on a predefined trajectory within an updraft. The wind fields around two typical updraft locations (a building and a hill) were analysed, and a simplified trajectory calculation method was developed by which trajectories for height gain can be calculated on-board the aircraft based on a priori knowledge of the wind field. The results of simulations are presented, demonstrating the behaviour of the system in both smooth and turbulent flows. Finally, the results from a series of flight tests are presented. Flight tests at the hill were consistently successful, while flights around the building could not be sustained for periods of more than approximately 20 s. The difficulty of operating near a building is attributable to significant levels of low-frequency unsteadiness (gustiness) in the oncoming wind during the flight tests, effectively resulting in a loss of updraft for sustained periods.

Development and Testing of Control Laws for the Active Aeroelastic Wing Program

NASA Technical Reports Server (NTRS)

Dibley, Ryan P.; Allen, Michael J.; Clarke, Robert; Gera, Joseph; Hodgkinson, John

2005-01-01

The Active Aeroelastic Wing research program was a joint program between the U.S. Air Force Research Laboratory and NASA established to investigate the characteristics of an aeroelastic wing and the technique of using wing twist for roll control. The flight test program employed the use of an F/A-18 aircraft modified by reducing the wing torsional stiffness and adding a custom research flight control system. The research flight control system was optimized to maximize roll rate using only wing surfaces to twist the wing while simultaneously maintaining design load limits, stability margins, and handling qualities. NASA Dryden Flight Research Center developed control laws using the software design tool called CONDUIT, which employs a multi-objective function optimization to tune selected control system design parameters. Modifications were made to the Active Aeroelastic Wing implementation in this new software design tool to incorporate the NASA Dryden Flight Research Center nonlinear F/A-18 simulation for time history analysis. This paper describes the design process, including how the control law requirements were incorporated into constraints for the optimization of this specific software design tool. Predicted performance is also compared to results from flight.

Induced Moment Effects of Formation Flight Using Two F/A-18 Aircraft

NASA Technical Reports Server (NTRS)

Hansen, Jennifer L.; Cobleigh, Brent R.

2002-01-01

Previous investigations into formation flight have shown the possibility for significant fuel savings through drag reduction. Using two F/A-18 aircraft, NASA Dryden Flight Research Center has investigated flying aircraft in autonomous formation. Positioning the trailing airplane for best drag reduction requires investigation of the wingtip vortex effects induced by the leading airplane. A full accounting of the vortex effect on the trailing airplane is desired to validate vortex-effect prediction methods and provide a database for the design of a formation flight autopilot. A recent flight phase has mapped the complete wingtip vortex effects at two flight conditions with the trailing airplane at varying distances behind the leading one. Force and moment data at Mach 0.56 and an altitude of 25,000 ft and Mach 0.86 and an altitude of 36,000 ft have been obtained with 20, 55, 110, and 190 ft of longitudinal distance between the aircraft. The moments induced by the vortex on the trailing airplane were well within the pilot's ability to control. This report discusses the data analysis methods and vortex-induced effects on moments and side force. An assessment of the impact of the nonlinear vortex effects on the design of a formation autopilot is offered.

Design and Test Requirements for Space Flight Pressurized Systems

DTIC Science & Technology

2014-11-26

5.7.2 Fire Resistant Fluids ......................................................................................... 24 5.7.3 Accumulators...5.1.7 Threaded Parts [5.1.7-1] All threaded parts in components shall be securely locked to resist uncoupling forces by acceptable safe design...event of force balance on both inlet ports may occur, thus causing the shutoff valve to restrict flow from the outlet port. 5.7.2 Fire Resistant

Biomechanics and biomimetics in insect-inspired flight systems.

PubMed

Liu, Hao; Ravi, Sridhar; Kolomenskiy, Dmitry; Tanaka, Hiroto

2016-09-26

Insect- and bird-size drones-micro air vehicles (MAV) that can perform autonomous flight in natural and man-made environments are now an active and well-integrated research area. MAVs normally operate at a low speed in a Reynolds number regime of 10(4)-10(5) or lower, in which most flying animals of insects, birds and bats fly, and encounter unconventional challenges in generating sufficient aerodynamic forces to stay airborne and in controlling flight autonomy to achieve complex manoeuvres. Flying insects that power and control flight by flapping wings are capable of sophisticated aerodynamic force production and precise, agile manoeuvring, through an integrated system consisting of wings to generate aerodynamic force, muscles to move the wings and a control system to modulate power output from the muscles. In this article, we give a selective review on the state of the art of biomechanics in bioinspired flight systems in terms of flapping and flexible wing aerodynamics, flight dynamics and stability, passive and active mechanisms in stabilization and control, as well as flapping flight in unsteady environments. We further highlight recent advances in biomimetics of flapping-wing MAVs with a specific focus on insect-inspired wing design and fabrication, as well as sensing systems.This article is part of the themed issue 'Moving in a moving medium: new perspectives on flight'. © 2016 The Author(s).

ED15-0104-78

NASA Image and Video Library

2015-04-09

The X-56A Multi-Utility Technology Testbed (MUTT) is greeted on an Edwards Air Force Base runway by a U.S. Air Force Research Laboratory (AFRL) team member. NASA’s Armstrong Flight Research Center and the AFRL, along with participants from Langley Research Center and Glenn Research Center, and support from Lockheed Martin, are using the second X-56A (dubbed “Buckeye”) to check out aircraft systems, evaluate handling qualities, characterize and expand the airplane’s performance envelope, and verify pre-flight predictions regarding aircraft behavior. The 20-minute flight marked the beginning of a research effort designed to yield significant advances in aeroservoelastic technology using a low-cost, modular, remotely piloted aerial vehicle.

Experimental verification of high energy laser-generated impulse for remote laser control of space debris.

PubMed

Lorbeer, Raoul-Amadeus; Zwilich, Michael; Zabic, Miroslav; Scharring, Stefan; Eisert, Lukas; Wilken, Jascha; Schumacher, Dennis; Roth, Markus; Eckel, Hans-Albert

2018-05-31

Walking along a beach one may notice debris being washed ashore from the vast oceans. Then, turning your head up at night you even might noticed a shooting star or a bright spot passing by. Chances are, that you witnessed space debris, endangering future space flight in lower earth orbit. If it was possible to turn cm-sized debris into shooting stars the problem might be averted. Unfortunately, these fragments counting in the 100 thousands are not controllable. To possibly regain control we demonstrate how to exert forces on a free falling debris object from a distance by ablating material with a high energy ns-laser-system. Thrust effects did scale as expected from simulations and led to speed gains above 0.3 m/s per laser pulse in an evacuated micro-gravity environment.

X-24A in Powered Flight after Drop from B-52 Mothership

NASA Technical Reports Server (NTRS)

1970-01-01

The X-24A lights its XLR-11 rocket engine and begins its powered flight after being drop launched from its B-52 mothership, seen here with high-altitude contrails streaming from its wings against a piercingly dark blue sky. The X-24 was one of a group of lifting bodies flown by the NASA Flight Research Center (now Dryden Flight Research Center), Edwards, California, in a joint program with the U.S. Air Force at Edwards Air Force Base from 1963 to 1975. The lifting bodies were used to demonstrate the ability of pilots to maneuver and safely land wingless vehicles designed to fly back to Earth from space and be landed like an airplane at a predetermined site. Lifting bodies' aerodynamic lift, essential to flight in the atmosphere, was obtained from their shape. The addition of fins and control surfaces allowed the pilots to stabilize and control the vehicles and regulate their flight paths. Built by Martin Aircraft Company, Maryland, for the U.S. Air Force, the X-24A was a bulbous vehicle shaped like a teardrop with three vertical fins at the rear for directional control. It weighed 6,270 pounds, was 24.5 feet long and 11.5 feet wide (measuring just the fuselage, not the distance between the tips of the outboard fins). Its first unpowered glide flight was on April 17, 1969, with Air Force Maj. Jerauld Gentry at the controls. Gentry also piloted its first powered flight on March 19, 1970. The X-24A was flown 28 times in the program that, like the HL-10, validated the concept that a Space Shuttle vehicle could be landed unpowered. The fastest speed achieved by the X-24A was 1,036 miles per hour (mph-Mach 1.6). Its maximum altitude was 71,400 feet. It was powered by an XLR-11 rocket engine with a maximum theoretical vacuum thrust of 8,480 pounds. The X-24A was later modified into the X-24B. The bulbous shape of the X-24A was converted into a 'flying flatiron' shape with a rounded top, flat bottom, and double delta platform that ended in a pointed nose. The X-24B demonstrated that accurate unpowered reentry vehicle landings were operationally feasible. Top speed achieved by the X-24B was 1,164 mph and the highest altitude it reached was 74,130 feet. The vehicle is on display at the Air Force Museum, Wright-Patterson Air Force Base, Ohio. The pilot on the last powered flight of the X-24B was Bill Dana, who also flew the last X-15 flight about seven years earlier. The X-24A shape was later borrowed for the X-38 Crew Return Vehicle (CRV) technology demonstrator for the International Space Station.

Walter C. Williams with Brig. General Albert Boyd

NASA Technical Reports Server (NTRS)

1950-01-01

Walter C. Williams, (behind airplane model) Head of the National Advisory Committee for Aeronautics High-Speed Flight Research Station at Edwards Air Force Base in California is examining a Northrop X-4 research airplane with Brig. Gen. Albert Boyd, Commander of Edwards Air Force Base. At Edwards, the Air Force Air Material Command ran a brief program on the X-4 during the summer of 1950 before delivering it to the NACA. Data was collected on these 14 flights, so they were logged as NACA test flights. General Boyd made flight number 13. Air Force and NACA pilots completed a total of 82 flights on X-4 #2 (46-677) between August 1950 and September 1953. There are three things that made the Mojave Desert, where Edwards Air Force Base is located, so well suited for flight research. The first was the area's flying conditions--clear skies with great visibility almost every day of the year. The second was the 44-square-mile Rogers Dry Lake, a natural landing site that General Boyd referred to as 'God's gift to the Air Force.' The third was the unpopulated area surrounding the lakebed, which led to fewer complaints about aircraft noise (including sonic booms) than would have occurred in more populated areas. There was also less chance of injury to the surrounding population in the event of an aircraft accident.

Linear Aerospike SR-71 Experiment (LASRE) refueling during first flight

NASA Technical Reports Server (NTRS)

1997-01-01

A NASA SR-71 refuels with an Edwards Air Force Base KC-135 during the first flight of the NASA/Rocketdyne/ Lockheed Martin Linear Aerospike SR-71 Experiment (LASRE). The flight took place Oct. 31 at NASA's Dryden Flight Research Center, Edwards, California. The SR-71 took off at 8:31 a.m. PST. The aircraft flew for one hour and fifty minutes, reaching a maximum speed of Mach 1.2 before landing at Edwards at 10:21 a.m. PST, successfully validating the SR-71/linear aerospike experiment configuration. The goal of the first flight was to evaluate the aerodynamic characteristics and the handling of the SR-71/linear aerospike experiment configuration. The engine was not fired during the flight. The LASRE experiment was designed to provide in-flight data to help Lockheed Martin evaluate the aerodynamic characteristics and the handling of the SR-71 linear aerospike experiment configuration. The goal of the project was to provide in-flight data to help Lockheed Martin validate the computational predictive tools it was using to determine the aerodynamic performance of a future reusable launch vehicle. The joint NASA, Rocketdyne (now part of Boeing), and Lockheed Martin Linear Aerospike SR-71 Experiment (LASRE) completed seven initial research flights at Dryden Flight Research Center. Two initial flights were used to determine the aerodynamic characteristics of the LASRE apparatus (pod) on the back of the SR-71. Five later flights focused on the experiment itself. Two were used to cycle gaseous helium and liquid nitrogen through the experiment to check its plumbing system for leaks and to test engine operational characteristics. During the other three flights, liquid oxygen was cycled through the engine. Two engine hot-firings were also completed on the ground. A final hot-fire test flight was canceled because of liquid oxygen leaks in the test apparatus. The LASRE experiment itself was a 20-percent-scale, half-span model of a lifting body shape (X-33) without the fins. It was rotated 90 degrees and equipped with eight thrust cells of an aerospike engine and was mounted on a housing known as the 'canoe,' which contained the gaseous hydrogen, helium, and instrumentation gear. The model, engine, and canoe together were called a 'pod.' The experiment focused on determining how a reusable launch vehicle's engine flume would affect the aerodynamics of its lifting-body shape at specific altitudes and speeds. The interaction of the aerodynamic flow with the engine plume could create drag; design refinements looked at minimizing this interaction. The entire pod was 41 feet in length and weighed 14,300 pounds. The experimental pod was mounted on one of NASA's SR-71s, which were at that time on loan to NASA from the U.S. Air Force. Lockheed Martin may use the information gained from the LASRE and X-33 Advanced Technology Demonstrator Projects to develop a potential future reusable launch vehicle. NASA and Lockheed Martin were partners in the X-33 program through a cooperative agreement. The goal of that program was to enable significant reductions in the cost of access to space and to promote creation and delivery of new space services and activities to improve the United States's economic competitiveness. In March 2001, however, NASA cancelled the X-33 program.

A NASA technician paints NASA's first Orion full-scale abort flight test crew module.

NASA Image and Video Library

2008-03-31

A full-scale flight-test mockup of the Constellation program's Orion crew vehicle arrived at NASA's Dryden Flight Research Center in late March 2008 to undergo preparations for the first short-range flight test of the spacecraft's astronaut escape system later that year. Engineers and technicians at NASA's Langley Research Center fabricated the structure, which precisely represents the size, outer shape and mass characteristics of the Orion space capsule. The Orion crew module mockup was ferried to NASA Dryden on an Air Force C-17. After painting in the Edwards Air Force Base paint hangar, the conical capsule was taken to Dryden for installation of flight computers, instrumentation and other electronics prior to being sent to the U.S. Army's White Sands Missile Range in New Mexico for integration with the escape system and the first abort flight test in late 2008. The tests were designed to ensure a safe, reliable method of escape for astronauts in case of an emergency.

Sporting a fresh paint job, NASA's first Orion full-scale abort flight test crew module awaits avionics and other equipment installation.

NASA Image and Video Library

2008-04-01

A full-scale flight-test mockup of the Constellation program's Orion crew vehicle arrived at NASA's Dryden Flight Research Center in late March 2008 to undergo preparations for the first short-range flight test of the spacecraft's astronaut escape system later that year. Engineers and technicians at NASA's Langley Research Center fabricated the structure, which precisely represents the size, outer shape and mass characteristics of the Orion space capsule. The Orion crew module mockup was ferried to NASA Dryden on an Air Force C-17. After painting in the Edwards Air Force Base paint hangar, the conical capsule was taken to Dryden for installation of flight computers, instrumentation and other electronics prior to being sent to the U.S. Army's White Sands Missile Range in New Mexico for integration with the escape system and the first abort flight test in late 2008. The tests were designed to ensure a safe, reliable method of escape for astronauts in case of an emergency.

The Ultrasonic Directional Tidal Breathing Pattern Sensor: Equitable Design Realization Based on Phase Information

PubMed Central

Sinharay, Arijit; Rakshit, Raj; Chakravarty, Tapas; Ghosh, Deb; Pal, Arpan

2017-01-01

Pulmonary ailments are conventionally diagnosed by spirometry. The complex forceful breathing maneuver as well as the extreme cost of spirometry renders it unsuitable in many situations. This work is aimed to facilitate an emerging direction of tidal breathing-based pulmonary evaluation by designing a novel, equitable, precise and portable device for acquisition and analysis of directional tidal breathing patterns, in real time. The proposed system primarily uses an in-house designed blow pipe, 40-kHz air-coupled ultrasound transreceivers, and a radio frequency (RF) phase-gain integrated circuit (IC). Moreover, in order to achieve high sensitivity in a cost-effective design philosophy, we have exploited the phase measurement technique, instead of selecting the contemporary time-of-flight (TOF) measurement; since application of the TOF principle in tidal breathing assessments requires sub-micro to nanosecond time resolution. This approach, which depends on accurate phase measurement, contributed to enhanced sensitivity using a simple electronics design. The developed system has been calibrated using a standard 3-L calibration syringe. The parameters of this system are validated against a standard spirometer, with maximum percentage error below 16%. Further, the extracted respiratory parameters related to tidal breathing have been found to be comparable with relevant prior works. The error in detecting respiration rate only is 3.9% compared to manual evaluation. These encouraging insights reveal the definite potential of our tidal breathing pattern (TBP) prototype for measuring tidal breathing parameters in order to extend the reach of affordable healthcare in rural regions and developing areas. PMID:28800103

Eclipse - tow flight closeup and release

NASA Technical Reports Server (NTRS)

1998-01-01

This clip, running 15 seconds in length, shows the QF-106 'Delta Dart' gear down, with the tow rope secured to the attachment point above the aircraft nose. First there is a view looking back from the C-141A, then looking forward from the nose of the QF-106, and finally a shot of the aircraft being released from the tow rope. NASA Dryden Flight Research Center, Edwards, California, supported a Kelly Space and Technology, Inc. (KST)/U.S. Air Force project known as Eclipse, which demonstrated a reusable tow launch vehicle concept. The purpose of the project was to demonstrate a reusable tow launch vehicle concept that had been conceived and patented by KST. Kelly Space obtained a contract with the USAF Research Laboratory for the tow launch demonstration project under the Small Business Innovation Research (SBIR) program. The USAF SBIR contract included the modifications to turn the QF-106 into the Experimental Demonstrator #1 (EXD-01), and the C141A aircraft to incorporate the tow provisions to link the two aircraft, as well as conducting flight tests. The demonstration consisted of ground and flight tests. These tests included a Combined Systems Test of both airplanes joined by a tow rope, a towed taxi test, and six towed flights. The primary goal of the project was demonstrating the tow phase of the Eclipse concept using a scaled-down tow aircraft (C-141A) and a representative aerodynamically-shaped aircraft (QF-106A) as a launch vehicle. This was successfully accomplished. On December 20, 1997, NASA research pilot Mark Stucky flew a QF-106 on the first towed flight behind an Air Force C-141 in the joint Eclipse project with KST to demonstrate a reusable tow launch vehicle concept developed by KST. Kelly Space and Technology hoped to use the data from the tow tests to validate a tow-to-launch procedure for reusable space launch vehicles. Stucky flew six successful tow tests between December 1997 and February 6, 1998. On February 6, 1998, the sixth and final towed flight brought the project to a successful completion. Preliminary flight results determined that the handling qualities of the QF-106 on tow were very stable; actual flight-measured values of tow rope tension were well within predictions made by the simulation, aerodynamic characteristics and elastic properties of the tow rope were a significant component of the towing system; and the Dryden high-fidelity simulation provided a representative model of the performance of the QF-106 and C-141A airplanes in tow configuration. Total time on tow for the entire project was 5 hours, 34 minutes, and 29 seconds. All six flights were highly productive, and all project objectives were achieved. All three of the project objectives were successfully accomplished. The objectives were: demonstration of towed takeoff, climb-out, and separation of the EXD-01 from the towing aircraft; validation of simulation models of the towed aircraft systems; and development of ground and flight procedures for towing and launching a delta-winged airplane configuration safely behind a transport-type aircraft. NASA Dryden served as the responsible test organization and had flight safety responsibility for the Eclipse project. Dryden also supplied engineering, simulation, instrumentation, range support, research pilots, and chase aircraft for the test series. Dryden personnel also performed the modifications to convert the QF-106 into the piloted EXD-01 aircraft. During the early flight phase of the project, Tracor, Inc. provided maintenance and ground support for the two QF-106 airplanes.The Air Force Flight Test Center (AFFTC), Edwards, California, provided the C-141A transport aircraft for the project, its flight and engineering support, and the aircrew. Kelly Space and Technology provided the modification design and fabrication of the hardware that was installed on the EXD-01 aircraft. Kelly Space and Technology hopes to use the data gleaned from the tow tests to develop a series of low-cost reusable launch vehicles, in particular to gain experience towing delta-wing aircraft having high wing loading, and in general to demonstrate various operational procedures such as ground processing and abort scenarios. The first successful towed flight occurred on Dec. 20, 1997. Prior to this first tow test flight, the C-141A and EXD-01 were used to conduct a series of tethered taxi tests to validate the tow procedures. Before these tethered taxi tests, a successful joint flight test was conducted in late October 1996, by Dryden, AFFTC, and KST, in which one of the Dryden F-18 chase aircraft flew at various ranges and locations behind the C-141A to define the wake turbulence and wingtip vortex environment. This flight test was replicated in July 1997, with an unmodified QF-106 flight proficiency aircraft.

Flight mechanics and control of escape manoeuvres in hummingbirds. II. Aerodynamic force production, flight control and performance limitations.

PubMed

Cheng, Bo; Tobalske, Bret W; Powers, Donald R; Hedrick, Tyson L; Wang, Yi; Wethington, Susan M; Chiu, George T-C; Deng, Xinyan

2016-11-15

The superior manoeuvrability of hummingbirds emerges from complex interactions of specialized neural and physiological processes with the unique flight dynamics of flapping wings. Escape manoeuvring is an ecologically relevant, natural behaviour of hummingbirds, from which we can gain understanding into the functional limits of vertebrate locomotor capacity. Here, we extend our kinematic analysis of escape manoeuvres from a companion paper to assess two potential limiting factors of the manoeuvring performance of hummingbirds: (1) muscle mechanical power output and (2) delays in the neural sensing and control system. We focused on the magnificent hummingbird (Eugenes fulgens, 7.8 g) and the black-chinned hummingbird (Archilochus alexandri, 3.1 g), which represent large and small species, respectively. We first estimated the aerodynamic forces, moments and the mechanical power of escape manoeuvres using measured wing kinematics. Comparing active-manoeuvring and passive-damping aerodynamic moments, we found that pitch dynamics were lightly damped and dominated by the effect of inertia, while roll dynamics were highly damped. To achieve observed closed-loop performance, pitch manoeuvres required faster sensorimotor transduction, as hummingbirds can only tolerate half the delay allowed in roll manoeuvres. Accordingly, our results suggested that pitch control may require a more sophisticated control strategy, such as those based on prediction. For the magnificent hummingbird, we estimated that escape manoeuvres required muscle mass-specific power 4.5 times that during hovering. Therefore, in addition to the limitation imposed by sensorimotor delays, muscle power could also limit the performance of escape manoeuvres. © 2016. Published by The Company of Biologists Ltd.

STS-99 Crew Interviews: Gorie

NASA Technical Reports Server (NTRS)

1999-01-01

Live footage of a preflight interview with Pilot Dominic L. Pudwell Gorie is seen. The interview addresses many different questions including why Gorie became an astronaut, the events that led to his interest, and his career path. Other interesting information that this one-on-one interview discusses is the purpose for the Shuttle Radar Topography Mission (SRTM). The main interest is on the importance of this SRTM flight, the knowledge we will learn gain from the 3D topographic map of the Earth, and the possible similarity to the tethered Satellite System Flight. The two antennas that will be taking the pictures, the involvement of the National Imagery and Mapping Agency (NIMA), mass deployment and retraction, gravity gradient force, flight cast maneuvers, EARTHCAM, a student-controlled camera on the Endeavour Orbiter, and Gorie's responsibility during this 24 hour mission.

STS-99 Crew Interviews: Gerhard P.J. Thiele

NASA Technical Reports Server (NTRS)

1999-01-01

Live footage of a preflight interview with Mission Specialist Gerhard P.J. Thiele is seen. The interview addresses many different questions including why Thiele became an astronaut, the events that led to his interest, and his career path. Other interesting information that this one-on-one interview discusses is the purpose for the Shuttle Radar Topography Mission (SRTM). The main interest is on the importance of this SRTM flight, the knowledge we will learn gain from the 3D topographic map of the Earth, and the possible similarity to the Tethered Satellite System Flight. The two antennas that will be taking the pictures, the involvement of the International partners, mass deployment and retraction, gravity gradient force, flight cast maneuvers, EARTHCAM, a student-controlled camera on the Endeavour Orbiter, and Thiele's responsibility during this 24 hour mission are also discussed.

Transient shifts in frontal and parietal circuits scale with enhanced visual feedback and changes in force variability and error

PubMed Central

Poon, Cynthia; Coombes, Stephen A.; Corcos, Daniel M.; Christou, Evangelos A.

2013-01-01

When subjects perform a learned motor task with increased visual gain, error and variability are reduced. Neuroimaging studies have identified a corresponding increase in activity in parietal cortex, premotor cortex, primary motor cortex, and extrastriate visual cortex. Much less is understood about the neural processes that underlie the immediate transition from low to high visual gain within a trial. This study used 128-channel electroencephalography to measure cortical activity during a visually guided precision grip task, in which the gain of the visual display was changed during the task. Force variability during the transition from low to high visual gain was characterized by an inverted U-shape, whereas force error decreased from low to high gain. Source analysis identified cortical activity in the same structures previously identified using functional magnetic resonance imaging. Source analysis also identified a time-varying shift in the strongest source activity. Superior regions of the motor and parietal cortex had stronger source activity from 300 to 600 ms after the transition, whereas inferior regions of the extrastriate visual cortex had stronger source activity from 500 to 700 ms after the transition. Force variability and electrical activity were linearly related, with a positive relation in the parietal cortex and a negative relation in the frontal cortex. Force error was nonlinearly related to electrical activity in the parietal cortex and frontal cortex by a quadratic function. This is the first evidence that force variability and force error are systematically related to a time-varying shift in cortical activity in frontal and parietal cortex in response to enhanced visual gain. PMID:23365186

X-15 ship #3 on lakebed

NASA Technical Reports Server (NTRS)

1961-01-01

The X-15-3 (56-6672), seen here on the lakebed at Edwards Air Force Base, Edwards, California, was a rocket-powered aircraft 50 ft long with a wingspan of 22 ft. It was a missile-shaped vehicle with an unusual wedge-shaped vertical tail, thin stubby wings, and unique side fairings that extended along the side of the fuselage. The X-15 weighed about 14,000 lb empty and approximately 34,000 lb at launch. The XLR-99 rocket engine, manufactured by Thiokol Chemical Corp., was pilot controlled and was capable of developing 57,000 lb of thrust. North American Aviation built three X-15 aircraft for the program. The X-15 research aircraft was developed to provide in-flight information and data on aerodynamics, structures, flight controls, and the physiological aspects of high-speed, high-altitude flight. A follow-on program used the aircraft as a testbed to carry various scientific experiments beyond the Earth's atmosphere on a repeated basis. For flight in the dense air of the usable atmosphere, the X-15 used conventional aerodynamic controls such as rudder surfaces on the vertical stabilizers to control yaw and movable horizontal stabilizers to control pitch when moving in synchronization or roll when moved differentially. For flight in the thin air outside of the appreciable Earth's atmosphere, the X-15 used a reaction control system. Hydrogen peroxide thrust rockets located on the nose of the aircraft provided pitch and yaw control. Those on the wings provided roll control. Because of the large fuel consumption, the X-15 was air launched from a B-52 aircraft at 45,000 ft and a speed of about 500 mph. Depending on the mission, the rocket engine provided thrust for the first 80 to 120 sec of flight. The remainder of the normal 10 to 11 min. flight was powerless and ended with a 200-mph glide landing. Generally, one of two types of X-15 flight profiles was used; a high-altitude flight plan that called for the pilot to maintain a steep rate of climb, or a speed profile that called for the pilot to push over and maintain a level altitude. The X-15 was flown over a period of nearly 10 years -- June 1959 to Oct. 1968 -- and set the world's unofficial speed and altitude records of 4,520 mph (Mach 6.7) and 354,200 ft in a program to investigate all aspects of manned hypersonic flight. Information gained from the highly successful X-15 program contributed to the development of the Mercury, Gemini, and Apollo manned spaceflight programs, and also the Space Shuttle program. The X-15s made a total of 199 flights, and were manufactured by North American Aviation. X-15-1 serial number 56-6670, seen in this photo, is now located at the National Air and Space Museum, Washington DC. North American X-15A-2, serial number 56-6671, is at the United States Air Force Museum, Wright-Patterson AFB, Ohio. The X-15-3, serial number 56-6672, crashed on 15 November 1967, resulting in the death of Maj. Michael J. Adams.

Proto-Flight Manipulator Arm (P-FMA)

NASA Technical Reports Server (NTRS)

Britton, W. R.

1977-01-01

The technical development of the Proto-Flight Manipulator Arm (P-FMA) which is a seven-degree-of-freedom general-purpose arm capable of being remotely operated in an earth orbital environment is discussed. The P-FMA is a unique manipulator, combining the capabilities of significant dexterity, high tip forces, precise motion control, gear backdriveability, high end effector grip forces and torques, and the quality of flightworthiness. The 2.4-meter (8-foot) arm weighs 52.2 kilograms (115 pounds).

Flight Qualification of the NASA's Super Pressure Balloon

NASA Astrophysics Data System (ADS)

Cathey, Henry; Said, Magdi; Fairbrother, Debora

Designs of new balloons to support space science require a number of actual flights under various flight conditions to qualify them to as standard balloon flight offerings to the science community. Development of the new Super Pressure Balloon for the National Aeronautics and Space Administration’s Balloon Program Office has entailed employing new design, analysis, and production techniques to advance the state of the art. Some of these advances have been evolutionary steps and some have been revolutionary steps requiring a maturing understanding of the materials, designs, and manufacturing approaches. The NASA Super Pressure Balloon development end goal is to produce a flight vehicle that is qualified to carry a ton of science instrumentation, at an altitude greater than 33 km while maintaining a near constant pressure altitude for extended periods of up to 100 days, and at any latitude on the globe. The NASA’s Balloon Program Office has pursued this development in a carefully executed incremental approach by gradually increasing payload carrying capability and increasing balloon volume to reach these end goal. A very successful test flight of a ~200,700 m3 balloon was launch in late 2008 from Antarctica. This balloon flew for over 54 days at a constant altitude and circled the Antarctic continent almost three times. A larger balloon was flown from Antarctica in early 2011. This ~422,400 m3 flew at a constant altitude for 22 days making one circuit around Antarctica. Although the performance was nominal, the flight was terminated via command to recover high valued assets from the payload. The balloon designed to reach the program goals is a ~532,200 m3 pumpkin shaped Super Pressure Balloon. A test flight of this balloon was launched from the Swedish Space Corporation’s Esrange Balloon Launch Facilities near Kiruna, Sweden on 14 August, 2012. This flight was another success for this development program. Valuable information was gained from this short test flight by successfully demonstrated balloon vehicle performance, obtained a large amount of videos, measured balloon differential pressure, obtained temperature and altitude data, assessed structure strength through pressurization, and demonstrated the balloon vehicles altitude stability. This flight was the first of several to qualify this design for the science community. Results of the most recent flights will be presented. Some of the related material characterization testing which is vital to the balloon design development for the balloon will also be presented. Additionally, this paper will provide a current overview of the development and qualification approach pursued for the NASA’s Super Pressure Balloon. Future plans and goals of future test flights will also be presented. This will include the projected balloon volumes, payload capabilities, test flight locations, and proposed flight schedule.

Pilot-in-the-Loop Analysis of Propulsive-Only Flight Control Systems

NASA Technical Reports Server (NTRS)

Chou, Hwei-Lan; Biezad, Daniel J.

1996-01-01

Longitudinal control system architectures are presented which directly couple flight stick motions to throttle commands for a multi-engine aircraft. This coupling enables positive attitude control with complete failure of the flight control system. The architectures chosen vary from simple feedback gains to classical lead-lag compensators with and without prefilters. Each architecture is reviewed for its appropriateness for piloted flight. The control systems are then analyzed with pilot-in-the-loop metrics related to bandwidth required for landing. Results indicate that current and proposed bandwidth requirements should be modified for throttles only flight control. Pilot ratings consistently showed better ratings than predicted by analysis. Recommendations are made for more robust design and implementation. The use of Quantitative Feedback Theory for compensator design is discussed. Although simple and effective augmented control can be achieved in a wide variety of failed configurations, a few configuration characteristics are dominant for pilot-in-the-loop control. These characteristics will be tested in a simulator study involving failed flight controls for a multi-engine aircraft.

Evaluation of balloon trajectory forecast routines for GAINS

NASA Astrophysics Data System (ADS)

Collander, R.; Girz, C.

The Global Air-ocean IN-situ System (GAINS) is a global observing system designed to augment current environmental observing and monitoring networks. GAINS is a network of long-duration, stratospheric platforms that carry onboard sensors and hundreds of dropsondes to acquire meteorological, air chemistry, and climate data over oceans and in remote land regions of the globe. Although GAINS platforms will include balloons and Remotely Operated Aircraft (ROA), the scope of this paper is limited to balloon-based platforms. A primary goal of GAINS balloon test flights is post-flight recovery of the balloon shell and payload, which requires information on the expected flight path and landing site prior to launch. Software has been developed for the prediction of the balloon trajectory and landing site, with separate versions written to generate predictions based upon rawinsonde data and model output. Balloon positions are calculated in 1-min increments based on wind data from the closest rawinsonde site or model grid point, given a known launch point, ascent and descent rate and flight duration. For short flights (< 6h), rawinsonde winds interpolated to 10-mb levels are used for trajectory calculations. Predictions for flight durations of 6 to 48h are based upon the initialization and 3 h forecast wind fields from NOAA's global aviation- (AVN) and Rapid Update Cycle (RUC) models. Given a limited number of actual balloon launches, trajectories computed from a chronological series of hourly RUC initializations are used as the baseline for comparison purposes. These baseline trajectories are compared to trajectory predictions from the rawinsonde and model-based versions on a monthly and seasonal basis over a 1-year period (January 1 - December 31, 2001) for flight durations of 3h, 6h and 48h. Predicted trajectories diverge from the baseline path, with the divergence increasing with increasing time. We examine the zonal, meridional and net magnitudes of these deviations, and attempt to determine directional biases in the predictions. This paper gives an overview of the software, including methods employed, physical considerations and limitations, and discusses results of this evaluation.

Wind-tunnel free-flight investigation of a 0.15-scale model of the F-106B airplane with vortex flaps

NASA Technical Reports Server (NTRS)

Yip, Long P.

1987-01-01

An investigation to determine the effects of vortex flaps on the flight dynamic characteristics of the F-106B in the area of low-speed, high-angle-of-attack flight was undertaken on a 0.15-scale model of the airplane in the Langley 30- by 60-Foot Tunnel. Static force tests, dynamic forced-oscillation tests, as well as free-flight tests were conducted to obtain a data base on the flight characteristics of the F-106B airplane with vortex flaps. Vortex flap configurations tested included a full-span gothic flap, a full-span constant-chord flap, and a part-span gothic flap.

Measuring Workload Differences Between Short-term Memory and Long-term Memory Scenarios in a Simulated Flight Environment

NASA Technical Reports Server (NTRS)

Berg, S. L.; Sheridan, T. B.

1984-01-01

Four highly experienced Air Force pilots each flew four simulated flight scenarios. Two scenarios required a great deal of aircraft maneuvering. The other two scenarios involved less maneuvering, but required remembering a number of items. All scenarios were designed to be equaly challenging. Pilot's Subjective Ratings for Activity-level, Complexity, Difficulty, Stress, and Workload were higher for the manuevering scenarios than the memory scenarios. At a moderate workload level, keeping the pilots active resulted in better aircraft control. When required to monitor and remember items, aircraft control tended to decrease. Pilots tended to weigh information about the spatial positioning and performance of their aircraft more heavily than other items.

Control of constraint forces and trajectories in a rich sensory and actuation environment.

PubMed

Hemami, Hooshang; Dariush, Behzad

2010-12-01

A simple control strategy is proposed and applied to a class of non-linear systems that have abundant sensory and actuation channels as in living systems. The main objective is the independent control of constrained trajectories of motion, and control of the corresponding constraint forces. The peripheral controller is a proportional, derivative and integral (PID) controller. A central controller produces, via pattern generators, reference signals that are the desired constrained position and velocity trajectories, and the desired constraint forces. The basic tenet of the this hybrid control strategy is the use of two mechanisms: 1. linear state and force feedback, and 2. non-linear constraint velocity feedback - sliding mode feedback. The first mechanism can be envisioned as a high gain feedback systems. The high gain attribute imitates the agonist-antagonist co-activation in natural systems. The strategy is applied to the control of the force and trajectory of a two-segment thigh-leg planar biped leg with a mass-less foot cranking a pedal that is analogous to a bicycle pedal. Five computational experiments are presented to show the effectiveness of the strategy and the performance of the controller. The findings of this paper are applicable to the design of orthoses and prostheses to supplement functional electrical stimulation for support purposes in the spinally injured cases. Copyright © 2010 Elsevier Inc. All rights reserved.

The free-flight response of Drosophila to motion of the visual environment.

PubMed

Mronz, Markus; Lehmann, Fritz-Olaf

2008-07-01

In the present study we investigated the behavioural strategies with which freely flying fruit flies (Drosophila) control their flight trajectories during active optomotor stimulation in a free-flight arena. We measured forward, turning and climbing velocities of single flies using high-speed video analysis and estimated the output of a 'Hassenstein-Reichardt' elementary motion detector (EMD) array and the fly's gaze to evaluate flight behaviour in response to a rotating visual panorama. In a stationary visual environment, flight is characterized by flight saccades during which the animals turn on average 120 degrees within 130 ms. In a rotating environment, the fly's behaviour typically changes towards distinct, concentric circular flight paths where the radius of the paths increases with increasing arena velocity. The EMD simulation suggests that this behaviour is driven by a rotation-sensitive EMD detector system that minimizes retinal slip on each compound eye, whereas an expansion-sensitive EMD system with a laterally centred visual focus potentially helps to achieve centring response on the circular flight path. We developed a numerical model based on force balance between horizontal, vertical and lateral forces that allows predictions of flight path curvature at a given locomotor capacity of the fly. The model suggests that turning flight in Drosophila is constrained by the production of centripetal forces needed to avoid side-slip movements. At maximum horizontal velocity this force may account for up to 70% of the fly's body weight during yaw turning. Altogether, our analyses are widely consistent with previous studies on Drosophila free flight and those on the optomotor response under tethered flight conditions.

Aerodynamic characteristics of the ventilated design for flapping wing micro air vehicle.

PubMed

Zhang, G Q; Yu, S C M

2014-01-01

Inspired by superior flight performance of natural flight masters like birds and insects and based on the ventilating flaps that can be opened and closed by the changing air pressure around the wing, a new flapping wing type has been proposed. It is known that the net lift force generated by a solid wing in a flapping cycle is nearly zero. However, for the case of the ventilated wing, results for the net lift force are positive which is due to the effect created by the "ventilation" in reducing negative lift force during the upstroke. The presence of moving flaps can serve as the variable in which, through careful control of the areas, a correlation with the decrease in negative lift can be generated. The corresponding aerodynamic characteristics have been investigated numerically by using different flapping frequencies and forward flight speeds.

Aerodynamic Characteristics of the Ventilated Design for Flapping Wing Micro Air Vehicle

PubMed Central

Zhang, G. Q.; Yu, S. C. M.

2014-01-01

Inspired by superior flight performance of natural flight masters like birds and insects and based on the ventilating flaps that can be opened and closed by the changing air pressure around the wing, a new flapping wing type has been proposed. It is known that the net lift force generated by a solid wing in a flapping cycle is nearly zero. However, for the case of the ventilated wing, results for the net lift force are positive which is due to the effect created by the “ventilation” in reducing negative lift force during the upstroke. The presence of moving flaps can serve as the variable in which, through careful control of the areas, a correlation with the decrease in negative lift can be generated. The corresponding aerodynamic characteristics have been investigated numerically by using different flapping frequencies and forward flight speeds. PMID:24683339

Space Science in Project SMART: A UNH High School Outreach Program

NASA Astrophysics Data System (ADS)

Smith, C. W.; Broad, L.; Goelzer, S.; Lessard, M.; Levergood, R.; Lugaz, N.; Moebius, E.; Schwadron, N.; Torbert, R. B.; Zhang, J.; Bloser, P. F.

2016-12-01

Every summer for the past 25 years the University of New Hampshire (UNH) has run a month-long, residential outreach program for high school students considering careers in mathematics, science, or engineering. Space science is one of the modules. Students work directly with UNH faculty performing original work with real spacecraft data and hardware and present the results of that effort at the end of the program. Recent research topics have included interplanetary waves and turbulence as recorded by the ACE and Voyager spacecraft, electromagnetic ion cyclotron (EMIC) waves seen by the RBSP spacecraft, interplanetary coronal mass ejections (ICME) acceleration and interstellar pickup ions as seen by the STEREO spacecraft, and prototyping CubeSat hardware. Student research efforts can provide useful results for future research efforts by the faculty while the students gain unique exposure to space physics and a science career. In addition, the students complete a team project. Since 2006, that project has been the construction and flight of a high-altitude balloon payload and instruments. The students typically build the instruments they fly. In the process, students learn circuit design and construction, microcontroller programming, and core atmospheric and space science. Our payload design has evolved significantly since the first flight of a simple rectangular box and now involves a stable descent vehicle that does not require a parachute, an on-board flight control computer, in-flight autonomous control and data acquisition of multiple student-built instruments, and real-time camera images sent to ground. This is a program that can be used as a model for other schools to follow and that high schools can initiate. More information can be found at .

[Aviation medicine: yesterday, today and tomorrow].

PubMed

Iamenskov, V V; Khafizov, N N; Morozov, A V

2011-04-01

The article describes the history and current state of medical support of the Air Force. On December 1, 2009 Air Force Medical Service was renamed into the service of aviation medicine of Command Air Forces, and many of the functions of medical support of the Air Force transferred to the newly formed Air Force Center of Aviation Medicine. April 29, 2011 office of Aviation Medicine of the High Command of the Air Force has stepped 95-year milestone. The changes affected all structures entrusted with the issues of medical support of the Air Force. Today, the medical service of the Air Force faces challenges--ensuring safety, the study of flight conditions and their impact on health, job performance and psychological characteristics of flight personnel.

KENNEDY SPACE CENTER, FLA. - One of four rudder speed brake actuators arrives at Cape Canaveral Air Force Station. The actuators, to be installed on the orbiter Discovery, are being X-rayed at the Radiographic High-Energy X-ray Facility to determine if the gears were installed correctly. Discovery has been assigned to the first Return to Flight mission, STS-114, a logistics flight to the International Space Station.

NASA Image and Video Library

2004-03-08

KENNEDY SPACE CENTER, FLA. - One of four rudder speed brake actuators arrives at Cape Canaveral Air Force Station. The actuators, to be installed on the orbiter Discovery, are being X-rayed at the Radiographic High-Energy X-ray Facility to determine if the gears were installed correctly. Discovery has been assigned to the first Return to Flight mission, STS-114, a logistics flight to the International Space Station.

Design and Fabrication of a Dielectric Total Internal Reflecting Solar Concentrator and Associated Flux Extractor for Extreme High Temperature (2500K) Applications

NASA Technical Reports Server (NTRS)

Soules, Jack A.; Buchele, Donald R.; Castle, Charles H.; Macosko, Robert P.

1997-01-01

The Analex Corporation, under contract to the NASA Lewis Research Center (LeRC), Cleveland, Ohio, recently evaluated the feasibility of utilizing refractive secondary concentrators for solar heat receivers operating at temperatures up to 2500K. The feasibility study pointed out a number of significant advantages provided by solid single crystal refractive devices over the more conventional hollow reflective compound parabolic concentrators (CPCs). In addition to the advantages of higher concentration ratio and efficiency, the refractive concentrator, when combined with a flux extractor rod, provides for flux tailoring within the heat receiver cavity. This is a highly desirable, almost mandatory, feature for solar thermal propulsion engine designs presently being considered for NASA and Air Force thermal applications. Following the feasibility evaluation, the NASA-LeRC, NASA-Marshall Space Flight Center (MSFC), and Analex Corporation teamed up to design, fabricate, and test a refractive secondary concentrator/flux extractor system for potential use in the NASA-MSFC "Shooting Star" flight experiment. This paper describes the advantages and technical challenges associated with the design methodologies developed and utilized and the material and fabrication limitations encountered.

The importance of the diurnal and annual cycle of air traffic for contrail radiative forcing.

PubMed

Stuber, Nicola; Forster, Piers; Rädel, Gaby; Shine, Keith

2006-06-15

Air traffic condensation trails, or contrails, are believed to have a net atmospheric warming effect, although one that is currently small compared to that induced by other sources of human emissions. However, the comparably large growth rate of air traffic requires an improved understanding of the resulting impact of aircraft radiative forcing on climate. Contrails have an effect on the Earth's energy balance similar to that of high thin ice clouds. Their trapping of outgoing longwave radiation emitted by the Earth and atmosphere (positive radiative forcing) is partly compensated by their reflection of incoming solar radiation (negative radiative forcing). On average, the longwave effect dominates and the net contrail radiative forcing is believed to be positive. Over daily and annual timescales, varying levels of air traffic, meteorological conditions, and solar insolation influence the net forcing effect of contrails. Here we determine the factors most important for contrail climate forcing using a sophisticated radiative transfer model for a site in southeast England, located in the entrance to the North Atlantic flight corridor. We find that night-time flights during winter (December to February) are responsible for most of the contrail radiative forcing. Night flights account for only 25 per cent of daily air traffic, but contribute 60 to 80 per cent of the contrail forcing. Further, winter flights account for only 22 per cent of annual air traffic, but contribute half of the annual mean forcing. These results suggest that flight rescheduling could help to minimize the climate impact of aviation.

Aerodynamics and flow features of a damselfly in takeoff flight.

PubMed

Bode-Oke, Ayodeji T; Zeyghami, Samane; Dong, Haibo

2017-09-26

Flight initiation is fundamental for survival, escape from predators and lifting payload from one place to another in biological fliers and can be broadly classified into jumping and non-jumping takeoffs. During jumping takeoffs, the legs generate most of the initial impulse. Whereas the wings generate most of the forces in non-jumping takeoffs, which are usually voluntary, slow, and stable. It is of great interest to understand how these non-jumping takeoffs occur and what strategies insects use to generate large amount of forces required for this highly demanding flight initiation mode. Here, for the first time, we report accurate wing and body kinematics measurements of a damselfly during a non-jumping takeoff. Furthermore, using a high fidelity computational fluid dynamics simulation, we identify the 3D flow features and compute the wing aerodynamics forces to unravel the key mechanisms responsible for generating large flight forces. Our numerical results show that a damselfly generates about three times its body weight during the first half-stroke for liftoff. In generating these forces, the wings flap through a steeply inclined stroke plane with respect to the horizon, slicing through the air at high angles of attack (45°-50°). Consequently, a leading edge vortex (LEV) is formed during both the downstroke and upstroke on all the four wings. The formation of the LEV, however, is inhibited in the subsequent upstrokes following takeoff. Accordingly, we observe a drastic reduction in the magnitude of the aerodynamic force, signifying the importance of LEV in augmenting force production. Our analysis also shows that forewing-hindwing interaction plays a favorable role in enhancing both lift and thrust production during takeoff.

F-18 HARV research pilot Dana Purifoy

NASA Technical Reports Server (NTRS)

1996-01-01

Dana D. Purifoy is an aerospace research pilot at NASA's Dryden Flight Research Center, Edwards, California. He joined NASA in August 1994. Purifoy is a former Air Force test pilot who served as a project pilot in the joint NASA/Air Force X-29 Forward Swept Wing research program conducted at Dryden from 1984 to 1991. His most recent assignment in the Air Force was flying U-2 aircraft as a test pilot at Air Force Plant 42, Palmdale, CA. In addition to flying the X-29 at Dryden as an Air Force pilot, Purifoy also served as project pilot and joint test force director with the AFTI F-16 (Advanced Fighter Technology Integration/F-16) program, also located at Dryden. Before his assignments as project pilot on the X-29 and AFTI/F-16 aircraft, Purifoy was chief of the Academics Systems Branch at the Air Force Test Pilot School at Edwards. Prior to becoming a test pilot, he flew F-111 and F-16 aircraft in Great Britain and Germany. He has accumulated 3800 hours of flying time in his career. The final flight for the F-18 High Alpha Research Vehicle (HARV) took place at NASA Dryden on May 29, 1996. The highly modified F-18 airplane flew 383 flights over a nine year period and demonstrated concepts that greatly increase fighter maneuverability. Among concepts proven in the aircraft is the use of paddles to direct jet engine exhaust in cases of extreme altitudes where conventional control surfaces lose effectiveness. Another concept, developed by NASA Langley Research Center, is a deployable wing-like surface installed on the nose of the aircraft for increased right and left (yaw) control on nose-high flight angles.

Adding tactile realism to a virtual reality laparoscopic surgical simulator with a cost-effective human interface device

NASA Astrophysics Data System (ADS)

Mack, Ian W.; Potts, Stephen; McMenemy, Karen R.; Ferguson, R. S.

2006-02-01

The laparoscopic technique for performing abdominal surgery requires a very high degree of skill in the medical practitioner. Much interest has been focused on using computer graphics to provide simulators for training surgeons. Unfortunately, these tend to be complex and have a very high cost, which limits availability and restricts the length of time over which individuals can practice their skills. With computer game technology able to provide the graphics required for a surgical simulator, the cost does not have to be high. However, graphics alone cannot serve as a training simulator. Human interface hardware, the equivalent of the force feedback joystick for a flight simulator game, is required to complete the system. This paper presents a design for a very low cost device to address this vital issue. The design encompasses: the mechanical construction, the electronic interfaces and the software protocols to mimic a laparoscopic surgical set-up. Thus the surgeon has the capability of practicing two-handed procedures with the possibility of force feedback. The force feedback and collision detection algorithms allow surgeons to practice realistic operating theatre procedures with a good degree of authenticity.

Vacuum Plasma Spray Forming of Tungsten Lorentz Force Accelerator Components

NASA Technical Reports Server (NTRS)

Zimmerman, Frank R.

2001-01-01

The Vacuum Plasma Spray (VPS) Laboratory at NASA's Marshall Space Flight Center has developed and demonstrated a fabrication technique using the VPS process to form anode sections for a Lorentz force accelerator from tungsten. Lorentz force accelerators are an attractive form of electric propulsion that provides continuous, high-efficiency propulsion at useful power levels for such applications as orbit transfers or deep space missions. The VPS process is used to deposit refractory metals such as tungsten onto a graphite mandrel of the desired shape. Because tungsten is reactive at high temperatures, it is thermally sprayed in an inert environment where the plasma gun melts and accelerates the metal powder onto the mandrel. A three-axis robot inside the chamber controls the motion of the plasma spray torch. A graphite mandrel acts as a male mold, forming the required contour and dimensions of the inside surface of the anode. This paper describes the processing techniques, design considerations, and process development associated with the VPS forming of the Lorentz force accelerator.

NASA's first Orion full-scale abort flight test crew module was placed in NASA Dryden's Abort Flight Test integration area for equipment installation.

NASA Image and Video Library

2008-04-01

A full-scale flight-test mockup of the Constellation program's Orion crew vehicle arrived at NASA's Dryden Flight Research Center in late March 2008 to undergo preparations for the first short-range flight test of the spacecraft's astronaut escape system later that year. Engineers and technicians at NASA's Langley Research Center fabricated the structure, which precisely represents the size, outer shape and mass characteristics of the Orion space capsule. The Orion crew module mockup was ferried to NASA Dryden on an Air Force C-17. After painting in the Edwards Air Force Base paint hangar, the conical capsule was taken to Dryden for installation of flight computers, instrumentation and other electronics prior to being sent to the U.S. Army's White Sands Missile Range in New Mexico for integration with the escape system and the first abort flight test in late 2008. The tests were designed to ensure a safe, reliable method of escape for astronauts in case of an emergency.

NASA Dryden Flight Research Center personnel accompany NASA's first Orion full-scale abort flight test crew module as it heads to its new home.

NASA Image and Video Library

2008-04-01

A full-scale flight-test mockup of the Constellation program's Orion crew vehicle arrived at NASA's Dryden Flight Research Center in late March 2008 to undergo preparations for the first short-range flight test of the spacecraft's astronaut escape system later that year. Engineers and technicians at NASA's Langley Research Center fabricated the structure, which precisely represents the size, outer shape and mass characteristics of the Orion space capsule. The Orion crew module mockup was ferried to NASA Dryden on an Air Force C-17. After painting in the Edwards Air Force Base paint hangar, the conical capsule was taken to Dryden for installation of flight computers, instrumentation and other electronics prior to being sent to the U.S. Army's White Sands Missile Range in New Mexico for integration with the escape system and the first abort flight test in late 2008. The tests were designed to ensure a safe, reliable method of escape for astronauts in case of an emergency.

Design, calibration and testing of a force balance for a hypersonic shock tunnel

NASA Astrophysics Data System (ADS)

Vadassery, Pravin

The forces acting on a flight vehicle are critical for determining its performance. Of particular interest is the hypersonic regime. Force measurements are much more complex in hypersonic flows, where those speeds are simulated in shock tunnels. A force balance for such facilities contains sensitive gages that measure stress waves and ultimately determine the different components of force acting on the model. An external force balance was designed and fabricated for the UTA Hypersonic shock tunnel to measure drag at Mach 10. Static and dynamic calibrations were performed to find the transfer function of the system. Forces were recovered using a deconvolution procedure. To validate the force balance, experiments were conducted on a blunt cone. The measured forces were compared to Newtonian theory.

Eclipse program C-141A aircraft

NASA Technical Reports Server (NTRS)

1997-01-01

This photograph shows the Air Force C-141A that was used in the Eclipse project as a tow vehicle. In 1997 and 1998, the Dryden Flight Research Center at Edwards, California, supported and hosted a Kelly Space & Technology, Inc. project called Eclipse, which sought to demonstrate the feasibility of a reusable tow-launch vehicle concept. The project goal was to successfully tow, inflight, a modified QF-106 delta-wing aircraft with an Air Force C-141A transport aircraft. This would demonstrate the possibility of towing and launching an actual launch vehicle from behind a tow plane. Dryden was the responsible test organization and had flight safety responsibility for the Eclipse project. Dryden provided engineering, instrumentation, simulation, modification, maintenance, range support, and research pilots for the test program. The Air Force Flight Test Center (AFFTC), Edwards, California, supplied the C-141A transport aircraft and crew and configured the aircraft as needed for the tests. The AFFTC also provided the concept and detail design and analysis as well as hardware for the tow system and QF-106 modifications. Dryden performed the modifications to convert the QF-106 drone into the piloted EXD-01 (Eclipse eXperimental Demonstrator-01) experimental aircraft. Kelly Space & Technology hoped to use the results gleaned from the tow test in developing a series of low-cost, reusable launch vehicles. These tests demonstrated the validity of towing a delta-wing aircraft having high wind loading, validated the tow simulation model, and demonstrated various operational procedures, such as ground processing of in-flight maneuvers and emergency abort scenarios.

X-15 #3 pedestal-mounted full-scale replica covered in snow

NASA Technical Reports Server (NTRS)

1997-01-01

The full scale mock-up of X-15 #3 was installed September 1995 at the NASA Dryden Flight Research Center, Edwards, California. The original X-15 #3, serial number 56-6672, was destroyed on 15 November 1967, in a crash that also fatally injured pilot Maj. Michael J. Adams. The X-15 was a rocket-powered aircraft 50 ft long with a wingspan of 22 ft. It was a missile-shaped vehicle with an unusual wedge-shaped vertical tail, thin stubby wings, and unique side fairings that extended along the side of the fuselage. The X-15 weighed about 14,000 lb empty and approximately 34,000 lb at launch. The XLR-99 rocket engine, manufactured by Thiokol Chemical Corp., was pilot controlled and was capable of developing 57,000 lb of thrust. North American Aviation built three X-15 aircraft for the program. The X-15 research aircraft was developed to provide in-flight information and data on aerodynamics, structures, flight controls, and the physiological aspects of high-speed, high-altitude flight. A follow on program used the aircraft as a testbed to carry various scientific experiments beyond the Earth's atmosphere on a repeated basis. For flight in the dense air of the usable atmosphere, the X-15 used conventional aerodynamic controls such as rudder surfaces on the vertical stabilizers to control yaw and movable horizontal stabilizers to control pitch when moving in synchronization or roll when moved differentially. For flight in the thin air outside of the appreciable Earth's atmosphere, the X-15 used a reaction control system. Hydrogen peroxide thrust rockets located on the nose of the aircraft provided pitch and yaw control. Those on the wings provided roll control. Because of the large fuel consumption, the X-15 was air launched from a B-52 aircraft at 45,000 ft and a speed of about 500 mph. Depending on the mission, the rocket engine provided thrust for the first 80 to 120 sec of flight. The remainder of the normal 10 to 11 min. flight was powerless and ended with a 200-mph glide landing. Generally, one of two types of X-15 flight profiles was used; a high-altitude flight plan that called for the pilot to maintain a steep rate of climb, or a speed profile that called for the pilot to push over and maintain a level altitude. The X-15 was flown over a period of nearly 10 years -- June 1959 to Oct. 1968 -- and set the world's unofficial speed and altitude records of 4,520 mph (Mach 6.7) and 354,200 ft in a program to investigate all aspects of manned hypersonic flight. Information gained from the highly successful X-15 program contributed to the development of the Mercury, Gemini, and Apollo manned spaceflight programs, and also the Space Shuttle program. The X-15s made a total of 199 flights, and were manufactured by North American Aviation. X-15-1, serial number 56-6670, is now located at the National Air and Space Museum, Washington DC. North American X-15A-2, serial number 56-6671, is at the United States Air Force Museum, Wright-Patterson AFB, Ohio. Parts of the crashed X-15-3, recovered in 1992 by Peter Merlin and Tony Moore (The X-Hunters) are on display at the Air Force Flight Test Center Museum at Edwards. The canopy from the X-15-3, recovered during the original search in 1967, is displayed at the San Diego Aerospace Museum, San Diego, California.

X-15A-2 with test pilot Pete Knight

NASA Technical Reports Server (NTRS)

1965-01-01

Air Force pilot William J. 'Pete' Knight is seen here in front of the X-15A-2 aircraft (56-6671). Pete Knight made 16 flights in the X-15, and set the world unofficial speed record for fixed wing aircraft, 4,520 mph (mach 6.7), in the X-15A-2. He also made one flight above 50 miles, qualifying him for astronaut wings. The X-15 was a rocket-powered aircraft 50 ft long with a wingspan of 22 ft. It was a missile-shaped vehicle with an unusual wedge-shaped vertical tail, thin stubby wings, and unique fairings that extended along the side of the fuselage. The X-15 weighed about 14,000 lb empty and approximately 34,000 lb at launch. The XLR-99 rocket engine, manufactured by Thiokol Chemical Corp., was pilot controlled and was capable of developing 57,000 lb of rated thrust (actual thrust reportedly climbed to 60,000 lb). North American Aviation built three X-15 aircraft for the program. The X-15 research aircraft was developed to provide in-flight information and data on aerodynamics, structures, flight controls, and the physiological aspects of high-speed, high-altitude flight. A follow-on program used the aircraft as a testbed to carry various scientific experiments beyond the Earth's atmosphere on a repeated basis. For flight in the dense air of the usable atmosphere, the X-15 used conventional aerodynamic controls such as rudder surfaces on the vertical stabilizers to control yaw and canted horizontal surfaces on the tail to control pitch when moving in synchronization or roll when moved differentially. For flight in the thin air outside of the appreciable Earth's atmosphere, the X-15 used a reaction control system. Hydrogen peroxide thrust rockets located on the nose of the aircraft provided pitch and yaw control. Those on the wings provided roll control. Because of the large fuel consumption, the X-15 was air launched from a B-52 aircraft at 45,000 ft and a speed of about 500 mph. Depending on the mission, the rocket engine provided thrust for the first 80 to 120 sec of flight. The remainder of the normal 10 to 11 min. flight was powerless and ended with a 200-mph glide landing. Generally, one of two types of X-15 flight profiles was used: a high-altitude flight plan that called for the pilot to maintain a steep rate of climb, or a speed profile that called for the pilot to push over and maintain a level altitude. The X-15 was flown over a period of nearly 10 years--June 1959 to Oct. 1968--and set the world's unofficial speed and altitude records of 4,520 mph (Mach 6.7) and 354,200 ft (over 67 mi) in a program to investigate all aspects of piloted hypersonic flight. Information gained from the highly successful X-15 program contributed to the development of the Mercury, Gemini, and Apollo manned spaceflight programs, and also the Space Shuttle program. The X-15s made a total of 199 flights and were manufactured by North American Aviation. X-15-1, serial number 56-6670, is now located at the National Air and Space Museum, Washington DC. North American X-15A-2, serial number 56-6671, is at the United States Air Force Museum, Wright-Patterson AFB, Ohio. The X-15-3, serial number 56-6672, crashed on 15 November 1967, resulting in the death of Maj. Michael J. Adams.

F-16XL Hybrid Reynolds-Averaged Navier-Stokes/Large Eddy Simulation on Unstructured Grids

NASA Technical Reports Server (NTRS)

Park, Michael A.; Abdol-Hamid, Khaled S.; Elmiligui, Alaa

2015-01-01

This study continues the Cranked Arrow Wing Aerodynamics Program, International (CAWAPI) investigation with the FUN3D and USM3D flow solvers. CAWAPI was established to study the F-16XL, because it provides a unique opportunity to fuse fight test, wind tunnel test, and simulation to understand the aerodynamic features of swept wings. The high-lift performance of the cranked-arrow wing planform is critical for recent and past supersonic transport design concepts. Simulations of the low speed high angle of attack Flight Condition 25 are compared: Detached Eddy Simulation (DES), Modi ed Delayed Detached Eddy Simulation (MDDES), and the Spalart-Allmaras (SA) RANS model. Iso- surfaces of Q criterion show the development of coherent primary and secondary vortices on the upper surface of the wing that spiral, burst, and commingle. SA produces higher pressure peaks nearer to the leading-edge of the wing than flight test measurements. Mean DES and MDDES pressures better predict the flight test measurements, especially on the outer wing section. Vorticies and vortex-vortex interaction impact unsteady surface pressures. USM3D showed many sharp tones in volume points spectra near the wing apex with low broadband noise and FUN3D showed more broadband noise with weaker tones. Spectra of the volume points near the outer wing leading-edge was primarily broadband for both codes. Without unsteady flight measurements, the flight pressure environment can not be used to validate the simulations containing tonal or broadband spectra. Mean forces and moment are very similar between FUN3D models and between USM3D models. Spectra of the unsteady forces and moment are broadband with a few sharp peaks for USM3D.

Design and Performance of Insect-Scale Flapping-Wing Vehicles

NASA Astrophysics Data System (ADS)

Whitney, John Peter

Micro-air vehicles (MAVs)---small versions of full-scale aircraft---are the product of a continued path of miniaturization which extends across many fields of engineering. Increasingly, MAVs approach the scale of small birds, and most recently, their sizes have dipped into the realm of hummingbirds and flying insects. However, these non-traditional biologically-inspired designs are without well-established design methods, and manufacturing complex devices at these tiny scales is not feasible using conventional manufacturing methods. This thesis presents a comprehensive investigation of new MAV design and manufacturing methods, as applicable to insect-scale hovering flight. New design methods combine an energy-based accounting of propulsion and aerodynamics with a one degree-of-freedom dynamic flapping model. Important results include analytical expressions for maximum flight endurance and range, and predictions for maximum feasible wing size and body mass. To meet manufacturing constraints, the use of passive wing dynamics to simplify vehicle design and control was investigated; supporting tests included the first synchronized measurements of real-time forces and three-dimensional kinematics generated by insect-scale flapping wings. These experimental methods were then expanded to study optimal wing shapes and high-efficiency flapping kinematics. To support the development of high-fidelity test devices and fully-functional flight hardware, a new class of manufacturing methods was developed, combining elements of rigid-flex printed circuit board fabrication with "pop-up book" folding mechanisms. In addition to their current and future support of insect-scale MAV development, these new manufacturing techniques are likely to prove an essential element to future advances in micro-optomechanics, micro-surgery, and many other fields.

Design and simulation of flight control system for man-portable micro reconnaissance quadcopter

NASA Astrophysics Data System (ADS)

Yin, Xinfan; Zhang, Daibing; Fang, Qiang; Shen, Lincheng

2017-10-01

The quadcopter has been widely used in the field of aerial photography and environmental detection, because of its advantages of VTOL, simple structure, and easy-control. In the field of urban anti-terrorism or special operations, micro reconnaissance quadcpter has its unique advantages such as all-weather taking off and landing, small noise and so on, and it is very popular with special forces and riot police. This paper aims at the flight control problem of the micro quadcopter, for the purposes of attitude stabilization control and trajectory tracking control of the micro quadcopter, first, the modeling of the micro quadcopter is presented. And using the MATLAB/SIMULINK toolbox to build the flight controller of the micro quadcopter, and then simulation analysis and real flight test are given. The results of the experiment show that the designed PID controller can correct the flight attitude shift effectively and track the planned tracks well, and can achieve the goal of stable and reliable flight of the quadcopter. It can be a useful reference for the flight control system design of future special operations micro UAV.

A program for calculating load coefficient matrices utilizing the force summation method, L218 (LOADS). Volume 1: Engineering and usage

NASA Technical Reports Server (NTRS)

Miller, R. D.; Anderson, L. R.

1979-01-01

The LOADS program L218, a digital computer program that calculates dynamic load coefficient matrices utilizing the force summation method, is described. The load equations are derived for a flight vehicle in straight and level flight and excited by gusts and/or control motions. In addition, sensor equations are calculated for use with an active control system. The load coefficient matrices are calculated for the following types of loads: translational and rotational accelerations, velocities, and displacements; panel aerodynamic forces; net panel forces; shears and moments. Program usage and a brief description of the analysis used are presented. A description of the design and structure of the program to aid those who will maintain and/or modify the program in the future is included.

The Shock and Vibration Bulletin. Part 4. Structural Dynamics, Systems Identification, Computer Applications.

DTIC Science & Technology

1977-09-01

Division, Barry Wright Corporation, Watertown, MA DESIGN OF ELASTOMERIC COMPONENTS BY USING THE FINITE -" b ELEMENT TECHNIQUE R.H. Finney and B.P. Gupta...Alabama in Huntsville, Huntsville, AL PAPERS APPEARING IN PART 2 Vibration Analysis SOME ASPECTS OF VIBRATION CONTROL SUPPORT DESIGN 0 P. Bezler and J.R...at the Air Force Flight August 1968, pp. 239-248. Dynamics Laboratory (AFFDL). The laser force measuring mounting brackets were designed and 5. G. K

HL-10 on lakebed with pilot Major Jerauld R. Gentry

NASA Technical Reports Server (NTRS)

1968-01-01

Pilot Major Jerauld R. Gentry stands in front of the HL-10 Lifting Body. Gentry was the Air Force project pilot for the HL-10 while it was making the early glide and powered flights in 1968 following its modification. He made a total of nine flights in the vehicle. For his work on the HL-10, Gentry was awarded the Harmon International Trophy for his outstanding contribution to the science of flying. He later became the Air Force pilot for the X-24A. The HL-10 was one of five heavyweight lifting-body designs flown at NASA's Flight Research Center (FRC--later Dryden Flight Research Center), Edwards, California, from July 1966 to November 1975 to study and validate the concept of safely maneuvering and landing a low lift-over-drag vehicle designed for reentry from space. Northrop Corporation built the HL-10 and M2-F2, the first two of the fleet of 'heavy' lifting bodies flown by the NASA Flight Research Center. The contract for construction of the HL-10 and the M2-F2 was $1.8 million. 'HL' stands for horizontal landing, and '10' refers to the tenth design studied by engineers at NASA's Langley Research Center, Hampton, Va. After delivery to NASA in January 1966, the HL-10 made its first flight on Dec. 22, 1966, with research pilot Bruce Peterson in the cockpit. Although an XLR-11 rocket engine was installed in the vehicle, the first 11 drop flights from the B-52 launch aircraft were powerless glide flights to assess handling qualities, stability, and control. In the end, the HL-10 was judged to be the best handling of the three original heavy-weight lifting bodies (M2-F2/F3, HL-10, X-24A). The HL-10 was flown 37 times during the lifting body research program and logged the highest altitude and fastest speed in the Lifting Body program. On Feb. 18, 1970, Air Force test pilot Peter Hoag piloted the HL-10 to Mach 1.86 (1,228 mph). Nine days later, NASA pilot Bill Dana flew the vehicle to 90,030 feet, which became the highest altitude reached in the program. Some new and different lessons were learned through the successful flight testing of the HL-10. These lessons, when combined with information from it's sister ship, the M2-F2/F3, provided an excellent starting point for designers of future entry vehicles, including the Space Shuttle.

Eclipse program QF-106 aircraft in flight

NASA Technical Reports Server (NTRS)

1997-01-01

This photo shows one of the QF-106s used in the Eclipse project in flight. In 1997 and 1998, the Dryden Flight Research Center at Edwards, California, supported and hosted a Kelly Space & Technology, Inc. project called Eclipse, which sought to demonstrate the feasibility of a reusable tow-launch vehicle concept. The project goal was to successfully tow, inflight, a modified QF-106 delta-wing aircraft with an Air Force C-141A transport aircraft. This would demonstrate the possibility of towing and launching an actual launch vehicle from behind a tow plane. Dryden was the responsible test organization and had flight safety responsibility for the Eclipse project. Dryden provided engineering, instrumentation, simulation, modification, maintenance, range support, and research pilots for the test program. The Air Force Flight Test Center (AFFTC), Edwards, California, supplied the C-141A transport aircraft and crew and configured the aircraft as needed for the tests. The AFFTC also provided the concept and detail design and analysis as well as hardware for the tow system and QF-106 modifications. Dryden performed the modifications to convert the QF-106 drone into the piloted EXD-01 (Eclipse eXperimental Demonstrator-01) experimental aircraft. Kelly Space & Technology hoped to use the results gleaned from the tow test in developing a series of low-cost, reusable launch vehicles. These tests demonstrated the validity of towing a delta-wing aircraft having high wing loading, validated the tow simulation model, and demonstrated various operational procedures, such as ground processing of in-flight maneuvers and emergency abort scenarios.

Planned flight test of a mercury ion auxiliary propulsion system. 1: Objectives, systems descriptions, and mission operations

NASA Technical Reports Server (NTRS)

Power, J. C.

1978-01-01

A planned flight test of an 8 cm diameter, electron-bombardment mercury ion thruster system is described. The primary objective of the test is to flight qualify the 5 mN (1 mlb.) thruster system for auxiliary propulsion applications. A seven year north-south stationkeeping mission was selected as the basis for the flight test operating profile. The flight test, which will employ two thruster systems, will also generate thruster system space performance data, measure thruster-spacecraft interactions, and demonstrate thruster operation in a number of operating modes. The flight test is designated as SAMSO-601 and will be flown aboard the shuttle-launched Air Force space test program P80-1 satellite in 1981. The spacecraft will be 3- axis stabilized in its final 740 km circular orbit, which will have an inclination of approximately greater than 73 degrees. The spacecraft design lifetime is three years.

The implementation and operation of a variable-response electronic throttle control system for a TF-104G aircraft

NASA Technical Reports Server (NTRS)

Neal, Bradford; Sengupta, Upal

1989-01-01

During some flight programs, researchers have encountered problems in the throttle response characteristics of high-performance aircraft. To study and to help solve these problems, the National Aeronautics and Space Administration Ames Research Center's Dryden Flight Research Facility (Ames-Dryden) conducted a study using a TF-104G airplane modified with a variable-response electronic throttle control system. Ames-Dryden investigated the effects of different variables on engine response and handling qualities. The system provided transport delay, lead and lag filters, second-order lags, command rate and position limits, and variable gain between the pilot's throttle command and the engine fuel controller. These variables could be tested individually or in combination. Ten research flights were flown to gather data on engine response and to obtain pilot ratings of the various system configurations. The results should provide design criteria for engine-response characteristics. The variable-response throttle components and how they were installed in the TF-104G aircraft are described. How the variable-response throttle was used in flight and some of the results of using this system are discussed.

Flying at no mechanical energy cost: disclosing the secret of wandering albatrosses.

PubMed

Sachs, Gottfried; Traugott, Johannes; Nesterova, Anna P; Dell'Omo, Giacomo; Kümmeth, Franz; Heidrich, Wolfgang; Vyssotski, Alexei L; Bonadonna, Francesco

2012-01-01

Albatrosses do something that no other birds are able to do: fly thousands of kilometres at no mechanical cost. This is possible because they use dynamic soaring, a flight mode that enables them to gain the energy required for flying from wind. Until now, the physical mechanisms of the energy gain in terms of the energy transfer from the wind to the bird were mostly unknown. Here we show that the energy gain is achieved by a dynamic flight manoeuvre consisting of a continually repeated up-down curve with optimal adjustment to the wind. We determined the energy obtained from the wind by analysing the measured trajectories of free flying birds using a new GPS-signal tracking method yielding a high precision. Our results reveal an evolutionary adaptation to an extreme environment, and may support recent biologically inspired research on robotic aircraft that might utilize albatrosses' flight technique for engineless propulsion.

Flying at No Mechanical Energy Cost: Disclosing the Secret of Wandering Albatrosses

PubMed Central

Sachs, Gottfried; Traugott, Johannes; Nesterova, Anna P.; Dell'Omo, Giacomo; Kümmeth, Franz; Heidrich, Wolfgang

2012-01-01

Albatrosses do something that no other birds are able to do: fly thousands of kilometres at no mechanical cost. This is possible because they use dynamic soaring, a flight mode that enables them to gain the energy required for flying from wind. Until now, the physical mechanisms of the energy gain in terms of the energy transfer from the wind to the bird were mostly unknown. Here we show that the energy gain is achieved by a dynamic flight manoeuvre consisting of a continually repeated up-down curve with optimal adjustment to the wind. We determined the energy obtained from the wind by analysing the measured trajectories of free flying birds using a new GPS-signal tracking method yielding a high precision. Our results reveal an evolutionary adaptation to an extreme environment, and may support recent biologically inspired research on robotic aircraft that might utilize albatrosses' flight technique for engineless propulsion. PMID:22957014

Application of spring tabs to elevator controls

NASA Technical Reports Server (NTRS)

Phillips, William H

1944-01-01

Equations are presented for calculating the stick-force characteristics obtained with a spring-tab type of elevator control. The main problems encountered in the design of a satisfactory elevator spring tab are to provide stick forces in the desired range, to maintain the force per g sufficiently constant throughout the speed range, to avoid undesirable "feel" of the control in ground handling or in flight at low airspeeds, and to prevent flutter. Examples are presented to show the design features of spring tabs required to solve these problems for airplanes of various sizes.

STS-39 Space Shuttle mission report

NASA Technical Reports Server (NTRS)

Fricke, Robert W.

1991-01-01

The STS-39 Space Shuttle Program Mission Report contains a summary of the vehicle subsystem operations during the fortieth flight of the Space Shuttle and the twelfth flight of the Orbiter Vehicle Discovery (OV-103). In addition to the Discovery vehicle, the flight vehicle consisted of the following: an External Tank (ET) (designated as ET-46 (LWT-39); three Space Shuttle main engines (SSME's) (serial numbers 2026, 2030, and 2029 in positions 1, 2, and 3, respectively); and two Solid Rocket Boosters (SRB's) designated as BI-043. The primary objective of this flight was to successfully perform the planned operations of the Infrared Background Signature Survey (IBSS), Air Force Payload (AFP)-675, Space Test Payload (STP)-1, and the Multipurpose Experiment Canister (MPEC) payloads.

Modeling, Analysis and Simulation Approaches Used in Development of the National Aeronautics and Space Administration Max Launch Abort System

NASA Technical Reports Server (NTRS)

Yuchnovicz, Daniel E.; Dennehy, Cornelius J.; Schuster, David M.

2011-01-01

The National Aeronautics and Space Administration (NASA) Engineering and Safety Center was chartered to develop an alternate launch abort system (LAS) as risk mitigation for the Orion Project. Its successful flight test provided data for the design of future LAS vehicles. Design of the flight test vehicle (FTV) and pad abort trajectory relied heavily on modeling and simulation including computational fluid dynamics for vehicle aero modeling, 6-degree-of-freedom kinematics models for flight trajectory modeling, and 3-degree-of-freedom kinematics models for parachute force modeling. This paper highlights the simulation techniques and the interaction between the aerodynamics, flight mechanics, and aerodynamic decelerator disciplines during development of the Max Launch Abort System FTV.

STS-39 Space Shuttle mission report

NASA Astrophysics Data System (ADS)

Fricke, Robert W.

1991-06-01

The STS-39 Space Shuttle Program Mission Report contains a summary of the vehicle subsystem operations during the fortieth flight of the Space Shuttle and the twelfth flight of the Orbiter Vehicle Discovery (OV-103). In addition to the Discovery vehicle, the flight vehicle consisted of the following: an External Tank (ET) (designated as ET-46 (LWT-39); three Space Shuttle main engines (SSME's) (serial numbers 2026, 2030, and 2029 in positions 1, 2, and 3, respectively); and two Solid Rocket Boosters (SRB's) designated as BI-043. The primary objective of this flight was to successfully perform the planned operations of the Infrared Background Signature Survey (IBSS), Air Force Payload (AFP)-675, Space Test Payload (STP)-1, and the Multipurpose Experiment Canister (MPEC) payloads.

Analysis of Eye-Tracking Data with Regards to the Complexity of Flight Deck Information Automation and Management - Inattentional Blindness, System State Awareness, and EFB Usage

NASA Technical Reports Server (NTRS)

Dill, Evan T.; Young, Steven D.

2015-01-01

In the constant drive to further the safety and efficiency of air travel, the complexity of avionics-related systems, and the procedures for interacting with these systems, appear to be on an ever-increasing trend. While this growing complexity often yields productive results with respect to system capabilities and flight efficiency, it can place a larger burden on pilots to manage increasing amounts of information and to understand intricate system designs. Evidence supporting this observation is becoming widespread, yet has been largely anecdotal or the result of subjective analysis. One way to gain more insight into this issue is through experimentation using more objective measures or indicators. This study utilizes and analyzes eye-tracking data obtained during a high-fidelity flight simulation study wherein many of the complexities of current flight decks, as well as those planned for the next generation air transportation system (NextGen), were emulated. The following paper presents the findings of this study with a focus on electronic flight bag (EFB) usage, system state awareness (SSA) and events involving suspected inattentional blindness (IB).

High angle of attack control law development for a free-flight wind tunnel model using direct eigenstructure assignment

NASA Technical Reports Server (NTRS)

Wendel, Thomas R.; Boland, Joseph R.; Hahne, David E.

1991-01-01

Flight-control laws are developed for a wind-tunnel aircraft model flying at a high angle of attack by using a synthesis technique called direct eigenstructure assignment. The method employs flight guidelines and control-power constraints to develop the control laws, and gain schedules and nonlinear feedback compensation provide a framework for considering the nonlinear nature of the attack angle. Linear and nonlinear evaluations show that the control laws are effective, a conclusion that is further confirmed by a scale model used for free-flight testing.

Design and development of a fast ion mass spectrometer

NASA Technical Reports Server (NTRS)

Burch, J. L.

1983-01-01

Two Fast Ion Mass Spectrometers (FIMS A and FIMS B) were developed. The design, development, construction, calibration, integration, and flight of these instruments, along with early results from the data analysis efforts are summarized. A medium energy ion mass spectrometer that covers mass velocity space with significantly higher time resolution, improved mass resolution, (particularly for heavier ions), and wider energy range than existing instruments had achieved was completed. The initial design consisted of a dual channel cylindrical electrostatic analyzer followed by a dual channel cylindrical velocity filter. The gain versus count rate characteristics of the high current channel electron multipliers (CEM's), which were chosen for ion detection, revealed a systematic behavior that can be used as a criterion for selection of CEM's for long counting lifetimes.

STS-61 Space Shuttle mission report

NASA Technical Reports Server (NTRS)

Fricke, Robert W., Jr.

1994-01-01

The STS-61 Space Shuttle Program Mission Report summarizes the Hubble Space Telescope (HST) servicing mission as well as the Orbiter, External Tank (ET), Solid Rocket Booster (SRB), Redesigned Solid Rocket Motor (RSRM), and the Space Shuttle main engine (SSME) systems performance during the fifty-ninth flight of the Space Shuttle Program and fifth flight of the Orbiter vehicle Endeavour (OV-105). In addition to the Orbiter, the flight vehicle consisted of an ET designated as ET-60; three SSME's which were designated as serial numbers 2019, 2033, and 2017 in positions 1, 2, and 3, respectively; and two SRB's which were designated BI-063. The RSRM's that were installed in each SRB were designated as 360L023A (lightweight) for the left SRB, and 360L023B (lightweight) for the right SRB. This STS-61 Space Shuttle Program Mission Report fulfills the Space Shuttle Program requirement as documented in NSTS 07700, Volume 8, Appendix E. That document requires that each major organizational element supporting the Program report the results of its hardware evaluation and mission performance plus identify all related in-flight anomalies. The primary objective of the STS-61 mission was to perform the first on-orbit servicing of the Hubble Space Telescope. The servicing tasks included the installation of new solar arrays, replacement of the Wide Field/Planetary Camera I (WF/PC I) with WF/PC II, replacement of the High Speed Photometer (HSP) with the Corrective Optics Space Telescope Axial Replacement (COSTAR), replacement of rate sensing units (RSU's) and electronic control units (ECU's), installation of new magnetic sensing systems and fuse plugs, and the repair of the Goddard High Resolution Spectrometer (GHRS). Secondary objectives were to perform the requirements of the IMAX Cargo Bay Camera (ICBC), the IMAX Camera, and the Air Force Maui Optical Site (AMOS) Calibration Test.

HL-10 pilots assist with pilot entry into lifting body

NASA Technical Reports Server (NTRS)

1969-01-01

Not every moment of a test pilot's day is serious business. In a moment of levity, NASA pilots Bill Dana (left) and John A. Manke try to drag Air Force test pilot Peter Hoag away from the HL-10 lifting body while Air Force Major Jerauld R. Gentry helps from the cockpit. These four men were the principal pilots for the HL-10 program. This was not the only prank involving the HL-10 and its pilots. Once 'Captain Midnight' (Gentry) and the 'Midnight skulkers' sneaked into the NASA hangar and put 'U.S. Air Force' on the aircraft using stick-on letters. Later, while Gentry was making a lifting-body flight, his 1954 Ford was 'borrowed' from the parking lot, painted with yellow-green zinc-chromate primer, and decorated with large stick-on flowers about one foot in diameter. After Gentry returned from the flight, he was surprised to see what had happened to his car. The HL-10 was one of five heavyweight lifting-body designs flown at NASA's Flight Research Center (FRC--later Dryden Flight Research Center), Edwards, California, from July 1966 to November 1975 to study and validate the concept of safely maneuvering and landing a low lift-over-drag vehicle designed for reentry from space. Northrop Corporation built the HL-10 and M2-F2, the first two of the fleet of 'heavy' lifting bodies flown by the NASA Flight Research Center. The contract for construction of the HL-10 and the M2-F2 was $1.8 million. 'HL' stands for horizontal landing, and '10' refers to the tenth design studied by engineers at NASA's Langley Research Center, Hampton, Va. After delivery to NASA in January 1966, the HL-10 made its first flight on Dec. 22, 1966, with research pilot Bruce Peterson in the cockpit. Although an XLR-11 rocket engine was installed in the vehicle, the first 11 drop flights from the B-52 launch aircraft were powerless glide flights to assess handling qualities, stability, and control. In the end, the HL-10 was judged to be the best handling of the three original heavy-weight lifting bodies (M2-F2/F3, HL-10, X-24A). The HL-10 was flown 37 times during the lifting body research program and logged the highest altitude and fastest speed in the Lifting Body program. On Feb. 18, 1970, Air Force test pilot Peter Hoag piloted the HL-10 to Mach 1.86 (1,228 mph). Nine days later, NASA pilot Bill Dana flew the vehicle to 90,030 feet, which became the highest altitude reached in the program. Some new and different lessons were learned through the successful flight testing of the HL-10. These lessons, when combined with information from it's sister ship, the M2-F2/F3, provided an excellent starting point for designers of future entry vehicles, including the Space Shuttle.

X-36 Tailless Fighter Agility Research Aircraft on lakebed during high-speed taxi tests

NASA Technical Reports Server (NTRS)

1996-01-01

The NASA/McDonnell Douglas Corporation (MDC) X-36 Tailless Fighter Agility Research Aircraft undergoes high-speed taxi tests on Rogers Dry Lake at NASA Dryden Flight Research Center, Edwards, California, on October 17, 1996. The aircraft was tested at speeds up to 85 knots. Normal takeoff speed would be 110 knots. More taxi and radio frequency tests were slated before it's first flight would be made. This took place on May 17, 1997. The NASA/Boeing X-36 Tailless Fighter Agility Research Aircraft program successfully demonstrated the tailless fighter design using advanced technologies to improve the maneuverability and survivability of possible future fighter aircraft. The program met or exceeded all project goals. For 31 flights during 1997 at the Dryden Flight Research Center, Edwards, California, the project team examined the aircraft's agility at low speed / high angles of attack and at high speed / low angles of attack. The aircraft's speed envelope reached up to 206 knots (234 mph). This aircraft was very stable and maneuverable. It handled very well. The X-36 vehicle was designed to fly without the traditional tail surfaces common on most aircraft. Instead, a canard forward of the wing was used as well as split ailerons and an advanced thrust-vectoring nozzle for directional control. The X-36 was unstable in both pitch and yaw axes, so an advanced, single-channel digital fly-by-wire control system (developed with some commercially available components) was put in place to stabilize the aircraft. Using a video camera mounted in the nose of the aircraft and an onboard microphone, the X-36 was remotely controlled by a pilot in a ground station virtual cockpit. A standard fighter-type head-up display (HUD) and a moving-map representation of the vehicle's position within the range in which it flew provided excellent situational awareness for the pilot. This pilot-in-the-loop approach eliminated the need for expensive and complex autonomous flight control systems and the risks associated with their inability to deal with unknown or unforeseen phenomena in flight. Fully fueled the X-36 prototype weighed approximately 1,250 pounds. It was 19 feet long and three feet high with a wingspan of just over 10 feet. A Williams International F112 turbofan engine provided close to 700 pounds of thrust. A typical research flight lasted 35 to 45 minutes from takeoff to touchdown. A total of 31 successful research flights were flown from May 17, 1997, to November 12, 1997, amassing 15 hours and 38 minutes of flight time. The aircraft reached an altitude of 20,200 feet and a maximum angle of attack of 40 degrees. In a follow-on effort, the Air Force Research Laboratory (AFRL), Wright-Patterson Air Force Base, Ohio, contracted with Boeing to fly AFRL's Reconfigurable Control for Tailless Fighter Aircraft (RESTORE) software as a demonstration of the adaptability of the neural-net algorithm to compensate for in-flight damage or malfunction of effectors, such as flaps, ailerons and rudders. Two RESTORE research flights were flown in December 1998, proving the viability of the software approach. The X-36 aircraft flown at the Dryden Flight Research Center in 1997 was a 28-percent scale representation of a theoretical advanced fighter aircraft. The Boeing Phantom Works (formerly McDonnell Douglas) in St. Louis, Missouri, built two of the vehicles in a cooperative agreement with the Ames Research Center, Moffett Field, California.

Design and performance test of a MEMS vibratory gyroscope with a novel AGC force rebalance control

NASA Astrophysics Data System (ADS)

Sung, Woon-Tahk; Sung, Sangkyung; Lee, Jang Gyu; Kang, Taesam

2007-10-01

In this paper, the development and performance test results of a laterally oscillating MEMS gyroscope using a novel force rebalance control strategy are presented. The micromachined structure and electrodes are fabricated using the deep reactive ion etching (DRIE) and anodic wafer bonding process. The high quality factor required for the resonance-based sensor is achieved using a vacuum-sealed device package. A systematic design approach of the force rebalance control is applied via a modified automatic gain control (AGC) method. The rebalance control design takes advantages of a novel AGC loop modification, which allows the approximation of the system's dynamics into a simple linear form. Using the proposed modification of AGC and the rebalance strategy that maintains a biased oscillation, a number of performance improvements including bandwidth extension and widened operating range were observed to be achieved. Finally, the experimental results of the gyroscope's practical application verify the feasibility and performance of the developed sensor.

A Wind Tunnel Investigation of Joined Wing Scissor Morphing

DTIC Science & Technology

2006-06-01

would use the low sweep for carrier landing and subsonic cruise, and use the high sweep for 12 supersonic flight [13]. According to Raymer [19...Wright-Patterson AFB, Ohio: Air Force Institute of Technology, 2005. 12. Katz, Joseph, Shaun Byrne, and Robert Hahl. "Stall Resistance Features of...Lifting-Body Airplane Configurations." Journal of Aircraft 2nd ser. 36 (1999): 471-474. 13. Kress, Robert W. "Variable Sweep Wing Design." AIAA 83

Guidance and control/ACEE

NASA Technical Reports Server (NTRS)

1981-01-01

Active controls improve airplane performance by stabilizing its flight, reducing departures from stable flight, and alleviating loads imposed by external forces such as gusts, turbulence, or maneuvers. Some uses for active control systems, the design of redundant and reliable stability augmentation systems, digital fly-by-wire, and NASA assessments of the technology of sensors and actuators are discussed. A series of trade-off studies to better define optimum flight control systems, and research by drone and full-scale models are described.

Intelligent redundant actuation system requirements and preliminary system design

NASA Technical Reports Server (NTRS)

Defeo, P.; Geiger, L. J.; Harris, J.

1985-01-01

Several redundant actuation system configurations were designed and demonstrated to satisfy the stringent operational requirements of advanced flight control systems. However, this has been accomplished largely through brute force hardware redundancy, resulting in significantly increased computational requirements on the flight control computers which perform the failure analysis and reconfiguration management. Modern technology now provides powerful, low-cost microprocessors which are effective in performing failure isolation and configuration management at the local actuator level. One such concept, called an Intelligent Redundant Actuation System (IRAS), significantly reduces the flight control computer requirements and performs the local tasks more comprehensively than previously feasible. The requirements and preliminary design of an experimental laboratory system capable of demonstrating the concept and sufficiently flexible to explore a variety of configurations are discussed.

MEMS Based Micro Aerial Vehicles

NASA Astrophysics Data System (ADS)

Joshi, Niranjan; Köhler, Elof; Enoksson, Peter

2016-10-01

Designing a flapping wing insect robot requires understanding of insect flight mechanisms, wing kinematics and aerodynamic forces. These subsystems are interconnected and their dependence on one another affects the overall performance. Additionally it requires an artificial muscle like actuator and transmission to power the wings. Several kinds of actuators and mechanisms are candidates for this application with their own strengths and weaknesses. This article provides an overview of the insect scaled flight mechanism along with discussion of various methods to achieve the Micro Aerial Vehicle (MAV) flight. Ongoing projects in Chalmers is aimed at developing a low cost and low manufacturing time MAV. The MAV design considerations and design specifications are mentioned. The wings are manufactured using 3D printed carbon fiber and are under experimental study.

VANDENBERG AIR FORCE BASE, CALIF. - The L-1011 carrier aircraft is ready for flight after undergoing a Combined Systems Test, an integrated test involving the Pegasus launch vehicle, SciSat-1 spacecraft and L-1011 aircraft. The SciSat-1 weighs approximately 330 pounds and after launch will be placed in a 400-mile-high polar orbit to investigate processes that control the distribution of ozone in the upper atmosphere. The data from the satellite will provide Canadian and international scientists with improved measurements relating to global ozone processes and help policymakers assess existing environmental policy and develop protective measures for improving the health of our atmosphere, preventing further ozone depletion. The mission is designed to last two years.

NASA Image and Video Library

2003-08-09

VANDENBERG AIR FORCE BASE, CALIF. - The L-1011 carrier aircraft is ready for flight after undergoing a Combined Systems Test, an integrated test involving the Pegasus launch vehicle, SciSat-1 spacecraft and L-1011 aircraft. The SciSat-1 weighs approximately 330 pounds and after launch will be placed in a 400-mile-high polar orbit to investigate processes that control the distribution of ozone in the upper atmosphere. The data from the satellite will provide Canadian and international scientists with improved measurements relating to global ozone processes and help policymakers assess existing environmental policy and develop protective measures for improving the health of our atmosphere, preventing further ozone depletion. The mission is designed to last two years.

Safety Considerations in the Ground Environment

NASA Technical Reports Server (NTRS)

Kirkpatrick, Paul D.; Palo, Thomas E.

2007-01-01

In the history of humankind, every great space adventure has begun on the ground. While this seems to be stating the obvious, mission and spacecraft designers who have overlooked this fact have paid a high price, either in loss or damage to the spacecraft pre-launch, or in mission failure or reduction. Spacecraft personnel may risk not only their flight hardware, but they may also risk their lives, their co-workers lives and even the general public by not heeding safety on the ground. Their eyes may be on the stars but their feet are on the ground! One additional comment: Although the design requirements are very different for human rated and nonhuman rated flight hardware, while on the ground that flight hardware (and its ground support equipment) doesn't care about what it is flying on. On the ground, additional requirements are often levied to protect the work force and general public. (Authors' Note: The source material for this chapter is primarily taken from the Kennedy Space Center Handbook (KHB) 1700.7/45 SW Handbook S-100 Space Shuttle Payload Ground Safety Handbook and the authors' personal experiences.

Rule-based navigation control design for autonomous flight

NASA Astrophysics Data System (ADS)

Contreras, Hugo; Bassi, Danilo

2008-04-01

This article depicts a navigation control system design that is based on a set of rules in order to follow a desired trajectory. The full control of the aircraft considered here comprises: a low level stability control loop, based on classic PID controller and the higher level navigation whose main job is to exercise lateral control (course) and altitude control, trying to follow a desired trajectory. The rules and PID gains were adjusted systematically according to the result of flight simulation. In spite of its simplicity, the rule-based navigation control proved to be robust, even with big perturbation, like crossing winds.

HL-10 on lakebed with Jerauld R. Gentry, Peter Hoag, John A. Manke, and Bill Dana

NASA Technical Reports Server (NTRS)

1969-01-01

The four principal HL-10 pilots are seen here with the lifting body aircraft. They are, left to right; Air Force Major Jerauld R. Gentry, Air Force test pilot Peter Hoag, and NASA pilots John A. Manke and Bill Dana. All are wearing the pressure suits needed for flying above 50,000 feet. The HL-10 was one of five heavyweight lifting-body designs flown at NASA's Flight Research Center (FRC--later Dryden Flight Research Center), Edwards, California, from July 1966 to November 1975 to study and validate the concept of safely maneuvering and landing a low lift-over-drag vehicle designed for reentry from space. Northrop Corporation built the HL-10 and M2-F2, the first two of the fleet of 'heavy' lifting bodies flown by the NASA Flight Research Center. The contract for construction of the HL-10 and the M2-F2 was $1.8 million. 'HL' stands for horizontal landing, and '10' refers to the tenth design studied by engineers at NASA's Langley Research Center, Hampton, Va. After delivery to NASA in January 1966, the HL-10 made its first flight on Dec. 22, 1966, with research pilot Bruce Peterson in the cockpit. Although an XLR-11 rocket engine was installed in the vehicle, the first 11 drop flights from the B-52 launch aircraft were powerless glide flights to assess handling qualities, stability, and control. In the end, the HL-10 was judged to be the best handling of the three original heavy-weight lifting bodies (M2-F2/F3, HL-10, X-24A). The HL-10 was flown 37 times during the lifting body research program and logged the highest altitude and fastest speed in the Lifting Body program. On Feb. 18, 1970, Air Force test pilot Peter Hoag piloted the HL-10 to Mach 1.86 (1,228 mph). Nine days later, NASA pilot Bill Dana flew the vehicle to 90,030 feet, which became the highest altitude reached in the program. Some new and different lessons were learned through the successful flight testing of the HL-10. These lessons, when combined with information from it's sister ship, the M2-F2/F3, provided an excellent starting point for designers of future entry vehicles, including the Space Shuttle.

Encoding properties of haltere neurons enable motion feature detection in a biological gyroscope

PubMed Central

Fox, Jessica L.; Fairhall, Adrienne L.; Daniel, Thomas L.

2010-01-01

The halteres of dipteran insects are essential sensory organs for flight control. They are believed to detect Coriolis and other inertial forces associated with body rotation during flight. Flies use this information for rapid flight control. We show that the primary afferent neurons of the haltere’s mechanoreceptors respond selectively with high temporal precision to multiple stimulus features. Although we are able to identify many stimulus features contributing to the response using principal component analysis, predictive models using only two features, common across the cell population, capture most of the cells’ encoding activity. However, different sensitivity to these two features permits each cell to respond to sinusoidal stimuli with a different preferred phase. This feature similarity, combined with diverse phase encoding, allows the haltere to transmit information at a high rate about numerous inertial forces, including Coriolis forces. PMID:20133721

Motor skills under varied gravitoinertial force in parabolic flight

NASA Astrophysics Data System (ADS)

Ross, Helen E.

Parabolic flight produces brief alternating periods of high and low gravitoinertial force. Subjects were tested on various paper-and-pencil aiming and tapping tasks during both normal and varied gravity in flight. It was found that changes in g level caused directional errors in the z body axis (the gravity axis), the arm aiming too high under 0g and too low under 2g. The standard deviation also increased for both vertical and lateral movements in the mid-frontal plane. Both variable and directional errors were greater under 0g than 2g. In an unpaced reciprocal tapping task subjects tended to increase their error rate rather than their movement time, but showed a non-significant trend towards slower speeds under 0g for all movement orientations. Larger variable errors or slower speeds were probably due to the difficulty of re-organising a motor skill in an unfamiliar force environment, combined with anchorage difficulties under 0g.

X-4 in flight

NASA Technical Reports Server (NTRS)

1951-01-01

In the early days of transonic flight research, many aerodynamicists believed that eliminating conventional tail surfaces could reduce the problems created by shock wave interaction with the tail's lifting surfaces. To address this issue, the Army Air Forces's Air Technical Service awarded a contract to Northrop Aircraft Corporation on 5 April 1946 to build a piloted 'flying laboratory.' Northrop already had experience with tailless flying wing designs such as the N-1M, N-9M, XB-35, and YB-49. Subsequently, the manufacturer built two semi-tailless X-4 research aircraft, the first of which flew half a century ago. The X-4 was designed to investigate transonic compressibility effects at speeds near Mach 0.85 to 0.88, slightly below the speed of sound. Northrop project engineer Arthur Lusk designed the aircraft with swept wings and a conventional fuselage that housed two turbojet engines. It had a vertical stabilizer, but no horizontal tail surfaces. It was one of the smallest X-planes ever built, and every bit of internal space was used for systems and instrumentation. The first X-4 arrived at Muroc Air Force Base by truck on 15 November 1948. Over the course of several weeks, engineers conducted static tests, and Northrop test pilot Charles Tucker made initial taxi runs. Although small of stature, he barely fit into the diminutive craft. Tucker, a veteran Northrop test pilot, had previously flown the XB-35 and YB-49 flying wing bomber prototypes. Prior to flying for Northrop, he had logged 400 hours in jet airplanes as a test pilot for Lockheed and the Air Force. He would now be responsible for completing the contractor phase of the X-4 flight test program. Finally, all was ready. Tucker climbed into the cockpit, and made the first flight on 15 December 1948. It only lasted 18 minutes, allowing just enough time for the pilot to become familiar with the basic handling qualities of the craft. The X-4 handled well, but Tucker noted some longitudinal instability at all speeds. Data generated during the initial flight led to design changes that improved handling and performance. While the first aircraft underwent modifications, the second X-4 arrived at Muroc in early 1949. Heavy rains flooded the dry lakebed at Muroc, delaying further flight testing until 27 April. The first X-4 made a few more flights, but was beset by mechanical problems. Soon, the second aircraft became the workhorse of the contractor test program. Its first flight was accomplished on 7 June 1949. It was more completely instrumented than its stablemate, and didn't suffer from the same plague of malfunctions. The National Advisory Committee for Aeronautics (NACA) expressed interest in using the second X-4 for research, but was frustrated by the slow pace of the contractor's test phase. Both X-4 aircraft were grounded temporarily for installation of spin recovery parachutes, and improvements to the landing gear uplock system. After these tasks were completed, the first X-4 made its tenth and final flight on 26 January 1950. It was grounded, and used as a source of spare parts for the second aircraft. On 17 February, the remaining X-4 completed the contractor testing phase. The aircraft was then turned over to the NACA and the Air Force for a joint research program. NACA technicians prepared the aircraft, and made a number of design improvements. The joint NACA/USAF program consisted of 82 flights to evaluate handling qualities, stability and control, and performance at various lift-to-drag ratios. NACA pilots during the program included Stanley Butchart, George Cooper, Scott Crossfield, John Griffith, Walter Jones, John 'Jack' McKay, and Joe Walker. The Air Force pilots included B/Gen. Albert Boyd, Col. Frank 'Pete' Everest, Lt. Col. Richard Johnson, Capt. J. S. Nash, and Maj. Charles 'Chuck' Yeager. The final X-4 flight took place in September 1953. Further flights were planned, but a chronic fuel leak lead to cancellation of the program. NACA engineers had acquired a wealth of data, and the cost of repairing the leak could not be justified. Although the X-4 was never designed to fly at supersonic speeds, it gathered transonic data that proved that conventional tailless swept-wing configurations were unsuitable for supersonic performance. The design suffered from noticeable instability in all directions, increasing as it approached the speed of sound. It was, however, a valuable tool for dynamic stability research. Additionally low lift-to-drag data gathered with the X-4 was later beneficial to development of the X-15. On 10 March 1954, both X-4 aircraft were returned to the Air Force. The ejection seat from the second X-4 was retained by the NACA for use in the X-1E. The first X-4 is currently on display at the United States Air Force Academy in Colorado. The second aircraft is displayed at the USAF Museum in Ohio. This movie clip running approximately 7 seconds shows the Northrop X-4 in an air-to-air view climbing away from its chase aircraft.

United States Air Force Research Initiation Program for 1988. Volume 3

DTIC Science & Technology

1990-04-01

Assignment for Dr. Kenneth M. Sobel Flight Control Design 210-9MG-035 90 Comparative Burning Rates and Duplex Dr. Forrest Thomas (1987) Loads of Solid...Patterson Air Force Base. The test configuration has been designed for injecting fuel droplets in a well controlled laminar on well-characterized turbulent...its counter response may be significant, our system has thus achieved some measure of control over when non -critical processing is actually performed

X-36 Taking off during First Flight

NASA Technical Reports Server (NTRS)

1997-01-01

The remotely-piloted X-36 Tailless Fighter Agility Research Aircraft lifts off from Rogers Dry Lake at the Dryden Flight Research Center on its first flight on May 17, 1997. The NASA/Boeing X-36 Tailless Fighter Agility Research Aircraft program successfully demonstrated the tailless fighter design using advanced technologies to improve the maneuverability and survivability of possible future fighter aircraft. The program met or exceeded all project goals. For 31 flights during 1997 at the Dryden Flight Research Center, Edwards, California, the project team examined the aircraft's agility at low speed / high angles of attack and at high speed / low angles of attack. The aircraft's speed envelope reached up to 206 knots (234 mph). This aircraft was very stable and maneuverable. It handled very well. The X-36 vehicle was designed to fly without the traditional tail surfaces common on most aircraft. Instead, a canard forward of the wing was used as well as split ailerons and an advanced thrust-vectoring nozzle for directional control. The X-36 was unstable in both pitch and yaw axes, so an advanced, single-channel digital fly-by-wire control system (developed with some commercially available components) was put in place to stabilize the aircraft. Using a video camera mounted in the nose of the aircraft and an onboard microphone, the X-36 was remotely controlled by a pilot in a ground station virtual cockpit. A standard fighter-type head-up display (HUD) and a moving-map representation of the vehicle's position within the range in which it flew provided excellent situational awareness for the pilot. This pilot-in-the-loop approach eliminated the need for expensive and complex autonomous flight control systems and the risks associated with their inability to deal with unknown or unforeseen phenomena in flight. Fully fueled the X-36 prototype weighed approximately 1,250 pounds. It was 19 feet long and three feet high with a wingspan of just over 10 feet. A Williams International F112 turbofan engine provided close to 700 pounds of thrust. A typical research flight lasted 35 to 45 minutes from takeoff to touchdown. A total of 31 successful research flights were flown from May 17, 1997, to November 12, 1997, amassing 15 hours and 38 minutes of flight time. The aircraft reached an altitude of 20,200 feet and a maximum angle of attack of 40 degrees. In a follow-on effort, the Air Force Research Laboratory (AFRL), Wright-Patterson Air Force Base, Ohio, contracted with Boeing to fly AFRL's Reconfigurable Control for Tailless Fighter Aircraft (RESTORE) software as a demonstration of the adaptability of the neural-net algorithm to compensate for in-flight damage or malfunction of effectors, such as flaps, ailerons and rudders. Two RESTORE research flights were flown in December 1998, proving the viability of the software approach. The X-36 aircraft flown at the Dryden Flight Research Center in 1997 was a 28-percent scale representation of a theoretical advanced fighter aircraft. The Boeing Phantom Works (formerly McDonnell Douglas) in St. Louis, Missouri, built two of the vehicles in a cooperative agreement with the Ames Research Center, Moffett Field, California.

The Effect of Systematic Error in Forced Oscillation Testing

NASA Technical Reports Server (NTRS)

Williams, Brianne Y.; Landman, Drew; Flory, Isaac L., IV; Murphy, Patrick C.

2012-01-01

One of the fundamental problems in flight dynamics is the formulation of aerodynamic forces and moments acting on an aircraft in arbitrary motion. Classically, conventional stability derivatives are used for the representation of aerodynamic loads in the aircraft equations of motion. However, for modern aircraft with highly nonlinear and unsteady aerodynamic characteristics undergoing maneuvers at high angle of attack and/or angular rates the conventional stability derivative model is no longer valid. Attempts to formulate aerodynamic model equations with unsteady terms are based on several different wind tunnel techniques: for example, captive, wind tunnel single degree-of-freedom, and wind tunnel free-flying techniques. One of the most common techniques is forced oscillation testing. However, the forced oscillation testing method does not address the systematic and systematic correlation errors from the test apparatus that cause inconsistencies in the measured oscillatory stability derivatives. The primary objective of this study is to identify the possible sources and magnitude of systematic error in representative dynamic test apparatuses. Sensitivities of the longitudinal stability derivatives to systematic errors are computed, using a high fidelity simulation of a forced oscillation test rig, and assessed using both Design of Experiments and Monte Carlo methods.

Wave Amplitude Dependent Engineering Model of Propellant Slosh in Spherical Tanks

NASA Technical Reports Server (NTRS)

Brodnick, Jacob; Westra, Douglas G.; Eberhart, Chad J.; Yang, Hong Q.; West, Jeffrey S.

2016-01-01

Liquid propellant slosh is often a concern for the controllability of flight vehicles. Anti-slosh devices are traditionally included in propellant tank designs to limit the amount of sloshing allowed during flight. These devices and any necessary supports can be quite heavy to meet various structural requirements. Some of the burden on anti-slosh devices can be relieved by exploiting the nonlinear behavior of slosh waves in bare smooth wall tanks. A nonlinear regime slosh model for bare spherical tanks was developed through a joint analytical and experimental effort by NASA/MSFC. The developed slosh model accounts for the large damping inherent in nonlinear slosh waves which is more accurate and drives conservatism from vehicle stability analyses that use traditional bare tank slosh models. A more accurate slosh model will result in more realistic predicted slosh forces during flight reducing or removing the need for active controls during a maneuver or baffles in the tank design. Lower control gains and smaller or fewer tank baffles can reduce cost and system complexity while increasing vehicle performance. Both Computational Fluid Dynamics (CFD) simulation and slosh testing of three different spherical tank geometries were performed to develop the proposed slosh model. Several important findings were made during this effort in addition to determining the parameters to the nonlinear regime slosh model. The linear regime slosh damping trend for spherical tanks reported in NASA SP-106 was shown to be inaccurate for certain regions of a tank. Additionally, transition to the nonlinear regime for spherical tanks was only found to occur at very large wave amplitudes in the lower hemisphere and was a strong function of the propellant fill level in the upper hemisphere. The nonlinear regime damping trend was also found to be a function of the propellant fill level.

The roller coaster flight strategy of bar-headed geese conserves energy during Himalayan migrations

USGS Publications Warehouse

Bishop, C.M.; Spivey, R.J.; Hawkes, L. A.; Batbayar, N.; Chua, B.; Frappell, P.B.; Milsom, W.K.; Natsagdorj, T.; Newman, S.H.; Scott, G.R.; Takekawa, John Y.; Wikelski, Martin; Butler, Patrick J.

2015-01-01

The physiological and biomechanical requirements of flight at high altitude have been the subject of much interest. Here, we uncover a steep relation between heart rate and wingbeat frequency (raised to the exponent 3.5) and estimated metabolic power and wingbeat frequency (exponent 7) of migratory bar-headed geese. Flight costs increase more rapidly than anticipated as air density declines, which overturns prevailing expectations that this species should maintain high-altitude flight when traversing the Himalayas. Instead, a "roller coaster" strategy, of tracking the underlying terrain and discarding large altitude gains only to recoup them later in the flight with occasional benefits from orographic lift, is shown to be energetically advantageous for flights over the Himalayas.

MSFC Skylab thermal and environmental control system mission evaluation

NASA Technical Reports Server (NTRS)

Hopson, G. D.; Littles, J. W.; Patterson, W. C.

1974-01-01

An evaluation of the performance of the Skylab thermal and environmental control system is presented. Actual performance is compared to design and functional requirements and anomalies and discrepancies and their resolution are discussed. The thermal and environmental control systems performed their intended role. Based on the experience gained in design, development and flight, recommendations are provided which may be beneficial to future system designs.

Diagram of a Hydrogen Fuel System on NACA’s Martin B-57B Canberra

NASA Image and Video Library

1957-02-21

This diagram shows a hydrogen fuel system designed by researchers at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory and installed on a Martin B-57B Canberra aircraft. Lewis researchers accelerated their studies of high energy propellants in the early 1950s. In late 1954, Lewis researchers studied the combustion characteristics of gaseous hydrogen in a turbojet combustor. It was found that the hydrogen provided a very high efficiency. Almost immediately thereafter, Associate Director Abe Silverstein became focused on the possibilities of hydrogen for aircraft propulsion. That fall, Silverstein secured a contract to work with the air force to examine the practicality of liquid hydrogen aircraft. A B-57B Canberra was obtained by the air force especially for this project, referred to as Project Bee. The aircraft was powered by two Wright J65 engines, one of which was modified so that it could be operated using either traditional or liquid hydrogen propellants. The engine and its liquid hydrogen fuel system were tested extensively in the Altitude Wind Tunnel and the Four Burner Area test cells in 1955 and 1956. A B-57B flight program was planned to test the system on an actual aircraft. The aircraft would take off using jet fuel, switch to liquid hydrogen while over Lake Erie, then after burning the hydrogen supply switch back to jet fuel for the landing. The third test flight, in February 1957, was a success, and the ensuing B-57B flights remain the only demonstration of hydrogen-powered aircraft.

32 CFR 245.21 - ESCAT air traffic priority list.

Code of Federal Regulations, 2013 CFR

2013-07-01

... aircraft post-maintenance test flights. (7) Federal aircraft post maintenance check flights in support of... identified as high threat targets. (c) Priority Three. (1) Forces being deployed or performing pre-deployment...

32 CFR 245.21 - ESCAT air traffic priority list.

Code of Federal Regulations, 2010 CFR

2010-07-01

... aircraft post-maintenance test flights. (7) Federal aircraft post maintenance check flights in support of... identified as high threat targets. (c) Priority Three. (1) Forces being deployed or performing pre-deployment...

32 CFR 245.21 - ESCAT air traffic priority list.

Code of Federal Regulations, 2014 CFR

2014-07-01

... aircraft post-maintenance test flights. (7) Federal aircraft post maintenance check flights in support of... identified as high threat targets. (c) Priority Three. (1) Forces being deployed or performing pre-deployment...

32 CFR 245.21 - ESCAT air traffic priority list.

Code of Federal Regulations, 2011 CFR

2011-07-01

... aircraft post-maintenance test flights. (7) Federal aircraft post maintenance check flights in support of... identified as high threat targets. (c) Priority Three. (1) Forces being deployed or performing pre-deployment...

KENNEDY SPACE CENTER, FLA. - Workers at Cape Canaveral Air Force Station place one of four rudder speed brake actuators onto a pallet for X-ray. The actuators, to be installed on the orbiter Discovery, are being X-rayed at the Radiographic High-Energy X-ray Facility to determine if the gears were installed correctly. Discovery has been assigned to the first Return to Flight mission, STS-114, a logistics flight to the International Space Station.

NASA Image and Video Library

2004-03-08

KENNEDY SPACE CENTER, FLA. - Workers at Cape Canaveral Air Force Station place one of four rudder speed brake actuators onto a pallet for X-ray. The actuators, to be installed on the orbiter Discovery, are being X-rayed at the Radiographic High-Energy X-ray Facility to determine if the gears were installed correctly. Discovery has been assigned to the first Return to Flight mission, STS-114, a logistics flight to the International Space Station.

How differential deflection of the inboard and outboard leading-edge flaps affected the handling qua

NASA Technical Reports Server (NTRS)

2002-01-01

How differential deflection of the inboard and outboard leading-edge flaps affected the handling qualities of this modified F/A-18A was evaluated during the first check flight in the Active Aeroelastic Wing program at NASA's Dryden Flight Research Center. The Active Aeroelastic Wing program at NASA's Dryden Flight Research Center seeks to determine the advantages of twisting flexible wings for primary maneuvering roll control at transonic and supersonic speeds, with traditional control surfaces such as ailerons and leading-edge flaps used to aerodynamically induce the twist. From flight test and simulation data, the program intends to develop structural modeling techniques and tools to help design lighter, more flexible high aspect-ratio wings for future high-performance aircraft, which could translate to more economical operation or greater payload capability. AAW flight tests began in November, 2002 with checkout and parameter-identification flights. Based on data obtained during the first flight series, new flight control software will be developed and a second series of research flights will then evaluate the AAW concept in a real-world environment. The program uses wings that were modified to the flexibility of the original pre-production F-18 wing. Other modifications include a new actuator to operate the outboard leading edge flap over a greater range and rate, and a research flight control system to host the aeroelastic wing control laws. The Active Aeroelastic Wing Program is jointly funded and managed by the Air Force Research Laboratory and NASA Dryden Flight Research Center, with Boeing's Phantom Works as prime contractor for wing modifications and flight control software development. The F/A-18A aircraft was provided by the Naval Aviation Systems Test Team and modified for its research role by NASA Dryden technicians.

KSC-2014-3158

NASA Image and Video Library

2014-06-30

CAPE CANAVERAL, Fla. – Inside the Delta Operations Center near Space Launch Complex 37 on Cape Canaveral Air Force Station in Florida, the second stage for the United Launch Alliance Delta IV Heavy for Exploration Flight Test-1, or EFT-1, is lifted high by crane for the move to a test cell. At the Horizontal Integration Facility, all three booster stages will be processed and checked out before being moved to the nearby launch pad and hoisted into position. During the EFT-1 mission, Orion will travel farther into space than any human spacecraft has gone in more than 40 years. The data gathered during the flight will influence design decisions, validate existing computer models and innovative new approaches to space systems development, as well as reduce overall mission risks and costs for later Orion flights. Liftoff of Orion on EFT-1 is planned for fall 2014. Photo credit: NASA/Frankie Martin

KSC-2014-3159

NASA Image and Video Library

2014-06-30

CAPE CANAVERAL, Fla. – Inside the Delta Operations Center near Space Launch Complex 37 on Cape Canaveral Air Force Station in Florida, the second stage for the United Launch Alliance Delta IV Heavy for Exploration Flight Test-1, or EFT-1, is lifted high by crane into a test cell. At the Horizontal Integration Facility, all three booster stages will be processed and checked out before being moved to the nearby launch pad and hoisted into position. During the EFT-1 mission, Orion will travel farther into space than any human spacecraft has gone in more than 40 years. The data gathered during the flight will influence design decisions, validate existing computer models and innovative new approaches to space systems development, as well as reduce overall mission risks and costs for later Orion flights. Liftoff of Orion on EFT-1 is planned for fall 2014. Photo credit: NASA/Frankie Martin

The Next Generation of High-Speed Dynamic Stability Wind Tunnel Testing (Invited)

NASA Technical Reports Server (NTRS)

Tomek, Deborah M.; Sewall, William G.; Mason, Stan E.; Szchur, Bill W. A.

2006-01-01

Throughout industry, accurate measurement and modeling of dynamic derivative data at high-speed conditions has been an ongoing challenge. The expansion of flight envelopes and non-conventional vehicle design has greatly increased the demand for accurate prediction and modeling of vehicle dynamic behavior. With these issues in mind, NASA Langley Research Center (LaRC) embarked on the development and shakedown of a high-speed dynamic stability test technique that addresses the longstanding problem of accurately measuring dynamic derivatives outside the low-speed regime. The new test technique was built upon legacy technology, replacing an antiquated forced oscillation system, and greatly expanding the capabilities beyond classic forced oscillation testing at both low and high speeds. The modern system is capable of providing a snapshot of dynamic behavior over a periodic cycle for varying frequencies, not just a damping derivative term at a single frequency.

X-36 on Ramp Viewed from Above

NASA Technical Reports Server (NTRS)

1997-01-01

This look-down view of the X-36 Tailless Fighter Agility Research Aircraft on the ramp at NASA's Dryden Flight Research Center, Edwards, California, clearly shows the unusual wing and canard design of the remotely-piloted aircraft. The NASA/Boeing X-36 Tailless Fighter Agility Research Aircraft program successfully demonstrated the tailless fighter design using advanced technologies to improve the maneuverability and survivability of possible future fighter aircraft. The program met or exceeded all project goals. For 31 flights during 1997 at the Dryden Flight Research Center, Edwards, California, the project team examined the aircraft's agility at low speed / high angles of attack and at high speed / low angles of attack. The aircraft's speed envelope reached up to 206 knots (234 mph). This aircraft was very stable and maneuverable. It handled very well. The X-36 vehicle was designed to fly without the traditional tail surfaces common on most aircraft. Instead, a canard forward of the wing was used as well as split ailerons and an advanced thrust-vectoring nozzle for directional control. The X-36 was unstable in both pitch and yaw axes, so an advanced, single-channel digital fly-by-wire control system (developed with some commercially available components) was put in place to stabilize the aircraft. Using a video camera mounted in the nose of the aircraft and an onboard microphone, the X-36 was remotely controlled by a pilot in a ground station virtual cockpit. A standard fighter-type head-up display (HUD) and a moving-map representation of the vehicle's position within the range in which it flew provided excellent situational awareness for the pilot. This pilot-in-the-loop approach eliminated the need for expensive and complex autonomous flight control systems and the risks associated with their inability to deal with unknown or unforeseen phenomena in flight. Fully fueled the X-36 prototype weighed approximately 1,250 pounds. It was 19 feet long and three feet high with a wingspan of just over 10 feet. A Williams International F112 turbofan engine provided close to 700 pounds of thrust. A typical research flight lasted 35 to 45 minutes from takeoff to touchdown. A total of 31 successful research flights were flown from May 17, 1997, to November 12, 1997, amassing 15 hours and 38 minutes of flight time. The aircraft reached an altitude of 20,200 feet and a maximum angle of attack of 40 degrees. In a follow-on effort, the Air Force Research Laboratory (AFRL), Wright-Patterson Air Force Base, Ohio, contracted with Boeing to fly AFRL's Reconfigurable Control for Tailless Fighter Aircraft (RESTORE) software as a demonstration of the adaptability of the neural-net algorithm to compensate for in-flight damage or malfunction of effectors, such as flaps, ailerons and rudders. Two RESTORE research flights were flown in December 1998, proving the viability of the software approach. The X-36 aircraft flown at the Dryden Flight Research Center in 1997 was a 28-percent scale representation of a theoretical advanced fighter aircraft. The Boeing Phantom Works (formerly McDonnell Douglas) in St. Louis, Missouri, built two of the vehicles in a cooperative agreement with the Ames Research Center, Moffett Field, California.

Modeling and Control of a Fixed Wing Tilt-Rotor Tri-Copter

NASA Astrophysics Data System (ADS)

Summers, Alexander

The following thesis considers modeling and control of a fixed wing tilt-rotor tri-copter. An emphasis of the conceptual design is made toward payload transport. Aerodynamic panel code and CAD design provide the base aerodynamic, geometric, mass, and inertia properties. A set of non-linear dynamics are created considering gravity, aerodynamics in vertical takeoff and landing (VTOL) and forward flight, and propulsion applied to a three degree of freedom system. A transition strategy, that removes trajectory planning by means of scheduled inputs, is theorized. Three discrete controllers, utilizing separate control techniques, are applied to ensure stability in the aerodynamic regions of VTOL, transition, and forward flight. The controller techniques include linear quadratic regulation, full state integral action, gain scheduling, and proportional integral derivative (PID) flight control. Simulation of the model control system for flight from forward to backward transition is completed with mass and center of gravity variation.

M2-F1 in flight

NASA Technical Reports Server (NTRS)

1964-01-01

The M2-F1 Lifting Body is seen here under tow by an unseen C-47 at the NASA Flight Research Center (later redesignated the Dryden Flight Research Center), Edwards, California. The low-cost vehicle was the first piloted lifting body to be test flown. The lifting-body concept originated in the mid-1950s at the National Advisory Committee for Aeronautics' Ames Aeronautical Laboratory, Mountain View California. By February 1962, a series of possible shapes had been developed, and R. Dale Reed was working to gain support for a research vehicle. The wingless, lifting body aircraft design was initially concieved as a means of landing an aircraft horizontally after atmospheric reentry. The absence of wings would make the extreme heat of re-entry less damaging to the vehicle. In 1962, Dryden management approved a program to build a lightweight, unpowered lifting body as a prototype to flight test the wingless concept. It would look like a 'flying bathtub,' and was designated the M2-F1, the 'M' referring to 'manned' and 'F' referring to 'flight' version. It featured a plywood shell placed over a tubular steel frame crafted at Dryden. Construction was completed in 1963. The first flight tests of the M2-F1 were over Rogers Dry Lake at the end of a tow rope attached to a hopped-up Pontiac convertible driven at speeds up to about 120 mph. These initial tests produced enough flight data about the M2-F1 to proceed with flights behind a NASA C-47 tow plane at greater altitudes. The C-47 took the craft to an altitude of 12,000 where free flights back to Rogers Dry Lake began. Pilot for the first series of flights of the M2-F1 was NASA research pilot Milt Thompson. Typical glide flights with the M2-F1 lasted about two minutes and reached speeds of 110 to l20 mph. More than 400 ground tows and 77 aircraft tow flights were carried out with the M2-F1. The success of Dryden's M2-F1 program led to NASA's development and construction of two heavyweight lifting bodies based on studies at NASA's Ames and Langley research centers--the M2-F2 and the HL-10, both built by the Northrop Corporation, and the U.S. Air Force's X-24 program. The Lifting Body program also heavily influenced the Space Shuttle program. The M2-F1 program demonstrated the feasibility of the lifting-body concept for horizontal landings of atmospheric entry vehicles. It also demonstrated a procurement and management concept for prototype flight research vehicles that produced rapid results at very low cost (approximately $50,000, excluding salaries of government employees assigned to the project).

Concept report: Experimental vector magnetograph (EXVM) operational configuration balloon flight assembly

NASA Technical Reports Server (NTRS)

1993-01-01

The observational limitations of earth bound solar studies has prompted a great deal of interest in recent months in being able to gain new scientific perspectives through, what should prove to be, relatively low cost flight of the magnetograph system. The ground work done by TBE for the solar balloon missions (originally planned for SOUP and GRID) as well as the rather advanced state of assembly of the EXVM has allowed the quick formulation of a mission concept for the 30 cm system currently being assembled. The flight system operational configuration will be discussed as it is proposed for short duration flight (on the order of one day) over the continental United States. Balloon hardware design requirements used in formulation of the concept are those set by the National Science Balloon Facility (NSBF), the support agency under NASA contract for flight services. The concept assumes that the flight hardware assembly would come together from three development sources: the scientific investigator package, the integration contractor package, and the NSBF support system. The majority of these three separate packages can be independently developed; however, the computer control interfaces and telemetry links would require extensive preplanning and coordination. A special section of this study deals with definition of a dedicated telemetry link to be provided by the integration contractor for video image data for pointing system performance verification. In this study the approach has been to capitalize to the maximum extent possible on existing hardware and system design. This is the most prudent step that can be taken to reduce eventual program cost for long duration flights. By fielding the existing EXVM as quickly as possible, experience could be gained from several short duration flight tests before it became necessary to commit to major upgrades for long duration flights of this system or of the larger 60 cm version being considered for eventual development.

Cockpit integration from a pilot's point of view

NASA Technical Reports Server (NTRS)

Green, D. L.

1982-01-01

Extensive experience in both operational and engineering test flight was used to suggest straightforward changes to helicopter cockpit and control system design that would improve pilot performance in marginal and instrument flight conditions. Needed control system improvements considered include: (1) separation of yaw from cyclic force trim; (2) pedal force proportional to displacement rate; and (3) integration of engine controls in collective stick. Display improvements needed include: (1) natural cuing of yaw rate in attitude indicator; (2) collective position indication and radar altimeter placed within primary scan; and (3) omnidirectional display of full range airspeed data.

Bone histological correlates of soaring and high-frequency flapping flight in the furculae of birds.

PubMed

Mitchell, Jessica; Legendre, Lucas J; Lefèvre, Christine; Cubo, Jorge

2017-06-01

The furcula is a specialized bone in birds involved in flight function. Its morphology has been shown to reflect different flight styles from soaring/gliding birds, subaqueous flight to high-frequency flapping flyers. The strain experienced by furculae can vary depending on flight type. Bone remodeling is a response to damage incurred from different strain magnitudes and types. In this study, we tested whether a bone microstructural feature, namely Haversian bone density, differs in birds with different flight styles, and reassessed previous work using phylogenetic comparative methods that assume an evolutionary model with additional taxa. We show that soaring birds have higher Haversian bone densities than birds with a flapping style of flight. This result is probably linked to the fact that the furculae of soaring birds provide less protraction force and more depression force than furculae of birds showing other kinds of flight. The whole bone area is another explanatory factor, which confirms the fact that size is an important consideration in Haversian bone development. All birds, however, display Haversian bone development in their furculae, and other factors like age could be affecting the response of Haversian bone development. Copyright © 2017 Elsevier GmbH. All rights reserved.

Evolution of Space Shuttle Range Safety (RS) Ascent Flight Envelope Design

NASA Technical Reports Server (NTRS)

Brewer, Joan D.

2011-01-01

Ascent flight envelopes are trajectories that define the normal operating region of a space vehicle s position from liftoff until the end of powered flight. They fulfill part of the RS data requirements imposed by the Air Force s 45th Space Wing (45SW) on space vehicles launching from the Eastern Range (ER) in Florida. The 45SW is chartered to protect the public by minimizing risks associated with the inherent hazards of launching a vehicle into space. NASA s Space Shuttle program has launched 130+ manned missions over a 30 year period from the ER. Ascent envelopes were delivered for each of those missions. The 45SW envelope requirements have remained largely unchanged during this time. However, the methodology and design processes used to generate the envelopes have evolved over the years to support mission changes, maintain high data quality, and reduce costs. The evolution of the Shuttle envelope design has yielded lessons learned that can be applied to future endevours. There have been numerous Shuttle ascent design enhancements over the years that have caused the envelope methodology to evolve. One of these Shuttle improvements was the introduction of onboard flight software changes implemented to improve launch probability. This change impacted the preflight nominal ascent trajectory, which is a key element in the RS envelope design. While the early Shuttle nominal trajectories were designed preflight using a representative monthly mean wind, the new software changes involved designing a nominal ascent trajectory on launch day using real-time winds. Because the actual nominal trajectory position was not known until launch day, the envelope analysis had to be customized to account for this nominal trajectory variation in addition to the other envelope components.

Investigation of interactions between limb-manipulator dynamics and effective vehicle roll control characteristics

NASA Technical Reports Server (NTRS)

Johnston, D. E.; Mcruer, D. T.

1986-01-01

A fixed-base simulation was performed to identify and quantify interactions between the pilot's hand/arm neuromuscular subsystem and such features of typical modern fighter aircraft roll rate command control system mechanization as: (1) force sensing side-stick type manipulator; (2) vehicle effective role time constant; and (3) flight control system effective time delay. The simulation results provide insight to high frequency pilot induced oscillations (PIO) (roll ratchet), low frequency PIO, and roll-to-right control and handling problems previously observed in experimental and production fly-by-wire control systems. The simulation configurations encompass and/or duplicate actual flight situations, reproduce control problems observed in flight, and validate the concept that the high frequency nuisance mode known as roll ratchet derives primarily from the pilot's neuromuscular subsystem. The simulations show that force-sensing side-stick manipulator force/displacement/command gradients, command prefilters, and flight control system time delays need to be carefully adjusted to minimize neuromuscular mode amplitude peaking (roll ratchet tendency) without restricting roll control bandwidth (with resulting sluggish or PIO prone control).

Digital controllers for VTOL aircraft

NASA Technical Reports Server (NTRS)

Stengel, R. F.; Broussard, J. R.; Berry, P. W.

1976-01-01

Using linear-optimal estimation and control techniques, digital-adaptive control laws have been designed for a tandem-rotor helicopter which is equipped for fully automatic flight in terminal area operations. Two distinct discrete-time control laws are designed to interface with velocity-command and attitude-command guidance logic, and each incorporates proportional-integral compensation for non-zero-set-point regulation, as well as reduced-order Kalman filters for sensor blending and noise rejection. Adaptation to flight condition is achieved with a novel gain-scheduling method based on correlation and regression analysis. The linear-optimal design approach is found to be a valuable tool in the development of practical multivariable control laws for vehicles which evidence significant coupling and insufficient natural stability.

Users manual for flight control design programs

NASA Technical Reports Server (NTRS)

Nalbandian, J. Y.

1975-01-01

Computer programs for the design of analog and digital flight control systems are documented. The program DIGADAPT uses linear-quadratic-gaussian synthesis algorithms in the design of command response controllers and state estimators, and it applies covariance propagation analysis to the selection of sampling intervals for digital systems. Program SCHED executes correlation and regression analyses for the development of gain and trim schedules to be used in open-loop explicit-adaptive control laws. A linear-time-varying simulation of aircraft motions is provided by the program TVHIS, which includes guidance and control logic, as well as models for control actuator dynamics. The programs are coded in FORTRAN and are compiled and executed on both IBM and CDC computers.

Flight-Proven Nano-Satellite Architecture for Hands-On Academic Training at the US Air Force Academy

NASA Astrophysics Data System (ADS)

Underwood, Craig I.; Sellers, Lt. Jerry, , Col.; Sweeting, Martin, , Sir

2002-01-01

This paper describes the use of "commercial-off-the-shelf" open-architecture satellite sub-systems, based on the flight- proven "SNAP" nanosatellite platform, to provide "hands-on" education and training at the United States Air Force Academy. The UK's first nanosatellite: SNAP-1, designed and built by Surrey Satellite Technology Ltd. (SSTL) and Surrey Space Centre staff - in less than a year - was launched in June 2000. The 6.5 kg spacecraft carries advanced, UK-developed, GPS navigation, computing, propulsion and attitude control technologies, which have been used to demonstrate orbital manoeuvring and full three-axis controlled body stabilisation. SNAP-1's primary payload is a machine vision system which has been used to image the in-orbit deployment of another SSTL-built spacecraft: Tsinghua-1. The highly successful, SNAP-1 mission has also demonstrated how the concept of using a standardised, modular nanosatellite bus can provide the core support units (power system, on-board data-handling and communications systems and standardised payload interface) for a practical nanosatellite to be constructed and flown in a remarkably short time-frame. Surrey's undergraduate and post-graduate students have made a major input to the SNAP concept over the last six years in the context of project work within the Space Centre. Currently, students at the USAF Academy are benefiting from this technology in the context of designing their own nanosatellite - FalconSAT-2. For the FalconSAT-2 project, the approach has been to focus on building up infrastructure, including design and development tools that can serve as a firm foundation to allow the satellite design to evolve steadily over the course of several missions. Specific to this new approach has been a major effort to bound the problem faced by the students. To do this, the program has leveraged the research carried out at the Surrey Space Centre, by "buying into" the SNAP architecture. Through this, the Academy program has achieved an "out of the box" solution for several critical subsystems; including power, communications and, most important, data handling. Using one set of SNAP hardware, the FalconSAT Avionics Simulation Testbed (FAST) was established in Fall 2000. FAST provides both a long-term facility for cadets to gain hands-on experience with spacecraft hardware and software, as well as overall program risk reduction by providing a facility for subsystem, software, and operational procedures development and testing. In addition, over the last two years, USAF cadets have been seconded to Surrey to help develop a MATLAB-based spacecraft simulator for SNAP, which itself is becoming a useful educational tool. While the use of the SNAP hardware has eased spacecraft design problem in many respects, considerable effort still remains in the areas of payload design and development, structures, attitude control, thermal control, solar panels, testing and operations -- more than enough to challenge even the most ambitious undergraduate students. This paper reviews our experience, both in the UK and in the US, in using a flight-proven nanosatellite in an educational context.

Simulation of Attitude and Trajectory Dynamics and Control of Multiple Spacecraft

NASA Technical Reports Server (NTRS)

Stoneking, Eric T.

2009-01-01

Agora software is a simulation of spacecraft attitude and orbit dynamics. It supports spacecraft models composed of multiple rigid bodies or flexible structural models. Agora simulates multiple spacecraft simultaneously, supporting rendezvous, proximity operations, and precision formation flying studies. The Agora environment includes ephemerides for all planets and major moons in the solar system, supporting design studies for deep space as well as geocentric missions. The environment also contains standard models for gravity, atmospheric density, and magnetic fields. Disturbance force and torque models include aerodynamic, gravity-gradient, solar radiation pressure, and third-body gravitation. In addition to the dynamic and environmental models, Agora supports geometrical visualization through an OpenGL interface. Prototype models are provided for common sensors, actuators, and control laws. A clean interface accommodates linking in actual flight code in place of the prototype control laws. The same simulation may be used for rapid feasibility studies, and then used for flight software validation as the design matures. Agora is open-source and portable across computing platforms, making it customizable and extensible. It is written to support the entire GNC (guidance, navigation, and control) design cycle, from rapid prototyping and design analysis, to high-fidelity flight code verification. As a top-down design, Agora is intended to accommodate a large range of missions, anywhere in the solar system. Both two-body and three-body flight regimes are supported, as well as seamless transition between them. Multiple spacecraft may be simultaneously simulated, enabling simulation of rendezvous scenarios, as well as formation flying. Built-in reference frames and orbit perturbation dynamics provide accurate modeling of precision formation control.

Crew activity and motion effects on the space station

NASA Technical Reports Server (NTRS)

Rochon, Brian V.; Scheer, Steven A.

1987-01-01

Among the significant sources of internal disturbances that must be considered in the design of space station vibration control systems are the loads induced on the structure from various crew activities. Flight experiment T013, flown on the second manned mission of Skylab, measured force and moment time histories for a range of preplanned crew motions and activities. This experiment has proved itself invaluable as a source of on-orbit crew induced loads that has allowed a space station forcing function data base to be built. This will enable forced response such as acceleration and deflections, attributable to crew activity, to be calculated. The flight experiment, resultant database and structural model pre-processor, analysis examples and areas of combined research shall be described.

Flight-Test-Determined Aerodynamic Force and Moment Characteristics of the X-43A at Mach 7.0

NASA Technical Reports Server (NTRS)

Davis. Marl C.; White, J. Terry

2006-01-01

The second flight of the Hyper-X program afforded a unique opportunity to determine the aerodynamic force and moment characteristics of an airframe-integrated scramjet-powered aircraft in hypersonic flight. These data were gathered via a repeated series of pitch, yaw, and roll doublets; frequency sweeps; and pushover-pullup maneuvers performed throughout the X-43A cowl-closed descent. Maneuvers were conducted at Mach numbers of 6.80 to 0.95 and altitudes from 92,000 ft msl to sea level. The dynamic pressure varied from 1300 psf to 400 psf with the angle of attack ranging from 0 deg to 14 deg. The flight-extracted aerodynamics were compared with preflight predictions based on wind-tunnel-test data. The X-43A flight-derived axial force was found to be 10 percent to 15 percent higher than prediction. Under-predictions of similar magnitude were observed for the normal force. For Mach numbers above 4.0, the flight-derived stability and control characteristics resulted in larger-than-predicted static margins, with the largest discrepancy approximately 5 in. forward along the x-axis center of gravity at Mach 6.0. This condition would result in less static margin in pitch. The predicted lateral-directional stability and control characteristics matched well with flight data when allowance was made for the high uncertainty in angle of sideslip.

X-43A Flight-Test-Determined Aerodynamic Force and Moment Characteristics at Mach 7.0

NASA Technical Reports Server (NTRS)

Davis, Mark C.; White, J. Terry

2008-01-01

The second flight of the Hyper-X program afforded a unique opportunity to determine the aerodynamic force and moment characteristics of an airframe-integrated scramjet-powered aircraft in hypersonic flight. These data were gathered via a repeated series of pitch, yaw, and roll doublets, frequency sweeps, and pushover-pullup maneuvers performed throughout the X-43A cowl-closed descent. Maneuvers were conducted at Mach numbers of 6.80-0.95 and at altitudes from 92,000 ft mean sea level to sea level. The dynamic pressure varied from 1300 to 400 psf with the angle of attack ranging from 0 to 14 deg. The flight-extracted aerodynamics were compared with preflight predictions based on wind-tunnel test data. The X-43A flight-derived axial force was found to be 10-15%higher than prediction. Underpredictions of similar magnitude were observed for the normal force. For Mach numbers above 4.0, the flight-derived stability and control characteristics resulted in larger-than-predicted static margins, with the largest discrepancy approximately 5 in. forward along the x-axis center of gravity at Mach 6.0. This condition would result in less static margin in pitch. The predicted lateral-directional stability and control characteristics matched well with flight data when allowance was made for the high uncertainty in angle of sideslip.

KSC-2014-3813

NASA Image and Video Library

2014-09-11

CAPE CANAVERAL, Fla. – Preparations are underway at the Neil Armstrong Operations and Checkout Building at NASA's Kennedy Space Center in Florida for the move of the Orion spacecraft for Exploration Flight Test-1 out of the high bay doors. Inside the high bay from left, are Jules Schneider, Lockheed Martin senior manager, and Kennedy Center Director Bob Cabana. The spacecraft will be transported to the Payload Hazardous Servicing Facility where it will be fueled ahead of its December flight test. Orion is the exploration spacecraft designed to carry astronauts to destinations not yet explored by humans, including an asteroid and Mars. It will have emergency abort capability, sustain the crew during space travel and provide safe re-entry from deep space return velocities. The first unpiloted test flight of the Orion is scheduled to launch atop a United Launch Alliance Delta IV Heavy rocket from Cape Canaveral Air Force Station in Florida to an altitude of 3,600 miles above the Earth's surface. The two-orbit, four-hour flight test will help engineers evaluate the systems critical to crew safety including the heat shield, parachute system and launch abort system. For more information, visit http://www.nasa.gov/orion. Photo credit: NASA/Daniel Casper

A high performance sensor for triaxial cutting force measurement in turning.

PubMed

Zhao, You; Zhao, Yulong; Liang, Songbo; Zhou, Guanwu

2015-04-03

This paper presents a high performance triaxial cutting force sensor with excellent accuracy, favorable natural frequency and acceptable cross-interference for high speed turning process. Octagonal ring is selected as sensitive element of the designed sensor, which is drawn inspiration from ring theory. A novel structure of two mutual-perpendicular octagonal rings is proposed and three Wheatstone full bridge circuits are specially organized in order to obtain triaxial cutting force components and restrain cross-interference. Firstly, the newly developed sensor is tested in static calibration; test results indicate that the sensor possesses outstanding accuracy in the range of 0.38%-0.83%. Secondly, impacting modal tests are conducted to identify the natural frequencies of the sensor in triaxial directions (i.e., 1147 Hz, 1122 Hz and 2035 Hz), which implies that the devised sensor can be used for cutting force measurement in a high speed lathe when the spindle speed does not exceed 17,205 rev/min in continuous cutting condition. Finally, an application of the sensor in turning process is operated to show its performance for real-time cutting force measurement; the measured cutting forces demonstrate a good accordance with the variation of cutting parameters. Thus, the developed sensor possesses perfect properties and it gains great potential for real-time cutting force measurement in turning.

A High Performance Sensor for Triaxial Cutting Force Measurement in Turning

PubMed Central

Zhao, You; Zhao, Yulong; Liang, Songbo; Zhou, Guanwu

2015-01-01

This paper presents a high performance triaxial cutting force sensor with excellent accuracy, favorable natural frequency and acceptable cross-interference for high speed turning process. Octagonal ring is selected as sensitive element of the designed sensor, which is drawn inspiration from ring theory. A novel structure of two mutual-perpendicular octagonal rings is proposed and three Wheatstone full bridge circuits are specially organized in order to obtain triaxial cutting force components and restrain cross-interference. Firstly, the newly developed sensor is tested in static calibration; test results indicate that the sensor possesses outstanding accuracy in the range of 0.38%–0.83%. Secondly, impacting modal tests are conducted to identify the natural frequencies of the sensor in triaxial directions (i.e., 1147 Hz, 1122 Hz and 2035 Hz), which implies that the devised sensor can be used for cutting force measurement in a high speed lathe when the spindle speed does not exceed 17,205 rev/min in continuous cutting condition. Finally, an application of the sensor in turning process is operated to show its performance for real-time cutting force measurement; the measured cutting forces demonstrate a good accordance with the variation of cutting parameters. Thus, the developed sensor possesses perfect properties and it gains great potential for real-time cutting force measurement in turning. PMID:25855035

Development of Variable Camber Continuous Trailing Edge Flap for Performance Adaptive Aeroelastic Wing

NASA Technical Reports Server (NTRS)

Nguyen, Nhan; Kaul, Upender; Lebofsky, Sonia; Ting, Eric; Chaparro, Daniel; Urnes, James

2015-01-01

This paper summarizes the recent development of an adaptive aeroelastic wing shaping control technology called variable camber continuous trailing edge flap (VCCTEF). As wing flexibility increases, aeroelastic interactions with aerodynamic forces and moments become an increasingly important consideration in aircraft design and aerodynamic performance. Furthermore, aeroelastic interactions with flight dynamics can result in issues with vehicle stability and control. The initial VCCTEF concept was developed in 2010 by NASA under a NASA Innovation Fund study entitled "Elastically Shaped Future Air Vehicle Concept," which showed that highly flexible wing aerodynamic surfaces can be elastically shaped in-flight by active control of wing twist and bending deflection in order to optimize the spanwise lift distribution for drag reduction. A collaboration between NASA and Boeing Research & Technology was subsequently funded by NASA from 2012 to 2014 to further develop the VCCTEF concept. This paper summarizes some of the key research areas conducted by NASA during the collaboration with Boeing Research and Technology. These research areas include VCCTEF design concepts, aerodynamic analysis of VCCTEF camber shapes, aerodynamic optimization of lift distribution for drag minimization, wind tunnel test results for cruise and high-lift configurations, flutter analysis and suppression control of flexible wing aircraft, and multi-objective flight control for adaptive aeroelastic wing shaping control.

Nano-mechanical properties and structural of a 3D-printed biodegradable biomimetic micro air vehicle wing

NASA Astrophysics Data System (ADS)

Salami, E.; Montazer, E.; Ward, T. A.; Ganesan, P. B.

2017-06-01

The biomimetic micro air vehicles (BMAV) are unmanned, micro-scaled aircraft that are bio-inspired from flying organisms to achieve the lift and thrust by flapping their wings. The main objectives of this study are to design a BMAV wing (inspired from the dragonfly) and analyse its nano-mechanical properties. In order to gain insights into the flight mechanics of dragonfly, reverse engineering methods were used to establish three-dimensional geometrical models of the dragonfly wings, so we can make a comparative analysis. Then mechanical test of the real dragonfly wings was performed to provide experimental parameter values for mechanical models in terms of nano-hardness and elastic modulus. The mechanical properties of wings were measured by nanoindentre. Finally, a simplified model was designed and the dragonfly-like wing frame structure was bio-mimicked and fabricated using a 3D printer. Then mechanical test of the BMAV wings was performed to analyse and compare the wings under a variety of simplified load regimes that are concentrated force, uniform line-load and a torque. This work opened up the possibility towards developing an engineering basis for the biomimetic design of BMAV wings.

Retired NASA research pilot and former astronaut Gordon Fullerton was greeted by scores of NASA Dryden staff who bid him farewell after his final NASA flight.

NASA Image and Video Library

2007-12-21

Long-time NASA Dryden research pilot and former astronaut C. Gordon Fullerton capped an almost 50-year flying career, including more than 38 years with NASA, with a final flight in a NASA F/A-18 on Dec. 21, 2007. Fullerton and Dryden research pilot Jim Smolka flew a 90-minute pilot proficiency formation aerobatics flight with another Dryden F/A-18 and a Dryden T-38 before concluding with two low-level formation flyovers of Dryden before landing. Fullerton was honored with a water-cannon spray arch provided by two fire trucks from the Edwards Air Force Base fire department as he taxied the F/A-18 up to the Dryden ramp, and was then greeted by his wife Marie and several hundred Dryden staff after his final flight. Fullerton began his flying career with the U.S. Air Force in 1958 after earning bachelor's and master's degrees in mechanical engineering from the California Institute of Technology. Initially trained as a fighter pilot, he later transitioned to multi-engine bombers and became a bomber operations test pilot after attending the Air Force Aerospace Research Pilot School at Edwards Air Force Base, Calif. He then was assigned to the flight crew for the planned Air Force Manned Orbital Laboratory in 1966. Upon cancellation of that program, the Air Force assigned Fullerton to NASA's astronaut corps in 1969. He served on the support crews for the Apollo 14, 15, 16 and 17 lunar missions, and was later assigned to one of the two flight crews that piloted the space shuttle prototype Enterprise during the Approach and Landing Test program at Dryden. He then logged some 382 hours in space when he flew on two early space shuttle missions, STS-3 on Columbia in 1982 and STS-51F on Challenger in 1985. He joined the flight crew branch at NASA Dryden after leaving the astronaut corps in 1986. During his 21 years at Dryden, Fullerton was project pilot on a number of high-profile research efforts, including the Propulsion Controlled Aircraft, the high-speed landing tests of

Two NASA Dryden F/A-18s flown by Gordon Fullerton and Nils Larson fly in tight formation Dec. 21, 2007 during Fullerton's final flight before his retirement.

NASA Image and Video Library

2007-12-21

Long-time NASA Dryden research pilot and former astronaut C. Gordon Fullerton capped an almost 50-year flying career, including more than 38 years with NASA, with a final flight in a NASA F/A-18 on Dec. 21, 2007. Fullerton and Dryden research pilot Jim Smolka flew a 90-minute pilot proficiency formation aerobatics flight with another Dryden F/A-18 and a Dryden T-38 before concluding with two low-level formation flyovers of Dryden before landing. Fullerton was honored with a water-cannon spray arch provided by two fire trucks from the Edwards Air Force Base fire department as he taxied the F/A-18 up to the Dryden ramp, and was then greeted by his wife Marie and several hundred Dryden staff after his final flight. Fullerton began his flying career with the U.S. Air Force in 1958 after earning bachelor's and master's degrees in mechanical engineering from the California Institute of Technology. Initially trained as a fighter pilot, he later transitioned to multi-engine bombers and became a bomber operations test pilot after attending the Air Force Aerospace Research Pilot School at Edwards Air Force Base, Calif. He then was assigned to the flight crew for the planned Air Force Manned Orbital Laboratory in 1966. Upon cancellation of that program, the Air Force assigned Fullerton to NASA's astronaut corps in 1969. He served on the support crews for the Apollo 14, 15, 16 and 17 lunar missions, and was later assigned to one of the two flight crews that piloted the space shuttle prototype Enterprise during the Approach and Landing Test program at Dryden. He then logged some 382 hours in space when he flew on two early space shuttle missions, STS-3 on Columbia in 1982 and STS-51F on Challenger in 1985. He joined the flight crew branch at NASA Dryden after leaving the astronaut corps in 1986. During his 21 years at Dryden, Fullerton was project pilot on a number of high-profile research efforts, including the Propulsion Controlled Aircraft, the high-speed landing tests of

X-36 Tailless Fighter Agility Research Aircraft on lakebed during high-speed taxi tests

NASA Technical Reports Server (NTRS)

1996-01-01

The NASA/McDonnell Douglas Corporation (MDC) X-36 Tailless Fighter Agility Research Aircraft undergoes high-speed taxi tests on Rogers Dry Lake at NASA Dryden Flight Research Center, Edwards, California, on October 17, 1996. The aircraft was tested at speeds up to 85 knots. Normal takeoff speed would be 110 knots. The NASA/Boeing X-36 Tailless Fighter Agility Research Aircraft program successfully demonstrated the tailless fighter design using advanced technologies to improve the maneuverability and survivability of possible future fighter aircraft. The program met or exceeded all project goals. For 31 flights during 1997 at the Dryden Flight Research Center, Edwards, California, the project team examined the aircraft's agility at low speed / high angles of attack and at high speed / low angles of attack. The aircraft's speed envelope reached up to 206 knots (234 mph). This aircraft was very stable and maneuverable. It handled very well. The X-36 vehicle was designed to fly without the traditional tail surfaces common on most aircraft. Instead, a canard forward of the wing was used as well as split ailerons and an advanced thrust-vectoring nozzle for directional control. The X-36 was unstable in both pitch and yaw axes, so an advanced, single-channel digital fly-by-wire control system (developed with some commercially available components) was put in place to stabilize the aircraft. Using a video camera mounted in the nose of the aircraft and an onboard microphone, the X-36 was remotely controlled by a pilot in a ground station virtual cockpit. A standard fighter-type head-up display (HUD) and a moving-map representation of the vehicle's position within the range in which it flew provided excellent situational awareness for the pilot. This pilot-in-the-loop approach eliminated the need for expensive and complex autonomous flight control systems and the risks associated with their inability to deal with unknown or unforeseen phenomena in flight. Fully fueled the X-36 prototype weighed approximately 1,250 pounds. It was 19 feet long and three feet high with a wingspan of just over 10 feet. A Williams International F112 turbofan engine provided close to 700 pounds of thrust. A typical research flight lasted 35 to 45 minutes from takeoff to touchdown. A total of 31 successful research flights were flown from May 17, 1997, to November 12, 1997, amassing 15 hours and 38 minutes of flight time. The aircraft reached an altitude of 20,200 feet and a maximum angle of attack of 40 degrees. In a follow-on effort, the Air Force Research Laboratory (AFRL), Wright-Patterson Air Force Base, Ohio, contracted with Boeing to fly AFRL's Reconfigurable Control for Tailless Fighter Aircraft (RESTORE) software as a demonstration of the adaptability of the neural-net algorithm to compensate for in-flight damage or malfunction of effectors, such as flaps, ailerons and rudders. Two RESTORE research flights were flown in December 1998, proving the viability of the software approach. The X-36 aircraft flown at the Dryden Flight Research Center in 1997 was a 28-percent scale representation of a theoretical advanced fighter aircraft. The Boeing Phantom Works (formerly McDonnell Douglas) in St. Louis, Missouri, built two of the vehicles in a cooperative agreement with the Ames Research Center, Moffett Field, California.

Improved Continuous-Time Higher Harmonic Control Using Hinfinity Methods

NASA Astrophysics Data System (ADS)

Fan, Frank H.

The helicopter is a versatile aircraft that can take-off and land vertically, hover efficiently, and maneuver in confined space. This versatility is enabled by the main rotor, which also causes undesired harmonic vibration during operation. This unwanted vibration has a negative impact on the practicality of the helicopter and also increases its operational cost. Passive control techniques have been applied to helicopter vibration suppression, but these methods are generally heavy and are not robust to changes in operating conditions. Feedback control offers the advantages of robustness and potentially higher performance over passive control techniques, and amongst the various feedback schemes, Shaw's higher harmonic control algorithm has been shown to be an effective method for attenuating harmonic disturbance in helicopters. In this thesis, the higher harmonic disturbance algorithm is further developed to achieve improved performance. One goal in this thesis is to determine the importance of periodicity in the helicopter rotor dynamics for control synthesis. Based on the analysis of wind tunnel data and simulation results, we conclude the helicopter rotor can be modeled reasonably well as linear and time-invariant for control design purposes. Modeling the helicopter rotor as linear time-invariant allows us to apply linear control theory concepts to the higher harmonic control problem. Another goal in this thesis is to find the limits of performance in harmonic disturbance rejection. To achieve this goal, we first define the metrics to measure the performance of the controller in terms of response speed and robustness to changes in the plant dynamics. The performance metrics are incorporated into an Hinfinity control problem. For a given plant, the resulting Hinfinity controller achieves the maximum performance, thus allowing us to identify the performance limitation in harmonic disturbance rejection. However, the Hinfinity controllers are of high order, and may have unstable poles, leading us to develop a design method to generate stable, fixed-order, and high performance controllers. Both the Hinfinity and the fixed-order controllers are designed for constant flight conditions. A gain-scheduled control law is used to reduce the vibration throughout the flight envelope. The gain-scheduling is accomplished by blending the outputs from fixed-order controllers designed for different flight conditions. The structure of the fixed-order controller allows the usage of a previously developed anti-windup scheme, and the blending function results in a bumpless full flight envelope control law. (Copies available exclusively from MIT Libraries, libraries.mit.edu/docs - docs mit.edu)

Precision star-tracking telescope

NASA Technical Reports Server (NTRS)

Fairbank, W. M.; Everitt, C. W. F.

1972-01-01

The design, construction, and preliminary testing of a new high accuracy star tracking telescope for the laboratory model of the Stanford gyro relativity experiment are described. The function of the telescope in the final flight experiment is to define (by reference to a suitable star) a direction in space for comparison with the relativistic precession of a group of gyroscopes. The design of the telescope has been strongly affected by designs for other portions of the overall experiments, for example the gyroscopes, the attitude control system of the satellite, and the instrumentation system used in processing relativity data. Main goals for the star tracker are: (1) independent readout of angular position in two planes; (2) absolute null stability over a one year period of mechanical parts; (3) readout linear to 0.001 arc-seconds over + or - 0.05 arc-second; (4) noise performance leading to a resolution of 0.05 arc-second in 0.1 second observation time of the chosen reference star; and (5) provision for automatic gain control capable of matching the gains of the gyroscopes and telescope readouts to 1% or better.

Effects of Inertial Power and Inertial Force on Bat Wings.

PubMed

Yin, Dongfu; Zhang, Zhisheng; Dai, Min

2016-06-01

The inertial power and inertial force of wings are important factors in evaluating the flight performance of native bats. Based on measurement data of wing size and motions of Eptesicus fuscus, we present a new computational bat wing model with divided fragments of skeletons and membrane. The motions of the model were verified by comparing the joint and tip trajectories with native bats. The influences of flap, sweep, elbow, wrist and digits motions, the effects of different bones and membrane of bat wing, the components on vertical, spanwise and fore-aft directions of the inertial power and force were analyzed. Our results indicate that the flap, sweep, and elbow motions contribute the main inertial power and force; the membrane occupies an important proportion of the inertial power and force; inertial power on flap direction was larger, while variations of inertial forces on different directions were not evident. These methods and results offer insights into flight dynamics in other flying animals and may contribute to the design of future robotic bats.

X-36 Tailless Fighter Agility Research Aircraft in flight

NASA Technical Reports Server (NTRS)

1997-01-01

The tailless X-36 technology demonstrator research aircraft cruises over the California desert at low altitude during a 1997 research flight. The NASA/Boeing X-36 Tailless Fighter Agility Research Aircraft program successfully demonstrated the tailless fighter design using advanced technologies to improve the maneuverability and survivability of possible future fighter aircraft. The program met or exceeded all project goals. For 31 flights during 1997 at the Dryden Flight Research Center, Edwards, California, the project team examined the aircraft's agility at low speed / high angles of attack and at high speed / low angles of attack. The aircraft's speed envelope reached up to 206 knots (234 mph). This aircraft was very stable and maneuverable. It handled very well. The X-36 vehicle was designed to fly without the traditional tail surfaces common on most aircraft. Instead, a canard forward of the wing was used as well as split ailerons and an advanced thrust-vectoring nozzle for directional control. The X-36 was unstable in both pitch and yaw axes, so an advanced, single-channel digital fly-by-wire control system (developed with some commercially available components) was put in place to stabilize the aircraft. Using a video camera mounted in the nose of the aircraft and an onboard microphone, the X-36 was remotely controlled by a pilot in a ground station virtual cockpit. A standard fighter-type head-up display (HUD) and a moving-map representation of the vehicle's position within the range in which it flew provided excellent situational awareness for the pilot. This pilot-in-the-loop approach eliminated the need for expensive and complex autonomous flight control systems and the risks associated with their inability to deal with unknown or unforeseen phenomena in flight. Fully fueled the X-36 prototype weighed approximately 1,250 pounds. It was 19 feet long and three feet high with a wingspan of just over 10 feet. A Williams International F112 turbofan engine provided close to 700 pounds of thrust. A typical research flight lasted 35 to 45 minutes from takeoff to touchdown. A total of 31 successful research flights were flown from May 17, 1997, to November 12, 1997, amassing 15 hours and 38 minutes of flight time. The aircraft reached an altitude of 20,200 feet and a maximum angle of attack of 40 degrees. In a follow-on effort, the Air Force Research Laboratory (AFRL), Wright-Patterson Air Force Base, Ohio, contracted with Boeing to fly AFRL's Reconfigurable Control for Tailless Fighter Aircraft (RESTORE) software as a demonstration of the adaptability of the neural-net algorithm to compensate for in-flight damage or malfunction of effectors, such as flaps, ailerons and rudders. Two RESTORE research flights were flown in December 1998, proving the viability of the software approach. The X-36 aircraft flown at the Dryden Flight Research Center in 1997 was a 28-percent scale representation of a theoretical advanced fighter aircraft. The Boeing Phantom Works (formerly McDonnell Douglas) in St. Louis, Missouri, built two of the vehicles in a cooperative agreement with the Ames Research Center, Moffett Field, California.

Commutated automatic gain control system

NASA Technical Reports Server (NTRS)

Yost, S. R.

1982-01-01

A commutated automatic gain control (AGC) system was designed and built for a prototype Loran C receiver. The receiver uses a microcomputer to control a memory aided phase-locked loop (MAPLL). The microcomputer also controls the input/output, latitude/longitude conversion, and the recently added AGC system. The circuit designed for the AGC is described, and bench and flight test results are presented. The AGC circuit described actually samples starting at a point 40 microseconds after a zero crossing determined by the software lock pulse ultimately generated by a 30 microsecond delay and add network in the receiver front end envelope detector.

Modular and Reusable Power System Design for the BRRISON Balloon Telescope

NASA Astrophysics Data System (ADS)

Truesdale, Nicholas A.

High altitude balloons are emerging as low-cost alternatives to orbital satellites in the field of telescopic observation. The near-space environment of balloons allows optics to perform near their diffraction limit. In practice, this implies that a telescope similar to the Hubble Space Telescope could be flown for a cost of tens of millions as opposed to billions. While highly feasible, the design of a balloon telescope to rival Hubble is limited by funding. Until a prototype is proven and more support for balloon science is gained, projects remain limited in both hardware costs and man hours. Thus, to effectively create and support balloon payloads, engineering designs must be efficient, modular, and if possible reusable. This thesis focuses specifically on a modular power system design for the BRRISON comet-observing balloon telescope. Time- and cost-saving techniques are developed that can be used for future missions. A modular design process is achieved through the development of individual circuit elements that span a wide range of capabilities. Circuits for power conversion, switching and sensing are designed to be combined in any configuration. These include DC-DC regulators, MOSFET drivers for switching, isolated switches, current sensors and voltage sensing ADCs. Emphasis is also given to commercially available hardware. Pre-fabricated DC-DC converters and an Arduino microcontroller simplify the design process and offer proven, cost-effective performance. The design of the BRRISON power system is developed from these low-level circuits elements. A board for main power distribution supports the majority of flight electronics, and is extensible to additional hardware in future applications. An ATX computer power supply is developed, allowing the use of a commercial ATX motherboard as the flight computer. The addition of new capabilities is explored in the form of a heater control board. Finally, the power system as a whole is described, and its overall performance analyzed. The success of the BRRISON power system during testing and flight proves its utility, both for BRRISON and for future balloon telescopes.

Development of helicopter attitude axes controlled hover flight without pilot assistance and vehicle crashes

NASA Astrophysics Data System (ADS)

Simon, Miguel

In this work, we show how to computerize a helicopter to fly attitude axes controlled hover flight without the assistance of a pilot and without ever crashing. We start by developing a helicopter research test bed system including all hardware, software, and means for testing and training the helicopter to fly by computer. We select a Remote Controlled helicopter with a 5 ft. diameter rotor and 2.2 hp engine. We equip the helicopter with a payload of sensors, computers, navigation and telemetry equipment, and batteries. We develop a differential GPS system with cm accuracy and a ground computerized navigation system for six degrees of freedom (6-DoF) free flight while tracking navigation commands. We design feedback control loops with yet-to-be-determined gains for the five control "knobs" available to a flying radio-controlled (RC) miniature helicopter: engine throttle, main rotor collective pitch, longitudinal cyclic pitch, lateral cyclic pitch, and tail rotor collective pitch. We develop helicopter flight equations using fundamental dynamics, helicopter momentum theory and blade element theory. The helicopter flight equations include helicopter rotor equations of motions, helicopter rotor forces and moments, helicopter trim equations, helicopter stability derivatives, and a coupled fuselage-rotor helicopter 6-DoF model. The helicopter simulation also includes helicopter engine control equations, a helicopter aerodynamic model, and finally helicopter stability and control equations. The derivation of a set of non-linear equations of motion for the main rotor is a contribution of this thesis work. We design and build two special test stands for training and testing the helicopter to fly attitude axes controlled hover flight, starting with one axis at a time and progressing to multiple axes. The first test stand is built for teaching and testing controlled flight of elevation and yaw (i.e., directional control). The second test stand is built for teaching and testing any one or combination of the following attitude axes controlled flight: (1) pitch, (2) roll and (3) yaw. The subsequent development of a novel method to decouple, stabilize and teach the helicopter hover flight is a primary contribution of this thesis. The novel method included the development of a non-linear modeling technique for linearizing the RPM state equation dynamics so that a simple but accurate transfer function is derivable between the "available torque of the engine" and RPM. Specifically, the main rotor and tail rotor torques are modeled accurately with a bias term plus a nonlinear term involving the product of RPM squared times the main rotor blade pitch angle raised to the three-halves power. Application of this non-linear modeling technique resulted in a simple, representative and accurate transfer function model of the open-loop plant for the entire helicopter system so that all the feedback control laws for autonomous flight purposes could be derived easily using classical control theory. This is one of the contributions of this dissertation work. After discussing the integration of hardware and software elements of our helicopter research test bed system, we perform a number of experiments and tests using the two specially built test stands. Feedback gains are derived for controlling the following: (1) engine throttle to maintain prescribed main rotor angular speed, (2) main rotor collective pitch to maintain constant elevation, (3) longitudinal cyclic pitch to maintain prescribed pitch angle, (4) lateral cyclic pitch to maintain prescribed roll angle, and (5) yaw axis to maintain prescribed compass direction. (Abstract shortened by UMI.)

Effects of Gravity on Cell Movement and Development

NASA Technical Reports Server (NTRS)

Wang, Yu-Li

2002-01-01

The main purpose of this project was to understand how the migration and growth of cultured cells respond to mechanical forces. We have made significant progress on all the proposed aims. The most important discoveries are that changes in the environmental mechanical input, such as during space flight, can induce profound changes in cell migration, growth, and programmed cell death. In addition, using genetically engineered cells, we have gained important insight into the molecular mechanism underlying such mechanosensing processes. The results are summarized.

The 20 GHz circularly polarized, high temperature superconducting microstrip antenna array

NASA Technical Reports Server (NTRS)

Morrow, Jarrett D.; Williams, Jeffery T.; Long, Stuart A.; Wolfe, John C.

1994-01-01

The primary goal was to design and characterize a four-element, 20 GHz, circularly polarized microstrip patch antenna fabricated from YBa2Cu3O(x) superconductor. The purpose is to support a high temperature superconductivity flight communications experiment between the space shuttle orbiter and the ACTS satellite. This study is intended to provide information into the design, construction, and feasibility of a circularly polarized superconducting 20 GHz downlink or cross-link antenna. We have demonstrated that significant gain improvements can be realized by using superconducting materials for large corporate fed array antennas. In addition, we have shown that when constructed from superconducting materials, the efficiency, and therefore the gain, of microstrip patches increases if the substrate is not so thick that the dominant loss mechanism for the patch is radiation into the surface waves of the conductor-backed substrate. We have considered two design configurations for a superconducting 20 GHz four-element circularly polarized microstrip antenna array. The first is the Huang array that uses properly oriented and phased linearly polarized microstrip patch elements to realize a circularly polarized pattern. The second is a gap-coupled array of circularly polarized elements. In this study we determined that although the Huang array operates well on low dielectric constant substrates, its performance becomes extremely sensitive to mismatches, interelement coupling, and design imperfections for substrates with high dielectric constants. For the gap-coupled microstrip array, we were able to fabricate and test circularly polarized elements and four-element arrays on LaAlO3 using sputtered copper films. These antennas were found to perform well, with relatively good circular polarization. In addition, we realized a four-element YBa2Cu3O(x) array of the same design and measured its pattern and gain relative to a room temperature copper array. The patterns were essentially the same as that for the copper array. The measured gain of the YBCO antenna was greater than that for the room temperature copper design at temperatures below 82K, reaching a value of 3.4 dB at the lowest temperatures.

Test pilots 1952 - Walker, Butchart, and Jones

NASA Technical Reports Server (NTRS)

1952-01-01

This photo shows test pilots, (Left-Right) Joseph A. Walker, Stanley P. Butchart and Walter P. Jones, standing in front of the Douglas D-558-II Skystreak, in 1952. These three test pilots at the National Advisory Committee for Aeronautics' High-Speed Flight Research Station probably were discussing their flights in the aircraft. Joe flew research flights on the D-558-I #3 (14 flights, first on June 29, 1951) investigating buffeting, tail loads, and longitudinal stability. He flew the D-558-II #2 (3 flights, first on April 29, 1955) and recorded data on lateral stability and control. He also made pilot check-out flights in the D-558-II #3 (2 flights, first on May 7, 1954). For fifteen years Walker served as a pilot at the Edwards flight research facility (today known as the National Aeronautics and Space Administration's Dryden Flight Research Center) on research flights as well as chase missions for other pilots on NASA and Air Force research programs. On June 8, 1966, he was flying chase in NASA's F-104N for the Air Force's experimental bomber, North American XB-70A, when he was fatally injured in a mid-air collision between the planes. Stan flew the D-558-I #3 (12 flights, first on October 19, 1951) to determine the dynamic longitudinal stability characteristics and investigations of the lateral stability and control. He made one flight in the D-558-II #3 on June 26, 1953, as a pilot check-out flight. Butchart retired from the NASA Dryden Flight Research Center at Edwards, California, on February 27, 1976, after a 25-year career in research aviation. Stan served as a research pilot, chief pilot, and director of flight operations. Walter P. Jones was a research pilot for NACA from the fall of 1950 to July 1952. He had been in the U.S. Air Force as a pilot before joining the Station. Jones flew the D-558-I #3 (5 flights, first on February 13, 1951) to study buffeting, tail loads and longitudinal stability. Jones made research flights on the D-558-II #3 ( 7 flights, first on July 20, 1951). These flights investigated pitch-up and evaluated outboard wing fences. Walt also made research flights in the Northrop X-4 (14 flights, first on March 26, 1952) and the Bell X-5 (8 flights, first on June 20, 1952). In July 1952, Walt left NACA's High-Speed Flight Research Station to join Northrop Corporation as a pilot. Returning from a test mission in a Northrop YF-89D Scorpion he was fatally injured on October 20, 1953, near Edwards Air Force Base.

A water-cannon salute from two Air Force fire trucks heralds NASA research pilot Gordon Fullerton's final mission as his NASA F/A-18 taxis beneath the spray.

NASA Image and Video Library

2007-12-21

Long-time NASA Dryden research pilot and former astronaut C. Gordon Fullerton capped an almost 50-year flying career, including more than 38 years with NASA, with a final flight in a NASA F/A-18 on Dec. 21, 2007. Fullerton and Dryden research pilot Jim Smolka flew a 90-minute pilot proficiency formation aerobatics flight with another Dryden F/A-18 and a Dryden T-38 before concluding with two low-level formation flyovers of Dryden before landing. Fullerton was honored with a water-cannon spray arch provided by two fire trucks from the Edwards Air Force Base fire department as he taxied the F/A-18 up to the Dryden ramp, and was then greeted by his wife Marie and several hundred Dryden staff after his final flight. Fullerton began his flying career with the U.S. Air Force in 1958 after earning bachelor's and master's degrees in mechanical engineering from the California Institute of Technology. Initially trained as a fighter pilot, he later transitioned to multi-engine bombers and became a bomber operations test pilot after attending the Air Force Aerospace Research Pilot School at Edwards Air Force Base, Calif. He then was assigned to the flight crew for the planned Air Force Manned Orbital Laboratory in 1966. Upon cancellation of that program, the Air Force assigned Fullerton to NASA's astronaut corps in 1969. He served on the support crews for the Apollo 14, 15, 16 and 17 lunar missions, and was later assigned to one of the two flight crews that piloted the space shuttle prototype Enterprise during the Approach and Landing Test program at Dryden. He then logged some 382 hours in space when he flew on two early space shuttle missions, STS-3 on Columbia in 1982 and STS-51F on Challenger in 1985. He joined the flight crew branch at NASA Dryden after leaving the astronaut corps in 1986. During his 21 years at Dryden, Fullerton was project pilot on a number of high-profile research efforts, including the Propulsion Controlled Aircraft, the high-speed landing tests of

Analysis of an Artificial Tailplane Icing Flight Test of a High-Wing, Twin-Engine Aircraft

NASA Astrophysics Data System (ADS)

Shaikh, Shehzad M.

The US Air Force Flight Test Center (AFFTC) conducted a civilian, Federal Aviation Administration (FAA) sponsored, evaluation of tailplane icing of a twin-turboprop business transport at Edwards Air Force Base. The flight test was conducted to evaluate ice shape growth and extent of ice on the tailplane for specific weather conditions of Liquid Water Content (LWC), droplet size, and ambient temperature. This work analyzes the flight test data comparing the drag for various tailplane icing conditions with respect to a flight test verified calibrated aircraft model. Although less than a third of the test aircraft was involved in the icing environment, the results of this analysis shows a significant increase in the aircraft drag with respect to the LWC, droplet size, and ambient temperature.

United States Air Force High School Apprenticeship Program. 1990 Program Management Report. Volume 3

DTIC Science & Technology

1991-04-18

User Guide Shelly Knupp 73 Computer-Aided Design (CAD) Area Christopher O’Dell 74 Electron Beam Lithography Suzette Yu 68 Flight Dynamics Laboratory 75...fabrication. I Mr. Ed Davis, for the background knowledge of device processes and I information on electron beam lithography . Captain Mike Cheney, for...researcher may write gates on to the wafer by a process called lithography . This is the most crucial and complex part of the process. Two types of proven

Interactions measurement payload for Shuttle

NASA Technical Reports Server (NTRS)

Guidice, D. A.; Pike, C. P.

1985-01-01

The Interactions Measurement Payload for Shuttle (IMPS) consisted of engineering experiments to determine the effects of the space environment on projected Air Force space systems. Measurements by IMPS on a polar-orbit Shuttle flight will lead to detailed knowledge of the interaction of the low-altitude polar-auroral environment on materials, equipment and technologies to be used in future large, high-power space systems. The results from the IMPS measurements will provide direct input to MIL-STD design guidelines and test standards that properly account for space-environment effects.

Geometry Modeling and Adaptive Control of Air-Breathing Hypersonic Vehicles

NASA Astrophysics Data System (ADS)

Vick, Tyler Joseph

Air-breathing hypersonic vehicles have the potential to provide global reach and affordable access to space. Recent technological advancements have made scramjet-powered flight achievable, as evidenced by the successes of the X-43A and X-51A flight test programs over the last decade. Air-breathing hypersonic vehicles present unique modeling and control challenges in large part due to the fact that scramjet propulsion systems are highly integrated into the airframe, resulting in strongly coupled and often unstable dynamics. Additionally, the extreme flight conditions and inability to test fully integrated vehicle systems larger than X-51 before flight leads to inherent uncertainty in hypersonic flight. This thesis presents a means to design vehicle geometries, simulate vehicle dynamics, and develop and analyze control systems for hypersonic vehicles. First, a software tool for generating three-dimensional watertight vehicle surface meshes from simple design parameters is developed. These surface meshes are compatible with existing vehicle analysis tools, with which databases of aerodynamic and propulsive forces and moments can be constructed. A six-degree-of-freedom nonlinear dynamics simulation model which incorporates this data is presented. Inner-loop longitudinal and lateral control systems are designed and analyzed utilizing the simulation model. The first is an output feedback proportional-integral linear controller designed using linear quadratic regulator techniques. The second is a model reference adaptive controller (MRAC) which augments this baseline linear controller with an adaptive element. The performance and robustness of each controller are analyzed through simulated time responses to angle-of-attack and bank angle commands, while various uncertainties are introduced. The MRAC architecture enables the controller to adapt in a nonlinear fashion to deviations from the desired response, allowing for improved tracking performance, stability, and robustness.

F-106 tow cable attachment and release mechanism for Eclipse program

NASA Technical Reports Server (NTRS)

1997-01-01

View of the tow cable attachment and release mechanism forward of the cockpit on the QF-106 Eclipse aircraft. This mechanism held and then released the Vectran rope used to tow the QF-106 behind an Air Force C-141A. In 1997 and 1998, the Dryden Flight Research Center at Edwards, California, supported and hosted a Kelly Space & Technology, Inc. project called Eclipse, which sought to demonstrate the feasibility of a reusable tow-launch vehicle concept. The project goal was to successfully tow, inflight, a modified QF-106 delta-wing aircraft with an Air Force C-141A transport aircraft. This would demonstrate the possibility of towing and launching an actual launch vehicle from behind a tow plane. Dryden was the responsible test organization and had flight safety responsibility for the Eclipse project. Dryden provided engineering, instrumentation, simulation, modification, maintenance, range support, and research pilots for the test program. The Air Force Flight Test Center (AFFTC), Edwards, California, supplied the C-141A transport aircraft and crew and configured the aircraft as needed for the tests. The AFFTC also provided the concept and detail design and analysis as well as hardware for the tow system and QF-106 modifications. Dryden performed the modifications to convert the QF-106 drone into the piloted EXD-01 (Eclipse eXperimental Demonstrator-01) experimental aircraft. Kelly Space & Technology hoped to use the results gleaned from the tow test in developing a series of low-cost, reusable launch vehicles. These tests demonstrated the validity of towing a delta-wing aircraft having high wing loading, validated the tow simulation model, and demonstrated various operational procedures, such as ground processing of in-flight maneuvers and emergency abort scenarios.

Wake Development behind Paired Wings with Tip and Root Trailing Vortices: Consequences for Animal Flight Force Estimates

PubMed Central

Horstmann, Jan T.; Henningsson, Per; Thomas, Adrian L. R.; Bomphrey, Richard J.

2014-01-01

Recent experiments on flapping flight in animals have shown that a variety of unrelated species shed a wake behind left and right wings consisting of both tip and root vortices. Here we present an investigation using Particle Image Velocimetry (PIV) of the behaviour and interaction of trailing vortices shed by paired, fixed wings that simplify and mimic the wake of a flying animal with a non-lifting body. We measured flow velocities at five positions downstream of two adjacent NACA 0012 aerofoils and systematically varied aspect ratio, the gap between the wings (corresponding to the width of a non-lifting body), angle of attack, and the Reynolds number. The range of aspect ratios and Reynolds number where chosen to be relevant to natural fliers and swimmers, and insect flight in particular. We show that the wake behind the paired wings deformed as a consequence of the induced flow distribution such that the wingtip vortices convected downwards while the root vortices twist around each other. Vortex interaction and wake deformation became more pronounced further downstream of the wing, so the positioning of PIV measurement planes in experiments on flying animals has an important effect on subsequent force estimates due to rotating induced flow vectors. Wake deformation was most severe behind wings with lower aspect ratios and when the distance between the wings was small, suggesting that animals that match this description constitute high-risk groups in terms of measurement error. Our results, therefore, have significant implications for experimental design where wake measurements are used to estimate forces generated in animal flight. In particular, the downstream distance of the measurement plane should be minimised, notwithstanding the animal welfare constraints when measuring the wake behind flying animals. PMID:24632825

A flight experiment to measure rarefied-flow aerodynamics

NASA Technical Reports Server (NTRS)

Blanchard, Robert C.

1990-01-01

A flight experiment to measure rarefied-flow aerodynamics of a blunt lifting body is being developed by NASA. This experiment, called the Rarefied-Flow Aerodynamic Measurement Experiment (RAME), is part of the Aeroassist Flight Experiment (AFE) mission, which is a Pathfinder design tool for aeroassisted orbital transfer vehicles. The RAME will use flight measurements from accelerometers, rate gyros, and pressure transducers, combined with knowledge of AFE in-flight mass properties and trajectory, to infer aerodynamic forces and moments in the rarefied-flow environment, including transition into the hypersonic continuum regime. Preflight estimates of the aerodynamic measurements are based upon environment models, existing computer simulations, and ground test results. Planned maneuvers at several altitudes will provide a first-time opportunity to examine gas-surface accommondation effects on aerodynamic coefficients in an environment of changing atmospheric composition. A description is given of the RAME equipment design.

Two Phase Technology Development Initiatives

NASA Technical Reports Server (NTRS)

Didion, Jeffrey R.

1999-01-01

Three promising thermal technology development initiatives, vapor compression thermal control system, electronics cooling, and electrohydrodynamics applications are outlined herein. These technologies will provide thermal engineers with additional tools to meet the thermal challenges presented by increased power densities and reduced architectural options that will be available in future spacecraft. Goddard Space Flight Center and the University of Maryland are fabricating and testing a 'proto- flight' vapor compression based thermal control system for the Ultra Long Duration Balloon (ULDB) Program. The vapor compression system will be capable of transporting approximately 400 W of heat while providing a temperature lift of 60C. The system is constructed of 'commercial off-the-shelf' hardware that is modified to meet the unique environmental requirements of the ULDB. A demonstration flight is planned for 1999 or early 2000. Goddard Space Flight Center has embarked upon a multi-discipline effort to address a number of design issues regarding spacecraft electronics. The program addressed the high priority design issues concerning the total mass of standard spacecraft electronics enclosures and the impact of design changes on thermal performance. This presentation reviews the pertinent results of the Lightweight Electronics Enclosure Program. Electronics cooling is a growing challenge to thermal engineers due to increasing power densities and spacecraft architecture. The space-flight qualification program and preliminary results of thermal performance tests of copper-water heat pipes are presented. Electrohydrodynamics (EHD) is an emerging technology that uses the secondary forces that result from the application of an electric field to a flowing fluid to enhance heat transfer and manage fluid flow. A brief review of current EHD capabilities regarding heat transfer enhancement of commercial heat exchangers and capillary pumped loops is presented. Goddard Space Flight Center research efforts applying this technique to fluid management and fluid pumping are discussed.

Control theory analysis of a three-axis VTOL flight director. M.S. Thesis - Pennsylvania State Univ.

NASA Technical Reports Server (NTRS)

Niessen, F. R.

1971-01-01

A control theory analysis of a VTOL flight director and the results of a fixed-based simulator evaluation of the flight-director commands are discussed. The VTOL configuration selected for this study is a helicopter-type VTOL which controls the direction of the thrust vector by means of vehicle-attitude changes and, furthermore, employs high-gain attitude stabilization. This configuration is the same as one which was simulated in actual instrument flight tests with a variable stability helicopter. Stability analyses are made for each of the flight-director commands, assuming a single input-output, multi-loop system model for each control axis. The analyses proceed from the inner-loops to the outer-loops, using an analytical pilot model selected on the basis of the innermost-loop dynamics. The time response of the analytical model of the system is primarily used to adjust system gains, while root locus plots are used to identify dominant modes and mode interactions.

KENNEDY SPACE CENTER, FLA. - An X-ray machine is in place to take images of four rudder speed brake actuators to be installed on the orbiter Discovery. The actuators are being X-rayed at the Cape Canaveral Air Force Station’s Radiographic High-Energy X-ray Facility to determine if the gears were installed correctly. Discovery has been assigned to the first Return to Flight mission, STS-114, a logistics flight to the International Space Station.

NASA Image and Video Library

2004-03-08

KENNEDY SPACE CENTER, FLA. - An X-ray machine is in place to take images of four rudder speed brake actuators to be installed on the orbiter Discovery. The actuators are being X-rayed at the Cape Canaveral Air Force Station’s Radiographic High-Energy X-ray Facility to determine if the gears were installed correctly. Discovery has been assigned to the first Return to Flight mission, STS-114, a logistics flight to the International Space Station.

Application of modern control design methodology to oblique wing research aircraft

NASA Technical Reports Server (NTRS)

Vincent, James H.

1991-01-01

A Linear Quadratic Regulator synthesis technique was used to design an explicit model following control system for the Oblique Wing Research Aircraft (OWRA). The forward path model (Maneuver Command Generator) was designed to incorporate the desired flying qualities and response decoupling. The LQR synthesis was based on the use of generalized controls, and it was structured to provide a proportional/integral error regulator with feedforward compensation. An unexpected consequence of this design approach was the ability to decouple the control synthesis into separate longitudinal and lateral directional designs. Longitudinal and lateral directional control laws were generated for each of the nine design flight conditions, and gain scheduling requirements were addressed. A fully coupled 6 degree of freedom open loop model of the OWRA along with the longitudinal and lateral directional control laws was used to assess the closed loop performance of the design. Evaluations were performed for each of the nine design flight conditions.

GPM High Gain Antenna System

NASA Image and Video Library

2013-11-14

The GPM High Gain Antenna System (HGAS) in integration and testing at Goddard Space Flight Center. Credit: Craig E. Huber, Chief Engineer SGT Inc, NASA Goddard Space Flight Center The Global Precipitation Measurement (GPM) mission is an international partnership co-led by NASA and the Japan Aerospace Exploration Agency (JAXA) that will provide next-generation global observations of precipitation from space. GPM will study global rain, snow and ice to better understand our climate, weather, and hydrometeorological processes. As of Novermber 2013 the GPM Core Observatory is in the final stages of testing at NASA Goddard Space Flight Center. The satellite will be flown to Japan in the fall of 2013 and launched into orbit on an HII-A rocket in early 2014. For more on the GPM mission, visit gpm.gsfc.nasa.gov/. NASA image use policy. NASA Goddard Space Flight Center enables NASA’s mission through four scientific endeavors: Earth Science, Heliophysics, Solar System Exploration, and Astrophysics. Goddard plays a leading role in NASA’s accomplishments by contributing compelling scientific knowledge to advance the Agency’s mission. Follow us on Twitter Like us on Facebook Find us on Instagram

Initial Flight Tests of the NASA F-15B Propulsion Flight Test Fixture

NASA Technical Reports Server (NTRS)

Palumbo, Nathan; Moes, Timothy R.; Vachon, M. Jake

2002-01-01

Flights of the F-15B/Propulsion Flight Test Fixture (PFTF) with a Cone Drag Experiment (CDE) attached have been accomplished at NASA Dryden Flight Research Center. Mounted underneath the fuselage of an F-15B airplane, the PFTF provides volume for experiment systems and attachment points for propulsion experiments. A unique feature of the PFTF is the incorporation of a six-degree-of-freedom force balance. The force balance mounts between the PFTF and experiment and measures three forces and moments. The CDE has been attached to the force balance for envelope expansion flights. This experiment spatially and inertially simulates a large propulsion test article. This report briefly describes the F-15B airplane, the PFTF, and the force balance. A detailed description of the CDE is provided. Force-balance ground testing and stiffness modifications are described. Flight profiles and selected flight data from the envelope expansion flights are provided and discussed, including force-balance data, the internal PFTF thermal and vibration environment, a handling qualities assessment, and performance capabilities of the F-15B airplane with the PFTF installed.

KSC-2014-2830

NASA Image and Video Library

2014-05-30

CAPE CANAVERAL, Fla. -- Lockheed Martin technicians and engineers attach the heat shield to the Orion crew module inside the Operations and Checkout Building high bay at NASA's Kennedy Space Center in Florida. Technicians have installed more than 200 instrumentation sensors on the heat shield for Exploration Flight Test-1, or EFT-1. The flight test will provide engineers with data about the heat shield's ability to protect Orion and its future crews from the 4,000-degree heat of reentry and an ocean splashdown following the spacecraft’s 20,000-mph reentry from space. Data gathered during the flight will inform decisions about design improvements on the heat shield and other Orion systems, and authenticate existing computer models and new approaches to space systems design and development. This process is critical to reducing overall risks and costs of future Orion missions. Orion is the exploration spacecraft designed to carry astronauts to destinations not yet explored by humans, including an asteroid and Mars. It will have emergency abort capability, sustain the crew during space travel and provide safe re-entry from deep space return velocities. The first unpiloted test flight of the Orion is scheduled to launch later this year atop a Delta IV rocket from Cape Canaveral Air Force Station in Florida to an altitude of 3,600 miles above the Earth's surface. The two-orbit, four-hour flight test will help engineers evaluate the systems critical to crew safety including the heat shield, parachute system and launch abort system. For more information, visit http://www.nasa.gov/orion. Photo credit: NASA/Daniel Casper

KSC-2014-2831

NASA Image and Video Library

2014-05-30

CAPE CANAVERAL, Fla. -- Lockheed Martin technicians and engineers attach the heat shield to the Orion crew module inside the Operations and Checkout Building high bay at NASA's Kennedy Space Center in Florida. Technicians have installed more than 200 instrumentation sensors on the heat shield for Exploration Flight Test-1, or EFT-1. The flight test will provide engineers with data about the heat shield's ability to protect Orion and its future crews from the 4,000-degree heat of reentry and an ocean splashdown following the spacecraft’s 20,000-mph reentry from space. Data gathered during the flight will inform decisions about design improvements on the heat shield and other Orion systems, and authenticate existing computer models and new approaches to space systems design and development. This process is critical to reducing overall risks and costs of future Orion missions. Orion is the exploration spacecraft designed to carry astronauts to destinations not yet explored by humans, including an asteroid and Mars. It will have emergency abort capability, sustain the crew during space travel and provide safe re-entry from deep space return velocities. The first unpiloted test flight of the Orion is scheduled to launch later this year atop a Delta IV rocket from Cape Canaveral Air Force Station in Florida to an altitude of 3,600 miles above the Earth's surface. The two-orbit, four-hour flight test will help engineers evaluate the systems critical to crew safety including the heat shield, parachute system and launch abort system. For more information, visit http://www.nasa.gov/orion. Photo credit: NASA/Daniel Casper

X-15 #3 in flight (USAF Photo)

NASA Technical Reports Server (NTRS)

1960-01-01

This U.S. Air Force photo shows the X-15 ship #3 (56-6672) in flight over the desert in the 1960s. Ship #3 made 65 flights during the program, attaining a top speed of Mach 5.65 and a maximum altitude of 354,200 feet. Only 10 of the 12 X-15 pilots flew Ship #3, and only eight of them earned their astronaut wings during the program. Robert White, Joseph Walker, Robert Rushworth, John 'Jack' McKay, Joseph Engle, William 'Pete' Knight, William Dana, and Michael Adams all earned their astronaut wings in Ship #3. Neil Armstrong and Milton Thompson also flew Ship #3. In fact, Armstrong piloted Ship #3 on its first flight, on 20 December 1961. On 15 November 1967, Ship #3 was launched over Delamar Lake, Nevada with Maj. Michael J. Adams at the controls. The vehicle soon reached a speed of Mach 5.2, and a peak altitude of 266,000 feet. During the climb, an electrical disturbance degraded the aircraft's controllability. Ship #3 began a slow drift in heading, which soon became a spin. Adams radioed that the X-15 'seems squirrelly' and then said 'I'm in a spin.' Through some combination of pilot technique and basic aerodynamic stability, Adams recovered from the spin and entered an inverted Mach 4.7 dive. As the X-15 plummeted into the increasingly thicker atmosphere, the Honeywell adaptive flight control system caused the vehicle to begin oscillating. As the pitching motion increased, aerodynamic forces finally broke the aircraft into several major pieces. Adams was killed when the forward fuselage impacted the desert. This was the only fatal accident during the entire X-15 program. The X-15 was a rocket-powered aircraft 50 ft long with a wingspan of 22 ft. It was a missile-shaped vehicle with an unusual wedge-shaped vertical tail, thin stubby wings, and unique side fairings that extended along the side of the fuselage. The X-15 weighed about 14,000 lb empty and approximately 34,000 lb at launch. The XLR-99 rocket engine, manufactured by Thiokol Chemical Corp., was pilot controlled and was capable of developing 57,000 lb of thrust. North American Aviation built three X-15 aircraft for the program. The X-15 research aircraft was developed to provide in-flight information and data on aerodynamics, structures, flight controls, and the physiological aspects of high-speed, high-altitude flight. A follow-on program used the aircraft as a testbed to carry various scientific experiments beyond the Earth's atmosphere on a repeated basis. For flight in the dense air of the usable atmosphere, the X-15 used conventional aerodynamic controls such as rudder surfaces on the vertical stabilizers to control yaw and movable horizontal stabilizers to control pitch when moving in synchronization or roll when moved differentially. For flight in the thin air outside of the appreciable Earth's atmosphere, the X-15 used a reaction control system. Hydrogen peroxide thrust rockets located on the nose of the aircraft provided pitch and yaw control. Those on the wings provided roll control. Because of the large fuel consumption, the X-15 was air launched from a B-52 aircraft at 45,000 ft and a speed of about 500 mph. Depending on the mission, the rocket engine provided thrust for the first 80 to 120 sec of flight. The remainder of the normal 10 to 11 min. flight was powerless and ended with a 200-mph glide landing. Generally, one of two types of X-15 flight profiles was used; a high-altitude flight plan that called for the pilot to maintain a steep rate of climb, or a speed profile that called for the pilot to push over and maintain a level altitude. The X-15 was flown over a period of nearly 10 years -- June 1959 to Oct. 1968 -- and set the world's unofficial speed and altitude records of 4,520 mph or Mach 6.7 (set by Ship #2) and 354,200 ft (set by Ship #3) in a program to investigate all aspects of manned hypersonic flight. Information gained from the highly successful X-15 program contributed to the development of the Mercury, Gemini, and Apollo manned spaceflight programs, and also the Space Shuttle program. The X-15s made a total of 199 flights, and were manufactured by North American Aviation. X-15-1, serial number 56-6670, is now located at the National Air and Space museum, Washington DC. North American X-15A-2, serial number 56-6671, is at the United States Air Force Museum, Wright-Patterson AFB, Ohio. Parts of the crashed X-15-3, serial number 56-6672, recovered in 1992 by Peter Merlin and Tony Moore (The X-Hunters) are on display at the Air Force Flight Test Center Museum at Edwards. The canopy from Ship #3, recovered during the original search in 1967, is displayed at the San Diego Aerospace Museum, San Diego, California.

Precise computer controlled positioning of robot end effectors using force sensors

NASA Technical Reports Server (NTRS)

Shieh, L. S.; Mcinnis, B. C.; Wang, J. C.

1988-01-01

A thorough study of combined position/force control using sensory feedback for a one-dimensional manipulator model, which may count for the spacecraft docking problem or be extended to the multi-joint robot manipulator problem, was performed. The additional degree of freedom introduced by the compliant force sensor is included in the system dynamics in the design of precise position control. State feedback based on the pole placement method and with integral control is used to design the position controller. A simple constant gain force controller is used as an example to illustrate the dependence of the stability and steady-state accuracy of the overall position/force control upon the design of the inner position controller. Supportive simulation results are also provided.

Entropy considerations applied to shock unsteadiness in hypersonic inlets

NASA Astrophysics Data System (ADS)

Bussey, Gillian Mary Harding

The stability of curved or rectangular shocks in hypersonic inlets in response to flow perturbations can be determined analytically from the principle of minimum entropy. Unsteady shock wave motion can have a significant effect on the flow in a hypersonic inlet or combustor. According to the principle of minimum entropy, a stable thermodynamic state is one with the lowest entropy gain. A model based on piston theory and its limits has been developed for applying the principle of minimum entropy to quasi-steady flow. Relations are derived for analyzing the time-averaged entropy gain flux across a shock for quasi-steady perturbations in atmospheric conditions and angle as a perturbation in entropy gain flux from the steady state. Initial results from sweeping a wedge at Mach 10 through several degrees in AEDC's Tunnel 9 indicates the bow shock becomes unsteady near the predicted normal Mach number. Several curved shocks of varying curvature are compared to a straight shock with the same mean normal Mach number, pressure ratio, or temperature ratio. The present work provides analysis and guidelines for designing an inlet robust to off- design flight or perturbations in flow conditions an inlet is likely to face. It also suggests that inlets with curved shocks are less robust to off-design flight than those with straight shocks such as rectangular inlets. Relations for evaluating entropy perturbations for highly unsteady flow across a shock and limits on their use were also developed. The normal Mach number at which a shock could be stable to high frequency upstream perturbations increases as the speed of the shock motion increases and slightly decreases as the perturbation size increases. The present work advances the principle of minimum entropy theory by providing additional validity for using the theory for time-varying flows and applying it to shocks, specifically those in inlets. While this analytic tool is applied in the present work for evaluating the stability of shocks in hypersonic inlets, it can be used for an arbitrary application with a shock.

Aerodynamic performance and particle image velocimetery of piezo actuated biomimetic manduca sexta engineered wings towards the design and application of a flapping wing flight vehicle

NASA Astrophysics Data System (ADS)

DeLuca, Anthony M.

Considerable research and investigation has been conducted on the aerodynamic performance, and the predominate flow physics of the Manduca Sexta size of biomimetically designed and fabricated wings as part of the AFIT FWMAV design project. Despite a burgeoning interest and research into the diverse field of flapping wing flight and biomimicry, the aerodynamics of flapping wing flight remains a nebulous field of science with considerable variance into the theoretical abstractions surrounding aerodynamic mechanisms responsible for aerial performance. Traditional FWMAV flight models assume a form of a quasi-steady approximation of wing aerodynamics based on an infinite wing blade element model (BEM). An accurate estimation of the lift, drag, and side force coefficients is a critical component of autonomous stability and control models. This research focused on two separate experimental avenues into the aerodynamics of AFIT's engineered hawkmoth wings|forces and flow visualization. 1. Six degree of freedom force balance testing, and high speed video analysis was conducted on 30°, 45°, and 60° angle stop wings. A novel, non-intrusive optical tracking algorithm was developed utilizing a combination of a Gaussian Mixture Model (GMM) and ComputerVision (OpenCV) tools to track the wing in motion from multiple cameras. A complete mapping of the wing's kinematic angles as a function of driving amplitude was performed. The stroke angle, elevation angle, and angle of attack were tabulated for all three wings at driving amplitudes ranging from A=0.3 to A=0.6. The wing kinematics together with the force balance data was used to develop several aerodynamic force coefficient models. A combined translational and rotational aerodynamic model predicted lift forces within 10%, and vertical forces within 6%. The total power consumption was calculated for each of the three wings, and a Figure of Merit was calculated for each wing as a general expression of the overall efficiency of the wing. Th 60° angle stop wing achieved the largest total stroke angle and generated the most lift for the lowest power consumption of the wings tested. 2. Phase averaged stereo Particle Image Velocimetry (PIV) data was collected at eight phases through the flap cycle on the 30°, 45°, and 60° angle stop wings. Wings were mounted transverse and parallel to the interrogating laser sheet, and planar velocity intersections at the wing mid-span, one chord below the wing, were compared to one another to verify data fidelity. A Rankine-Froude actuator disk model was adapted to calculate the approximate vertical thrust generated from the total momentum flux through the flapping semi-disk using the velocity field measurements. Three component stereo u, v, and w-velocity contour measurements confirmed the presence of extensive vortical structures in the vicinity of the wing. The leading edge vortex was successfully tracked through the stroke cycle appearing at approximately 25% span, increasing in circulatory strength and translational velocity down the span toward the tip, and dissipating just after 75% span. Thrust calculations showed the vertically mounted wing more accurately represented the vertical forces when compared to its corresponding force balance measurement than the horizontally mounted wing. The mid-span showed the highest vertical velocity profile below the wing; and hence, was the location responsible for the majority of lift production along the span.

Achieving the Air Force’s Energy Vision

DTIC Science & Technology

2011-01-01

sor of logistics and supply chain management at AFIT. The US Air Force is the largest con -sumer of energy in the federal gov-ernment, spending $9...technology, motors , advanced bat- teries, and ultracapacitors. AFIT is playing a critical role in meeting Air Force and industry demand for more...Flight Design displayed a parallel hybrid-electric propul- sion system with an ICE and electric motor (fig. 2) for a general aviation aircraft. A

The 1985 National Aeronautics and Space Administration's Summer High School Apprenticeship Research Program (SHARP)

NASA Technical Reports Server (NTRS)

1985-01-01

In 1985, a total of 126 talented high school students gained first hand knowledge about science and engineering careers by working directly with a NASA scientist or engineer during the summer. This marked the sixth year of operation for NASA's Summer High School Apprenticeship Research Program (SHARP). The major priority of maintaining the high standards and success of prior years was satisfied. The following eight sites participated in the Program: Ames Research Center, Ames' Dryden Flight Research Facility, Goddard Space Flight Center, Goddard's Wallop Flight Facility, Kennedy Space Center, Langley Research Center, Lewis Research Center, and Marshall Space Flight Center. Tresp Associates served as the SHARP contractor and worked closely with NASA staff at headquarters and the sites just mentioned to plan, implement, and evaluate the program.

The cart before the horse: Mariner spacecraft and launch vehicles

NASA Technical Reports Server (NTRS)

1984-01-01

Evolution of unmanned space exploration (Pioneer, Ranger, Surveyor, and Prospector) up to 1960, and the problems in the design and use of the Atlas Centaur launch vehicle were discussed. The Mariner Program was developed from the experience gained from the previous unmanned flights.

HL-10 in flight over lakebed

NASA Technical Reports Server (NTRS)

1969-01-01

The HL-10 lifting body is seen here in flight over Rogers Dry Lake at Edwards AFB. After the vehicle's fins were modified following its first flight, the HL-10 proved to be the best handling of the heavy-weight lifting bodies flown at Edwards Air Force Base. The HL-10 flew much better than the M2-F2, and pilots were eager to fly it. The HL-10 was one of five heavyweight lifting-body designs flown at NASA's Flight Research Center (FRC--later Dryden Flight Research Center), Edwards, California, from July 1966 to November 1975 to study and validate the concept of safely maneuvering and landing a low lift-over-drag vehicle designed for reentry from space. Northrop Corporation built the HL-10 and M2-F2, the first two of the fleet of 'heavy' lifting bodies flown by the NASA Flight Research Center. The contract for construction of the HL-10 and the M2-F2 was $1.8 million. 'HL' stands for horizontal landing, and '10' refers to the tenth design studied by engineers at NASA's Langley Research Center, Hampton, Va. After delivery to NASA in January 1966, the HL-10 made its first flight on Dec. 22, 1966, with research pilot Bruce Peterson in the cockpit. Although an XLR-11 rocket engine was installed in the vehicle, the first 11 drop flights from the B-52 launch aircraft were powerless glide flights to assess handling qualities, stability, and control. In the end, the HL-10 was judged to be the best handling of the three original heavy-weight lifting bodies (M2-F2/F3, HL-10, X-24A). The HL-10 was flown 37 times during the lifting body research program and logged the highest altitude and fastest speed in the Lifting Body program. On Feb. 18, 1970, Air Force test pilot Peter Hoag piloted the HL-10 to Mach 1.86 (1,228 mph). Nine days later, NASA pilot Bill Dana flew the vehicle to 90,030 feet, which became the highest altitude reached in the program. Some new and different lessons were learned through the successful flight testing of the HL-10. These lessons, when combined with information from it's sister ship, the M2-F2/F3, provided an excellent starting point for designers of future entry vehicles, including the Space Shuttle.

Pilot Peter Hoag and HL-10

NASA Technical Reports Server (NTRS)

1969-01-01

Air Force Major Peter Hoag stands in front of the HL-10 Lifting Body. Maj. Hoag joined the HL-10 program in 1969 and made his first glide flight on June 6, 1969. He made a total of 8 flights in the HL-10. They included the fastest lifting-body flight, which reached Mach 1.861 on Feb. 18, 1970. The HL-10 was one of five heavyweight lifting-body designs flown at NASA's Flight Research Center (FRC--later Dryden Flight Research Center), Edwards, California, from July 1966 to November 1975 to study and validate the concept of safely maneuvering and landing a low lift-over-drag vehicle designed for reentry from space. Northrop Corporation built the HL-10 and M2-F2, the first two of the fleet of 'heavy' lifting bodies flown by the NASA Flight Research Center. The contract for construction of the HL-10 and the M2-F2 was $1.8 million. 'HL' stands for horizontal landing, and '10' refers to the tenth design studied by engineers at NASA's Langley Research Center, Hampton, Va. After delivery to NASA in January 1966, the HL-10 made its first flight on Dec. 22, 1966, with research pilot Bruce Peterson in the cockpit. Although an XLR-11 rocket engine was installed in the vehicle, the first 11 drop flights from the B-52 launch aircraft were powerless glide flights to assess handling qualities, stability, and control. In the end, the HL-10 was judged to be the best handling of the three original heavy-weight lifting bodies (M2-F2/F3, HL-10, X-24A). The HL-10 was flown 37 times during the lifting body research program and logged the highest altitude and fastest speed in the Lifting Body program. On Feb. 18, 1970, Air Force test pilot Peter Hoag piloted the HL-10 to Mach 1.86 (1,228 mph). Nine days later, NASA pilot Bill Dana flew the vehicle to 90,030 feet, which became the highest altitude reached in the program. Some new and different lessons were learned through the successful flight testing of the HL-10. These lessons, when combined with information from it's sister ship, the M2-F2/F3, provided an excellent starting point for designers of future entry vehicles, including the Space Shuttle.

Digital Fly-By-Wire Flight Control Validation Experience

NASA Technical Reports Server (NTRS)

Szalai, K. J.; Jarvis, C. R.; Krier, G. E.; Megna, V. A.; Brock, L. D.; Odonnell, R. N.

1978-01-01

The experience gained in digital fly-by-wire technology through a flight test program being conducted by the NASA Dryden Flight Research Center in an F-8C aircraft is described. The system requirements are outlined, along with the requirements for flight qualification. The system is described, including the hardware components, the aircraft installation, and the system operation. The flight qualification experience is emphasized. The qualification process included the theoretical validation of the basic design, laboratory testing of the hardware and software elements, systems level testing, and flight testing. The most productive testing was performed on an iron bird aircraft, which used the actual electronic and hydraulic hardware and a simulation of the F-8 characteristics to provide the flight environment. The iron bird was used for sensor and system redundancy management testing, failure modes and effects testing, and stress testing in many cases with the pilot in the loop. The flight test program confirmed the quality of the validation process by achieving 50 flights without a known undetected failure and with no false alarms.

Development and Validation of High Performance Unshrouded Centrifugal Impeller

NASA Technical Reports Server (NTRS)

Chen, Wei-Chung; Williams, M.; Paris, John K.; Prueger, G. H.; Williams, Robert; Turner, James E. (Technical Monitor)

2001-01-01

The feasibility of using a two-stage unshrouded impeller turbopump to replace the current three-stage reusable launch vehicle engine shrouded impeller hydrogen pump has been evaluated from the standpoint of turbopump weight reduction and overall payload improvement. These advantages are a by-product of the higher tip speeds that an unshrouded impeller can sustain. The issues associated with the effect of unshrouded impeller tip clearance on pump efficiency and head have been evaluated with one-dimensional tools and full three-dimensional rotordynamic fluid reaction forces and coefficients have been established through time dependent computational fluid dynamics (CFD) simulation of the whole 360 degree impeller with different rotor eccentricities and whirling ratios. Unlike the shrouded impeller, the unshrouded impeller forces are evaluated as the sum of the pressure forces on the blade and the pressure forces on the hub using the CFD results. The turbopump axial thrust control has been optimized by adjusting the first stage impeller backend wear ring seal diameter and diverting the second stage backend balance piston flow to the proper location. The structural integrity associated with the high tip speed has been checked by analyzing a 3D-Finite Element Model at maximum design conditions (6% higher than the design speed). This impeller was fabricated and tested in the NASA Marshall Space Flight Center water-test rig. The experimental data will be compared with the analytical predictions and presented in another paper. The experimental data provides validation data for the numerical design and analysis methodology. The validated numerical methodology can be used to help design different unshrouded impeller configurations.

X-15 with test pilot Major Robert M. White

NASA Technical Reports Server (NTRS)

1961-01-01

Major Robert M. White is seen here next to the X-15 aircraft after a research flight. White was one of the initial pilots selected for the X-15 program, representing the Air Force in the joint program with NASA, the Navy, and North American Aviation. Between 13 April 1960 and 14 December 1962, he made 16 flights in the rocket-powered aircraft. He was the first pilot to fly to Mach 4, 5, and 6 (respectively 4, 5, and 6 times the speed of sound). He also flew to the altitude of 314,750 feet on 17 July 1962, setting a world altitude record. This was 59.6 miles, significantly higher than the 50 miles the Air Force accepted as the beginning of space, qualifying White for astronaut wings. The X-15 was a rocket-powered aircraft. The original three aircraft were about 50 ft long with a wingspan of 22 ft. The modified #2 aircraft (X-15A-2 was longer.) They were a missile-shaped vehicles with unusual wedge-shaped vertical tails, thin stubby wings, and unique side fairings that extended along the side of the fuselage. The X-15 weighed about 14,000 lb empty and approximately 34,000 lb at launch. The XLR-99 rocket engine, manufactured by Thiokol Chemical Corp., was pilot controlled and was rated at 57,000 lb of thrust, although there are indications that it actually achieved up to 60,000 lb. North American Aviation built three X-15 aircraft for the program. The X-15 research aircraft was developed to provide in-flight information and data on aerodynamics, structures, flight controls, and the physiological aspects of high-speed, high-altitude flight. A follow-on program used the aircraft as testbeds to carry various scientific experiments beyond the Earth's atmosphere on a repeated basis. For flight in the dense air of the usable atmosphere, the X-15 used conventional aerodynamic controls such as rudder surfaces on the vertical stabilizers to control yaw and movable horizontal stabilizers to control pitch when moving in synchronization or roll when moved differentially. For flight in the thin air outside of the appreciable Earth's atmosphere, the X-15 used a reaction control system. Hydrogen peroxide thrust rockets located on the nose of the aircraft provided pitch and yaw control. Those on the wings provided roll control. Because of the large fuel consumption, the X-15 was air launched from a B-52 aircraft at approximately 45,000 ft and a speed of about 500 mph. Depending on the mission, the rocket engine provided thrust for the first 80 to 120 sec of flight. The remainder of the normal 10 to 11 min. flight was powerless and ended with a 200-mph glide landing. Generally, one of two types of X-15 flight profiles was used; a high-altitude flight plan that called for the pilot to maintain a steep rate of climb, or a speed profile that called for the pilot to push over and maintain a level altitude. The X-15 was flown over a period of nearly 10 years -- June 1959 to Oct. 1968 -- and set the world's unofficial speed and altitude records of 4,520 mph (Mach 6.7) and 354,200 ft in a program to investigate all aspects of manned hypersonic flight. Information gained from the highly successful X-15 program contributed to the development of the Mercury, Gemini, and Apollo manned spaceflight programs, and also the Space Shuttle program. The X-15s made a total of 199 flights, and were manufactured by North American Aviation. X-15-1, serial number 56-6670, is now located at the National Air and Space Museum, Washington DC. North American X-15A-2, serial number 56-6671, is at the United States Air Force Museum, Wright-Patterson AFB, Ohio. X-15-3, serial number 56-6672, crashed on 15 November 1967, resulting in the death of Maj. Michael J. Adams.

X-36 Tailless Fighter Agility Research Aircraft in flight

NASA Technical Reports Server (NTRS)

1997-01-01

The X-36 technology demonstrator shows off its distinctive shape as the remotely piloted aircraft flies a research mission over the Southern California desert on October 30, 1997. The NASA/Boeing X-36 Tailless Fighter Agility Research Aircraft program successfully demonstrated the tailless fighter design using advanced technologies to improve the maneuverability and survivability of possible future fighter aircraft. The program met or exceeded all project goals. For 31 flights during 1997 at the Dryden Flight Research Center, Edwards, California, the project team examined the aircraft's agility at low speed / high angles of attack and at high speed / low angles of attack. The aircraft's speed envelope reached up to 206 knots (234 mph). This aircraft was very stable and maneuverable. It handled very well. The X-36 vehicle was designed to fly without the traditional tail surfaces common on most aircraft. Instead, a canard forward of the wing was used as well as split ailerons and an advanced thrust-vectoring nozzle for directional control. The X-36 was unstable in both pitch and yaw axes, so an advanced, single-channel digital fly-by-wire control system (developed with some commercially available components) was put in place to stabilize the aircraft. Using a video camera mounted in the nose of the aircraft and an onboard microphone, the X-36 was remotely controlled by a pilot in a ground station virtual cockpit. A standard fighter-type head-up display (HUD) and a moving-map representation of the vehicle's position within the range in which it flew provided excellent situational awareness for the pilot. This pilot-in-the-loop approach eliminated the need for expensive and complex autonomous flight control systems and the risks associated with their inability to deal with unknown or unforeseen phenomena in flight. Fully fueled the X-36 prototype weighed approximately 1,250 pounds. It was 19 feet long and three feet high with a wingspan of just over 10 feet. A Williams International F112 turbofan engine provided close to 700 pounds of thrust. A typical research flight lasted 35 to 45 minutes from takeoff to touchdown. A total of 31 successful research flights were flown from May 17, 1997, to November 12, 1997, amassing 15 hours and 38 minutes of flight time. The aircraft reached an altitude of 20,200 feet and a maximum angle of attack of 40 degrees. In a follow-on effort, the Air Force Research Laboratory (AFRL), Wright-Patterson Air Force Base, Ohio, contracted with Boeing to fly AFRL's Reconfigurable Control for Tailless Fighter Aircraft (RESTORE) software as a demonstration of the adaptability of the neural-net algorithm to compensate for in-flight damage or malfunction of effectors, such as flaps, ailerons and rudders. Two RESTORE research flights were flown in December 1998, proving the viability of the software approach. The X-36 aircraft flown at the Dryden Flight Research Center in 1997 was a 28-percent scale representation of a theoretical advanced fighter aircraft. The Boeing Phantom Works (formerly McDonnell Douglas) in St. Louis, Missouri, built two of the vehicles in a cooperative agreement with the Ames Research Center, Moffett Field, California.

X-36 Tailless Fighter Agility Research Aircraft arrival at Dryden

NASA Technical Reports Server (NTRS)

1996-01-01

The NASA/McDonnell Douglas Corporation (MDC) X-36 Tailless Fighter Agility Research Aircraft in it's hangar at NASA Dryden Flight Research Center, Edwards, California, following its arrival on July 2, 1996. The NASA/Boeing X-36 Tailless Fighter Agility Research Aircraft program successfully demonstrated the tailless fighter design using advanced technologies to improve the maneuverability and survivability of possible future fighter aircraft. The program met or exceeded all project goals. For 31 flights during 1997 at the Dryden Flight Research Center, Edwards, California, the project team examined the aircraft's agility at low speed / high angles of attack and at high speed / low angles of attack. The aircraft's speed envelope reached up to 206 knots (234 mph). This aircraft was very stable and maneuverable. It handled very well. The X-36 vehicle was designed to fly without the traditional tail surfaces common on most aircraft. Instead, a canard forward of the wing was used as well as split ailerons and an advanced thrust-vectoring nozzle for directional control. The X-36 was unstable in both pitch and yaw axes, so an advanced, single-channel digital fly-by-wire control system (developed with some commercially available components) was put in place to stabilize the aircraft. Using a video camera mounted in the nose of the aircraft and an onboard microphone, the X-36 was remotely controlled by a pilot in a ground station virtual cockpit. A standard fighter-type head-up display (HUD) and a moving-map representation of the vehicle's position within the range in which it flew provided excellent situational awareness for the pilot. This pilot-in-the-loop approach eliminated the need for expensive and complex autonomous flight control systems and the risks associated with their inability to deal with unknown or unforeseen phenomena in flight. Fully fueled the X-36 prototype weighed approximately 1,250 pounds. It was 19 feet long and three feet high with a wingspan of just over 10 feet. A Williams International F112 turbofan engine provided close to 700 pounds of thrust. A typical research flight lasted 35 to 45 minutes from takeoff to touchdown. A total of 31 successful research flights were flown from May 17, 1997, to November 12, 1997, amassing 15 hours and 38 minutes of flight time. The aircraft reached an altitude of 20,200 feet and a maximum angle of attack of 40 degrees. In a follow-on effort, the Air Force Research Laboratory (AFRL), Wright-Patterson Air Force Base, Ohio, contracted with Boeing to fly AFRL's Reconfigurable Control for Tailless Fighter Aircraft (RESTORE) software as a demonstration of the adaptability of the neural-net algorithm to compensate for in-flight damage or malfunction of effectors, such as flaps, ailerons and rudders. Two RESTORE research flights were flown in December 1998, proving the viability of the software approach. The X-36 aircraft flown at the Dryden Flight Research Center in 1997 was a 28-percent scale representation of a theoretical advanced fighter aircraft. The Boeing Phantom Works (formerly McDonnell Douglas) in St. Louis, Missouri, built two of the vehicles in a cooperative agreement with the Ames Research Center, Moffett Field, California.

Eclipse project QF-106 and C-141A climbs out under tow on first tethered flight December 20, 1997

NASA Technical Reports Server (NTRS)

1997-01-01

TOW LAUNCH DEMONSTRATION - The Kelly Space & Technology (KST)/USAF/NASA Eclipse project's modified QF-106 climbs out under tow by a USAF C-141A on the project's first tethered flight on December 20, 1997. The successful 18-minute-long flight reached an altitude of 10,000 feet. NASA's Dryden Flight Research Center, Edwards, California, hosted the project, providing engineering and facility support as well as the project pilot. In 1997 and 1998, the Dryden Flight Research Center at Edwards, California, supported and hosted a Kelly Space & Technology, Inc. project called Eclipse, which sought to demonstrate the feasibility of a reusable tow-launch vehicle concept. The project goal was to successfully tow, inflight, a modified QF-106 delta-wing aircraft with an Air Force C-141A transport aircraft. This would demonstrate the possibility of towing and launching an actual launch vehicle from behind a tow plane. Dryden was the responsible test organization and had flight safety responsibility for the Eclipse project. Dryden provided engineering, instrumentation, simulation, modification, maintenance, range support, and research pilots for the test program. The Air Force Flight Test Center (AFFTC), Edwards, California, supplied the C-141A transport aircraft and crew and configured the aircraft as needed for the tests. The AFFTC also provided the concept and detail design and analysis as well as hardware for the tow system and QF-106 modifications. Dryden performed the modifications to convert the QF-106 drone into the piloted EXD-01 (Eclipse eXperimental Demonstrator-01) experimental aircraft. Kelly Space & Technology hoped to use the results gleaned from the tow test in developing a series of low-cost, reusable launch vehicles. These tests demonstrated the validity of towing a delta-wing aircraft having high wing loading, validated the tow simulation model, and demonstrated various operational procedures, such as ground processing of in-flight maneuvers and emergency abort scenarios.

Eclipse project QF-106 and C-141A takeoff on first tethered flight December 20, 1997

NASA Technical Reports Server (NTRS)

1997-01-01

TOW ROPE TAKEOFF - The Kelly Space & Technology (KST)/USAF Eclipse project's modified QF-106 and a USAF C-141A takeoff for the project's first tethered flight on December 20, 1997. The successful 18-minute-long flight reached an altitude of 10,000 feet. NASA's Dryden Flight Research Center, Edwards, California, hosted the project, providing engineering and facility support as well as the project pilot. In 1997 and 1998, the Dryden Flight Research Center at Edwards, California, supported and hosted a Kelly Space & Technology, Inc. project called Eclipse, which sought to demonstrate the feasibility of a reusable tow-launch vehicle concept. The project goal was to successfully tow, inflight, a modified QF-106 delta-wing aircraft with an Air Force C-141A transport aircraft. This would demonstrate the possibility of towing and launching an actual launch vehicle from behind a tow plane. Dryden was the responsible test organization and had flight safety responsibility for the Eclipse project. Dryden provided engineering, instrumentation, simulation, modification, maintenance, range support, and research pilots for the test program. The Air Force Flight Test Center (AFFTC), Edwards, California, supplied the C-141A transport aircraft and crew and configured the aircraft as needed for the tests. The AFFTC also provided the concept and detail design and analysis as well as hardware for the tow system and QF-106 modifications. Dryden performed the modifications to convert the QF-106 drone into the piloted EXD-01 (Eclipse eXperimental Demonstrator-01) experimental aircraft. Kelly Space & Technology hoped to use the results gleaned from the tow test in developing a series of low-cost, reusable launch vehicles. These tests demonstrated the validity of towing a delta-wing aircraft having high wing loading, validated the tow simulation model, and demonstrated various operational procedures, such as ground processing of in-flight maneuvers and emergency abort scenarios.

Eclipse project closeup of QF-106 under tow on takeoff on first flight December 20, 1997

NASA Technical Reports Server (NTRS)

1997-01-01

OFF THE GROUND - The Kelly Space & Technology (KST)/USAF/NASA Eclipse project's modified QF-106 lifts off under tow on the project's first tethered flight on December 20, 1997. The successful 18-minute-long flight reached an altitude of 10,000 feet. NASA's Dryden Flight Research Center, Edwards, California, hosted the project, providing engineering and facility support as well as the project pilot. In 1997 and 1998, the Dryden Flight Research Center at Edwards, California, supported and hosted a Kelly Space & Technology, Inc. project called Eclipse, which sought to demonstrate the feasibility of a reusable tow-launch vehicle concept. The project goal was to successfully tow, inflight, a modified QF-106 delta-wing aircraft with an Air Force C-141A transport aircraft. This would demonstrate the possibility of towing and launching an actual launch vehicle from behind a tow plane. Dryden was the responsible test organization and had flight safety responsibility for the Eclipse project. Dryden provided engineering, instrumentation, simulation, modification, maintenance, range support, and research pilots for the test program. The Air Force Flight Test Center (AFFTC), Edwards, California, supplied the C-141A transport aircraft and crew and configured the aircraft as needed for the tests. The AFFTC also provided the concept and detail design and analysis as well as hardware for the tow system and QF-106 modifications. Dryden performed the modifications to convert the QF-106 drone into the piloted EXD-01 (Eclipse eXperimental Demonstrator-01) experimental aircraft. Kelly Space & Technology hoped to use the results gleaned from the tow test in developing a series of low-cost, reusable launch vehicles. These tests demonstrated the validity of towing a delta-wing aircraft having high wing loading, validated the tow simulation model, and demonstrated various operational procedures, such as ground processing of in-flight maneuvers and emergency abort scenarios.

Eclipse project closeup of QF-106 under tow on first tethered flight December 20, 1997

NASA Technical Reports Server (NTRS)

1997-01-01

The Kelly Space and Technology (KST)/USAF/NASA Eclipse project's modified QF-106 is shown under tow on the project's first tethered flight on December 20, 1997. The successful 18-minute-long flight reached an altitude of 10,000 feet. NASA's Dryden Flight Research Center, Edwards, California, is hosting the project, providing engineering and facility support as well as the project pilot, Mark Stucky. In 1997 and 1998, the Dryden Flight Research Center at Edwards, California, supported and hosted a Kelly Space & Technology, Inc. project called Eclipse, which sought to demonstrate the feasibility of a reusable tow-launch vehicle concept. The project goal was to successfully tow, inflight, a modified QF-106 delta-wing aircraft with an Air Force C-141A transport aircraft. This would demonstrate the possibility of towing and launching an actual launch vehicle from behind a tow plane. Dryden was the responsible test organization and had flight safety responsibility for the Eclipse project. Dryden provided engineering, instrumentation, simulation, modification, maintenance, range support, and research pilots for the test program. The Air Force Flight Test Center (AFFTC), Edwards, California, supplied the C-141A transport aircraft and crew and configured the aircraft as needed for the tests. The AFFTC also provided the concept and detail design and analysis as well as hardware for the tow system and QF-106 modifications. Dryden performed the modifications to convert the QF-106 drone into the piloted EXD-01 (Eclipse eXperimental Demonstrator-01) experimental aircraft. Kelly Space & Technology hoped to use the results gleaned from the tow test in developing a series of low-cost, reusable launch vehicles. These tests demonstrated the validity of towing a delta-wing aircraft having high wing loading, validated the tow simulation model, and demonstrated various operational procedures, such as ground processing of in-flight maneuvers and emergency abort scenarios.

KSC-2014-3788

NASA Image and Video Library

2014-09-10

CAPE CANAVERAL, Fla. – Inside the Neil Armstrong Operations and Checkout Building high bay at NASA's Kennedy Space Center in Florida, members of the Brevard Police and Fire Pipes and Drums lead NASA and Lockheed Martin workers out of the high bay after a ceremony to turn over the Orion spacecraft for Exploration Flight Test-1 to Lockheed Martin Ground Operations from Orion Assembly, Integration and Production. Orion is the exploration spacecraft designed to carry astronauts to destinations not yet explored by humans, including an asteroid and Mars. It will have emergency abort capability, sustain the crew during space travel and provide safe re-entry from deep space return velocities. The first unpiloted test flight of the Orion is scheduled to launch atop a United Launch Alliance Delta IV Heavy rocket from Cape Canaveral Air Force Station in Florida in December to an altitude of 3,600 miles above the Earth's surface. The two-orbit, four-hour flight test will help engineers evaluate the systems critical to crew safety including the heat shield, parachute system and launch abort system. For more information, visit http://www.nasa.gov/orion. Photo credit: NASA/Daniel Casper

Flight Loads and Environments Initiative

NASA Technical Reports Server (NTRS)

Kaufman, Daniel; Kern, Dennis

2005-01-01

A viewgraph presentation on the design of a lightweight non-intrusive force measurement device (FMD) to reduce the cost per effective payload (PL) mass into orbit (CPMO) by improving launch vehicle (LV) loads and environments.

Space Launch System Ascent Flight Control Design

NASA Technical Reports Server (NTRS)

VanZwieten, Tannen S.; Orr, Jeb S.; Wall, John H.; Hall, Charles E.

2014-01-01

A robust and flexible autopilot architecture for NASA's Space Launch System (SLS) family of launch vehicles is presented. As the SLS configurations represent a potentially significant increase in complexity and performance capability of the integrated flight vehicle, it was recognized early in the program that a new, generalized autopilot design should be formulated to fulfill the needs of this new space launch architecture. The present design concept is intended to leverage existing NASA and industry launch vehicle design experience and maintain the extensibility and modularity necessary to accommodate multiple vehicle configurations while relying on proven and flight-tested control design principles for large boost vehicles. The SLS flight control architecture combines a digital three-axis autopilot with traditional bending filters to support robust active or passive stabilization of the vehicle's bending and sloshing dynamics using optimally blended measurements from multiple rate gyros on the vehicle structure. The algorithm also relies on a pseudo-optimal control allocation scheme to maximize the performance capability of multiple vectored engines while accommodating throttling and engine failure contingencies in real time with negligible impact to stability characteristics. The architecture supports active in-flight load relief through the use of a nonlinear observer driven by acceleration measurements, and envelope expansion and robustness enhancement is obtained through the use of a multiplicative forward gain modulation law based upon a simple model reference adaptive control scheme.

The Air Force concentrating photovoltaic array program

NASA Technical Reports Server (NTRS)

Geis, Jack W.

1987-01-01

A summary is given of Air Force solar concentrator projects beginning with the Rockwell International study program in 1977. The Satellite Materials Hardening Programs (SMATH) explored and developed techniques for hardening planar solar cell array power systems to the combined nuclear and laser radiation threat environments. A portion of program dollars was devoted to developing a preliminary design for a hardened solar concentrator. The results of the Survivable Concentrating Photovoltaic Array (SCOPA) program, and the design, fabrication and flight qualification of a hardened concentrator panel are discussed.

Vacuum Plasma Spray Forming of Tungsten Lorentz Force Accelerator Components

NASA Technical Reports Server (NTRS)

Zimmerman, Frank R.

2004-01-01

The Vacuum Plasma Spray (VPS) Laboratory at NASA's Marshall Space Flight Center, working with the Jet Propulsion Laboratory, has developed and demonstrated a fabrication technique using the VPS process to form anode and cathode sections for a Lorentz force accelerator made from tungsten. Lorentz force accelerators are an attractive form of electric propulsion that provides continuous, high-efficiency propulsion at useful power levels for such applications as orbit transfers or deep space missions. The VPS process is used to deposit refractory metals such as tungsten onto a graphite mandrel of the desired shape. Because tungsten is reactive at high temperatures, it is thermally sprayed in an inert environment where the plasma gun melts and deposits the molten metal powder onto a mandrel. A three-axis robot inside the chamber controls the motion of the plasma spray torch. A graphite mandrel acts as a male mold, forming the required contour and dimensions for the inside surface of the anode or cathode of the accelerator. This paper describes the processing techniques, design considerations, and process development associated with the VPS forming of Lorentz force accelerator components.

KSC-2014-4177

NASA Image and Video Library

2014-10-01

CAPE CANAVERAL, Fla. – The United Launch Alliance Delta IV Heavy rocket for Exploration Flight Test-1 is being lifted to the vertical position at the pad at Space Launch Complex 37 at Cape Canaveral Air Force Station in Florida. The Delta IV Heavy is being readied to launch Orion on its first flight test. During its first flight test, Orion will travel farther into space than any human spacecraft has gone in more than 40 years. The data gathered during the flight will influence design decisions, validate existing computer models and innovative new approaches to space systems development, as well as reduce overall mission risks and costs for later Orion flights. Liftoff of Orion on the first flight test is planned for December 2014. Photo credit: NASA/Daniel Casper

EFT-1 Delta IV Heavy lift to vertical

NASA Image and Video Library

2014-10-01

This close-up view shows the United Launch Alliance Delta IV Heavy rocket for Exploration Flight Test-1 being raised into the vertical position at the pad at Space Launch Complex 37 at Cape Canaveral Air Force Station in Florida. The Delta IV Heavy is being readied to launch Orion on its first flight test. During its first flight test, Orion will travel farther into space than any human spacecraft has gone in more than 40 years. The data gathered during the flight will influence design decisions, validate existing computer models and innovative new approaches to space systems development, as well as reduce overall mission risks and costs for later Orion flights. Liftoff of Orion on the first flight test is planned for December 2014.

EFT-1 Delta IV Heavy lift to vertical

NASA Image and Video Library

2014-10-01

The United Launch Alliance Delta IV Heavy rocket for Exploration Flight Test-1 is lifted to the vertical position in the mobile service tower on the pad at Space Launch Complex 37 at Cape Canaveral Air Force Station in Florida. The Delta IV Heavy is being readied to launch Orion on its first flight test. During its first flight test, Orion will travel farther into space than any human spacecraft has gone in more than 40 years. The data gathered during the flight will influence design decisions, validate existing computer models and innovative new approaches to space systems development, as well as reduce overall mission risks and costs for later Orion flights. Liftoff of Orion on the first flight test is planned for December 2014.

KSC-2014-4176

NASA Image and Video Library

2014-10-01

CAPE CANAVERAL, Fla. – The United Launch Alliance Delta IV Heavy rocket for Exploration Flight Test-1 is being lifted to the vertical position at the pad at Space Launch Complex 37 at Cape Canaveral Air Force Station in Florida. The Delta IV Heavy is being readied to launch Orion on its first flight test. During its first flight test, Orion will travel farther into space than any human spacecraft has gone in more than 40 years. The data gathered during the flight will influence design decisions, validate existing computer models and innovative new approaches to space systems development, as well as reduce overall mission risks and costs for later Orion flights. Liftoff of Orion on the first flight test is planned for December 2014. Photo credit: NASA/Daniel Casper

EFT-1 Delta IV Heavy lift to vertical

NASA Image and Video Library

2014-10-01

The United Launch Alliance Delta IV Heavy rocket for Exploration Flight Test-1 is being lifted to the vertical position at the pad at Space Launch Complex 37 at Cape Canaveral Air Force Station in Florida. The Delta IV Heavy is being readied to launch Orion on its first flight test. During its first flight test, Orion will travel farther into space than any human spacecraft has gone in more than 40 years. The data gathered during the flight will influence design decisions, validate existing computer models and innovative new approaches to space systems development, as well as reduce overall mission risks and costs for later Orion flights. Liftoff of Orion on the first flight test is planned for December 2014.

KSC-2014-4170

NASA Image and Video Library

2014-09-30

CAPE CANAVERAL, Fla. – The United Launch Alliance Delta IV Heavy rocket for Exploration Flight Test-1 arrives at the pad at Space Launch Complex 37 at Cape Canaveral Air Force Station in Florida. The rocket is secured on the Elevated Platform Transporter. The Delta IV Heavy will launch Orion on its first flight test. During its first flight test, Orion will travel farther into space than any human spacecraft has gone in more than 40 years. The data gathered during the flight will influence design decisions, validate existing computer models and innovative new approaches to space systems development, as well as reduce overall mission risks and costs for later Orion flights. Liftoff of Orion on the first flight test is planned for December 2014. Photo credit: NASA/Dimitri Gerondidakis

KSC-2014-4174

NASA Image and Video Library

2014-09-30

CAPE CANAVERAL, Fla. – The United Launch Alliance Delta IV Heavy rocket for Exploration Flight Test-1 has arrived at the pad at Space Launch Complex 37 at Cape Canaveral Air Force Station in Florida. The rocket is secured on the Elevated Platform Transporter. The Delta IV Heavy will launch Orion on its first flight test. During its first flight test, Orion will travel farther into space than any human spacecraft has gone in more than 40 years. The data gathered during the flight will influence design decisions, validate existing computer models and innovative new approaches to space systems development, as well as reduce overall mission risks and costs for later Orion flights. Liftoff of Orion on the first flight test is planned for December 2014. Photo credit: NASA/Dimitri Gerondidakis

KSC-2014-4171

NASA Image and Video Library

2014-09-30

CAPE CANAVERAL, Fla. – The United Launch Alliance Delta IV Heavy rocket for Exploration Flight Test-1 arrives at the pad at Space Launch Complex 37 at Cape Canaveral Air Force Station in Florida. The rocket is secured on the Elevated Platform Transporter. The Delta IV Heavy will launch Orion on its first flight test. During its first flight test, Orion will travel farther into space than any human spacecraft has gone in more than 40 years. The data gathered during the flight will influence design decisions, validate existing computer models and innovative new approaches to space systems development, as well as reduce overall mission risks and costs for later Orion flights. Liftoff of Orion on the first flight test is planned for December 2014. Photo credit: NASA/Dimitri Gerondidakis

Orbital construction demonstration study. Volume 2: Technical

NASA Technical Reports Server (NTRS)

1977-01-01

The following items are discussed in reference to OCDA requirements; (1) flight mechanics and control, (2) effects of sun angle, (3) disturbance torques, (4) control system requirements, (5) OCDA orbit decay profile, and (6) aerodynamic drag forces. Structural design requirements are also given as well as basic design definition.

Software Partitioning Schemes for Advanced Simulation Computer Systems. Final Report.

ERIC Educational Resources Information Center

Clymer, S. J.

Conducted to design software partitioning techniques for use by the Air Force to partition a large flight simulator program for optimal execution on alternative configurations, this study resulted in a mathematical model which defines characteristics for an optimal partition, and a manually demonstrated partitioning algorithm design which…

APEX 3D Propeller Test Preliminary Design

NASA Technical Reports Server (NTRS)

Colozza, Anthony J.

2002-01-01

A low Reynolds number, high subsonic mach number flight regime is fairly uncommon in aeronautics. Most flight vehicles do not fly under these aerodynamic conditions. However, recently there have been a number of proposed aircraft applications (such as high altitude observation platforms and Mars aircraft) that require flight within this regime. One of the main obstacles to flight under these conditions is the ability to reliably generate sufficient thrust for the aircraft. For a conventional propulsion system, the operation and design of the propeller is the key aspect to its operation. Due to the difficulty in experimentally modeling the flight conditions in ground-based facilities, it has been proposed to conduct propeller experiments from a high altitude gliding platform (APEX). A preliminary design of a propeller experiment under the low Reynolds number, high mach number flight conditions has been devised. The details of the design are described as well as the potential data that will be collected.

Some historical trends in the research and development of aircraft

NASA Technical Reports Server (NTRS)

Spearman, M. L.

1983-01-01

A survey of some trends in aircraft design was made in an effort to determine the relation between research, development, test, and evaluation (RDT and E) and aircraft mission capability, requirements, and objectives. Driving forces in the history of aircraft include the quest for speed which involved design concepts incorporating jet propulsion systems and low drag features. The study of high speed design concepts promoted new experimental and analytical research techniques. These research techniques, in turn, have lead to concepts offering new performance potential. Design trends were directed toward increased speed, efficiency, productivity, and safety. Generally speaking, the research and development effort has been evolutionary in nature and, with the exception of the transition to supersonic flight, little has occurred since the origin of flight that has drastically changed the basic design fundamentals of aircraft. However, this does not preclude the possibility of dramatic changes in the future since the products of research are frequently unpredictable. Advances should be expected and sought in improved aerodynamics (reduced drag, enhanced lift, flow field exploitation); propulsion (improved engine cycles, multimode engines, alternate fuels, alternate power sources); structures (new materials, manufacturing techniques); all with a view toward increased efficiency and utility.

Diving-flight aerodynamics of a peregrine falcon (Falco peregrinus).

PubMed

Ponitz, Benjamin; Schmitz, Anke; Fischer, Dominik; Bleckmann, Horst; Brücker, Christoph

2014-01-01

This study investigates the aerodynamics of the falcon Falco peregrinus while diving. During a dive peregrines can reach velocities of more than 320 km h⁻¹. Unfortunately, in freely roaming falcons, these high velocities prohibit a precise determination of flight parameters such as velocity and acceleration as well as body shape and wing contour. Therefore, individual F. peregrinus were trained to dive in front of a vertical dam with a height of 60 m. The presence of a well-defined background allowed us to reconstruct the flight path and the body shape of the falcon during certain flight phases. Flight trajectories were obtained with a stereo high-speed camera system. In addition, body images of the falcon were taken from two perspectives with a high-resolution digital camera. The dam allowed us to match the high-resolution images obtained from the digital camera with the corresponding images taken with the high-speed cameras. Using these data we built a life-size model of F. peregrinus and used it to measure the drag and lift forces in a wind-tunnel. We compared these forces acting on the model with the data obtained from the 3-D flight path trajectory of the diving F. peregrinus. Visualizations of the flow in the wind-tunnel uncovered details of the flow structure around the falcon's body, which suggests local regions with separation of flow. High-resolution pictures of the diving peregrine indicate that feathers pop-up in the equivalent regions, where flow separation in the model falcon occurred.

Structural safety of trams in case of misguidance in a switch

NASA Astrophysics Data System (ADS)

Schindler, Christian; Schwickert, Martin; Simonis, Andreas

2010-08-01

Tram vehicles mainly operate on street tracks where sometimes misguidance in switches occurs due to unfavourable conditions. Generally, in this situation, the first running gear of the vehicle follows the bend track while the next running gears continue straight ahead. This leads to a constraint that can only be solved if the vehicle's articulation is damaged or the wheel derails. The last-mentioned situation is less critical in terms of safety and costs. Five different tram types, one of them high floor, the rest low floor, were examined analytically. Numerical simulation was used to determine which wheel would be the first to derail and what level of force is needed in the articulation area between two carbodies to make a tram derail. It was shown that with pure analytical simulation, only an idea of which tram type behaves better or worse in such a situation can be gained, while a three-dimensional computational simulation gives more realistic values for the forces that arise. Three of the four low-floor tram types need much higher articulation forces to make a wheel derail in a switch misguidance situation. One particular three-car type with two single-axle running gears underneath the centre car must be designed to withstand nearly three times higher articulation forces than a conventional high-floor articulated tram. Tram designers must be aware of that and should design the carbody accordingly.

X-36 Being Prepared on Lakebed for First Flight

NASA Technical Reports Server (NTRS)

1997-01-01

Lit by the rays of the morning sunrise on Rogers Dry Lake, adjacent to NASA's Dryden Flight Research Center, Edwards, California, technicians prepare the remotely-piloted X-36 Tailless Fighter Agility Research Aircraft for its first flight in May 1997. The NASA/Boeing X-36 Tailless Fighter Agility Research Aircraft program successfully demonstrated the tailless fighter design using advanced technologies to improve the maneuverability and survivability of possible future fighter aircraft. The program met or exceeded all project goals. For 31 flights during 1997 at the Dryden Flight Research Center, Edwards, California, the project team examined the aircraft's agility at low speed / high angles of attack and at high speed / low angles of attack. The aircraft's speed envelope reached up to 206 knots (234 mph). This aircraft was very stable and maneuverable. It handled very well. The X-36 vehicle was designed to fly without the traditional tail surfaces common on most aircraft. Instead, a canard forward of the wing was used as well as split ailerons and an advanced thrust-vectoring nozzle for directional control. The X-36 was unstable in both pitch and yaw axes, so an advanced, single-channel digital fly-by-wire control system (developed with some commercially available components) was put in place to stabilize the aircraft. Using a video camera mounted in the nose of the aircraft and an onboard microphone, the X-36 was remotely controlled by a pilot in a ground station virtual cockpit. A standard fighter-type head-up display (HUD) and a moving-map representation of the vehicle's position within the range in which it flew provided excellent situational awareness for the pilot. This pilot-in-the-loop approach eliminated the need for expensive and complex autonomous flight control systems and the risks associated with their inability to deal with unknown or unforeseen phenomena in flight. Fully fueled the X-36 prototype weighed approximately 1,250 pounds. It was 19 feet long and three feet high with a wingspan of just over 10 feet. A Williams International F112 turbofan engine provided close to 700 pounds of thrust. A typical research flight lasted 35 to 45 minutes from takeoff to touchdown. A total of 31 successful research flights were flown from May 17, 1997, to November 12, 1997, amassing 15 hours and 38 minutes of flight time. The aircraft reached an altitude of 20,200 feet and a maximum angle of attack of 40 degrees. In a follow-on effort, the Air Force Research Laboratory (AFRL), Wright-Patterson Air Force Base, Ohio, contracted with Boeing to fly AFRL's Reconfigurable Control for Tailless Fighter Aircraft (RESTORE) software as a demonstration of the adaptability of the neural-net algorithm to compensate for in-flight damage or malfunction of effectors, such as flaps, ailerons and rudders. Two RESTORE research flights were flown in December 1998, proving the viability of the software approach. The X-36 aircraft flown at the Dryden Flight Research Center in 1997 was a 28-percent scale representation of a theoretical advanced fighter aircraft. The Boeing Phantom Works (formerly McDonnell Douglas) in St. Louis, Missouri, built two of the vehicles in a cooperative agreement with the Ames Research Center, Moffett Field, California.

X-36 in Flight over Mojave Desert during 5th Flight

NASA Technical Reports Server (NTRS)

1997-01-01

The unusual lines of the X-36 Tailless Fighter Agility Research Aircraft contrast sharply with the desert floor as the remotely-piloted aircraft flies over the Mojave Desert on a June 1997 research flight. The NASA/Boeing X-36 Tailless Fighter Agility Research Aircraft program successfully demonstrated the tailless fighter design using advanced technologies to improve the maneuverability and survivability of possible future fighter aircraft. The program met or exceeded all project goals. For 31 flights during 1997 at the Dryden Flight Research Center, Edwards, California, the project team examined the aircraft's agility at low speed / high angles of attack and at high speed / low angles of attack. The aircraft's speed envelope reached up to 206 knots (234 mph). This aircraft was very stable and maneuverable. It handled very well. The X-36 vehicle was designed to fly without the traditional tail surfaces common on most aircraft. Instead, a canard forward of the wing was used as well as split ailerons and an advanced thrust-vectoring nozzle for directional control. The X-36 was unstable in both pitch and yaw axes, so an advanced, single-channel digital fly-by-wire control system (developed with some commercially available components) was put in place to stabilize the aircraft. Using a video camera mounted in the nose of the aircraft and an onboard microphone, the X-36 was remotely controlled by a pilot in a ground station virtual cockpit. A standard fighter-type head-up display (HUD) and a moving-map representation of the vehicle's position within the range in which it flew provided excellent situational awareness for the pilot. This pilot-in-the-loop approach eliminated the need for expensive and complex autonomous flight control systems and the risks associated with their inability to deal with unknown or unforeseen phenomena in flight. Fully fueled the X-36 prototype weighed approximately 1,250 pounds. It was 19 feet long and three feet high with a wingspan of just over 10 feet. A Williams International F112 turbofan engine provided close to 700 pounds of thrust. A typical research flight lasted 35 to 45 minutes from takeoff to touchdown. A total of 31 successful research flights were flown from May 17, 1997, to November 12, 1997, amassing 15 hours and 38 minutes of flight time. The aircraft reached an altitude of 20,200 feet and a maximum angle of attack of 40 degrees. In a follow-on effort, the Air Force Research Laboratory (AFRL), Wright-Patterson Air Force Base, Ohio, contracted with Boeing to fly AFRL's Reconfigurable Control for Tailless Fighter Aircraft (RESTORE) software as a demonstration of the adaptability of the neural-net algorithm to compensate for in-flight damage or malfunction of effectors, such as flaps, ailerons and rudders. Two RESTORE research flights were flown in December 1998, proving the viability of the software approach. The X-36 aircraft flown at the Dryden Flight Research Center in 1997 was a 28-percent scale representation of a theoretical advanced fighter aircraft. The Boeing Phantom Works (formerly McDonnell Douglas) in St. Louis, Missouri, built two of the vehicles in a cooperative agreement with the Ames Research Center, Moffett Field, California.

X-36 during First Flight

NASA Technical Reports Server (NTRS)

1997-01-01

The remotely-piloted X-36 Tailless Fighter Agility Research Aircraft climbs out from Rogers Dry Lake at the Dryden Flight Research Center on its first flight in May 1997. The aircraft flew for five minutes and reached an altitude of approximately 4,900 feet. The NASA/Boeing X-36 Tailless Fighter Agility Research Aircraft program successfully demonstrated the tailless fighter design using advanced technologies to improve the maneuverability and survivability of possible future fighter aircraft. The program met or exceeded all project goals. For 31 flights during 1997 at the Dryden Flight Research Center, Edwards, California, the project team examined the aircraft's agility at low speed / high angles of attack and at high speed / low angles of attack. The aircraft's speed envelope reached up to 206 knots (234 mph). This aircraft was very stable and maneuverable. It handled very well. The X-36 vehicle was designed to fly without the traditional tail surfaces common on most aircraft. Instead, a canard forward of the wing was used as well as split ailerons and an advanced thrust-vectoring nozzle for directional control. The X-36 was unstable in both pitch and yaw axes, so an advanced, single-channel digital fly-by-wire control system (developed with some commercially available components) was put in place to stabilize the aircraft. Using a video camera mounted in the nose of the aircraft and an onboard microphone, the X-36 was remotely controlled by a pilot in a ground station virtual cockpit. A standard fighter-type head-up display (HUD) and a moving-map representation of the vehicle's position within the range in which it flew provided excellent situational awareness for the pilot. This pilot-in-the-loop approach eliminated the need for expensive and complex autonomous flight control systems and the risks associated with their inability to deal with unknown or unforeseen phenomena in flight. Fully fueled the X-36 prototype weighed approximately 1,250 pounds. It was 19 feet long and three feet high with a wingspan of just over 10 feet. A Williams International F112 turbofan engine provided close to 700 pounds of thrust. A typical research flight lasted 35 to 45 minutes from takeoff to touchdown. A total of 31 successful research flights were flown from May 17, 1997, to November 12, 1997, amassing 15 hours and 38 minutes of flight time. The aircraft reached an altitude of 20,200 feet and a maximum angle of attack of 40 degrees. In a follow-on effort, the Air Force Research Laboratory (AFRL), Wright-Patterson Air Force Base, Ohio, contracted with Boeing to fly AFRL's Reconfigurable Control for Tailless Fighter Aircraft (RESTORE) software as a demonstration of the adaptability of the neural-net algorithm to compensate for in-flight damage or malfunction of effectors, such as flaps, ailerons and rudders. Two RESTORE research flights were flown in December 1998, proving the viability of the software approach. The X-36 aircraft flown at the Dryden Flight Research Center in 1997 was a 28-percent scale representation of a theoretical advanced fighter aircraft. The Boeing Phantom Works (formerly McDonnell Douglas) in St. Louis, Missouri, built two of the vehicles in a cooperative agreement with the Ames Research Center, Moffett Field, California.

X-36 Being Prepared on Lakebed for First Flight

NASA Technical Reports Server (NTRS)

1997-01-01

Lit by the rays of the morning sunrise on Rogers Dry Lake, adjacent to NASA's Dryden Flight Research Center, Edwards, California, a technician prepares the remotely-piloted X-36 Tailless Fighter Agility Research Aircraft for its first flight on May 17, 1997. The NASA/Boeing X-36 Tailless Fighter Agility Research Aircraft program successfully demonstrated the tailless fighter design using advanced technologies to improve the maneuverability and survivability of possible future fighter aircraft. The program met or exceeded all project goals. For 31 flights during 1997 at the Dryden Flight Research Center, Edwards, California, the project team examined the aircraft's agility at low speed / high angles of attack and at high speed / low angles of attack. The aircraft's speed envelope reached up to 206 knots (234 mph). This aircraft was very stable and maneuverable. It handled very well. The X-36 vehicle was designed to fly without the traditional tail surfaces common on most aircraft. Instead, a canard forward of the wing was used as well as split ailerons and an advanced thrust-vectoring nozzle for directional control. The X-36 was unstable in both pitch and yaw axes, so an advanced, single-channel digital fly-by-wire control system (developed with some commercially available components) was put in place to stabilize the aircraft. Using a video camera mounted in the nose of the aircraft and an onboard microphone, the X-36 was remotely controlled by a pilot in a ground station virtual cockpit. A standard fighter-type head-up display (HUD) and a moving-map representation of the vehicle's position within the range in which it flew provided excellent situational awareness for the pilot. This pilot-in-the-loop approach eliminated the need for expensive and complex autonomous flight control systems and the risks associated with their inability to deal with unknown or unforeseen phenomena in flight. Fully fueled the X-36 prototype weighed approximately 1,250 pounds. It was 19 feet long and three feet high with a wingspan of just over 10 feet. A Williams International F112 turbofan engine provided close to 700 pounds of thrust. A typical research flight lasted 35 to 45 minutes from takeoff to touchdown. A total of 31 successful research flights were flown from May 17, 1997, to November 12, 1997, amassing 15 hours and 38 minutes of flight time. The aircraft reached an altitude of 20,200 feet and a maximum angle of attack of 40 degrees. In a follow-on effort, the Air Force Research Laboratory (AFRL), Wright-Patterson Air Force Base, Ohio, contracted with Boeing to fly AFRL's Reconfigurable Control for Tailless Fighter Aircraft (RESTORE) software as a demonstration of the adaptability of the neural-net algorithm to compensate for in-flight damage or malfunction of effectors, such as flaps, ailerons and rudders. Two RESTORE research flights were flown in December 1998, proving the viability of the software approach. The X-36 aircraft flown at the Dryden Flight Research Center in 1997 was a 28-percent scale representation of a theoretical advanced fighter aircraft. The Boeing Phantom Works (formerly McDonnell Douglas) in St. Louis, Missouri, built two of the vehicles in a cooperative agreement with the Ames Research Center, Moffett Field, California.

X-36 Being Prepared on Lakebed for First Flight

NASA Technical Reports Server (NTRS)

1997-01-01

Lit by the rays of the morning sunrise on Rogers Dry Lake, adjacent to NASA's Dryden Flight Research Center, Edwards, California, technicians prepares the remotely-piloted X-36 Tailless Fighter Agility Research Aircraft for its first flight on May 17, 1997. The NASA/Boeing X-36 Tailless Fighter Agility Research Aircraft program successfully demonstrated the tailless fighter design using advanced technologies to improve the maneuverability and survivability of possible future fighter aircraft. The program met or exceeded all project goals. For 31 flights during 1997 at the Dryden Flight Research Center, Edwards, California, the project team examined the aircraft's agility at low speed / high angles of attack and at high speed / low angles of attack. The aircraft's speed envelope reached up to 206 knots (234 mph). This aircraft was very stable and maneuverable. It handled very well. The X-36 vehicle was designed to fly without the traditional tail surfaces common on most aircraft. Instead, a canard forward of the wing was used as well as split ailerons and an advanced thrust-vectoring nozzle for directional control. The X-36 was unstable in both pitch and yaw axes, so an advanced, single-channel digital fly-by-wire control system (developed with some commercially available components) was put in place to stabilize the aircraft. Using a video camera mounted in the nose of the aircraft and an onboard microphone, the X-36 was remotely controlled by a pilot in a ground station virtual cockpit. A standard fighter-type head-up display (HUD) and a moving-map representation of the vehicle's position within the range in which it flew provided excellent situational awareness for the pilot. This pilot-in-the-loop approach eliminated the need for expensive and complex autonomous flight control systems and the risks associated with their inability to deal with unknown or unforeseen phenomena in flight. Fully fueled the X-36 prototype weighed approximately 1,250 pounds. It was 19 feet long and three feet high with a wingspan of just over 10 feet. A Williams International F112 turbofan engine provided close to 700 pounds of thrust. A typical research flight lasted 35 to 45 minutes from takeoff to touchdown. A total of 31 successful research flights were flown from May 17, 1997, to November 12, 1997, amassing 15 hours and 38 minutes of flight time. The aircraft reached an altitude of 20,200 feet and a maximum angle of attack of 40 degrees. In a follow-on effort, the Air Force Research Laboratory (AFRL), Wright-Patterson Air Force Base, Ohio, contracted with Boeing to fly AFRL's Reconfigurable Control for Tailless Fighter Aircraft (RESTORE) software as a demonstration of the adaptability of the neural-net algorithm to compensate for in-flight damage or malfunction of effectors, such as flaps, ailerons and rudders. Two RESTORE research flights were flown in December 1998, proving the viability of the software approach. The X-36 aircraft flown at the Dryden Flight Research Center in 1997 was a 28-percent scale representation of a theoretical advanced fighter aircraft. The Boeing Phantom Works (formerly McDonnell Douglas) in St. Louis, Missouri, built two of the vehicles in a cooperative agreement with the Ames Research Center, Moffett Field, California.

X-36 Taking off During First Flight

NASA Technical Reports Server (NTRS)

1997-01-01

The X-36 remotely piloted aircraft lifts off on its first flight, May 17, 1997, at NASA's Dryden Flight Research Center, Edwards, California. The aircraft flew for five minutes and reached an altitude of approximately 4,900 feet. The NASA/Boeing X-36 Tailless Fighter Agility Research Aircraft program successfully demonstrated the tailless fighter design using advanced technologies to improve the maneuverability and survivability of possible future fighter aircraft. The program met or exceeded all project goals. For 31 flights during 1997 at the Dryden Flight Research Center, Edwards, California, the project team examined the aircraft's agility at low speed / high angles of attack and at high speed / low angles of attack. The aircraft's speed envelope reached up to 206 knots (234 mph). This aircraft was very stable and maneuverable. It handled very well. The X-36 vehicle was designed to fly without the traditional tail surfaces common on most aircraft. Instead, a canard forward of the wing was used as well as split ailerons and an advanced thrust-vectoring nozzle for directional control. The X-36 was unstable in both pitch and yaw axes, so an advanced, single-channel digital fly-by-wire control system (developed with some commercially available components) was put in place to stabilize the aircraft. Using a video camera mounted in the nose of the aircraft and an onboard microphone, the X-36 was remotely controlled by a pilot in a ground station virtual cockpit. A standard fighter-type head-up display (HUD) and a moving-map representation of the vehicle's position within the range in which it flew provided excellent situational awareness for the pilot. This pilot-in-the-loop approach eliminated the need for expensive and complex autonomous flight control systems and the risks associated with their inability to deal with unknown or unforeseen phenomena in flight. Fully fueled the X-36 prototype weighed approximately 1,250 pounds. It was 19 feet long and three feet high with a wingspan of just over 10 feet. A Williams International F112 turbofan engine provided close to 700 pounds of thrust. A typical research flight lasted 35 to 45 minutes from takeoff to touchdown. A total of 31 successful research flights were flown from May 17, 1997, to November 12, 1997, amassing 15 hours and 38 minutes of flight time. The aircraft reached an altitude of 20,200 feet and a maximum angle of attack of 40 degrees. In a follow-on effort, the Air Force Research Laboratory (AFRL), Wright-Patterson Air Force Base, Ohio, contracted with Boeing to fly AFRL's Reconfigurable Control for Tailless Fighter Aircraft (RESTORE) software as a demonstration of the adaptability of the neural-net algorithm to compensate for in-flight damage or malfunction of effectors, such as flaps, ailerons and rudders. Two RESTORE research flights were flown in December 1998, proving the viability of the software approach. The X-36 aircraft flown at the Dryden Flight Research Center in 1997 was a 28-percent scale representation of a theoretical advanced fighter aircraft. The Boeing Phantom Works (formerly McDonnell Douglas) in St. Louis, Missouri, built two of the vehicles in a cooperative agreement with the Ames Research Center, Moffett Field, California.

X-36 Being Prepared on Lakebed for First Flight

NASA Technical Reports Server (NTRS)

1997-01-01

As the sun creeps above the horizon of Rogers Dry Lake at NASA's Dryden Flight Research Center, Edwards, California, technicians make final preparations for the first flight of the X-36 Tailless Fighter Agility Research Aircraft. The NASA/Boeing X-36 Tailless Fighter Agility Research Aircraft program successfully demonstrated the tailless fighter design using advanced technologies to improve the maneuverability and survivability of possible future fighter aircraft. The program met or exceeded all project goals. For 31 flights during 1997 at the Dryden Flight Research Center, Edwards, California, the project team examined the aircraft's agility at low speed / high angles of attack and at high speed / low angles of attack. The aircraft's speed envelope reached up to 206 knots (234 mph). This aircraft was very stable and maneuverable. It handled very well. The X-36 vehicle was designed to fly without the traditional tail surfaces common on most aircraft. Instead, a canard forward of the wing was used as well as split ailerons and an advanced thrust-vectoring nozzle for directional control. The X-36 was unstable in both pitch and yaw axes, so an advanced, single-channel digital fly-by-wire control system (developed with some commercially available components) was put in place to stabilize the aircraft. Using a video camera mounted in the nose of the aircraft and an onboard microphone, the X-36 was remotely controlled by a pilot in a ground station virtual cockpit. A standard fighter-type head-up display (HUD) and a moving-map representation of the vehicle's position within the range in which it flew provided excellent situational awareness for the pilot. This pilot-in-the-loop approach eliminated the need for expensive and complex autonomous flight control systems and the risks associated with their inability to deal with unknown or unforeseen phenomena in flight. Fully fueled the X-36 prototype weighed approximately 1,250 pounds. It was 19 feet long and three feet high with a wingspan of just over 10 feet. A Williams International F112 turbofan engine provided close to 700 pounds of thrust. A typical research flight lasted 35 to 45 minutes from takeoff to touchdown. A total of 31 successful research flights were flown from May 17, 1997, to November 12, 1997, amassing 15 hours and 38 minutes of flight time. The aircraft reached an altitude of 20,200 feet and a maximum angle of attack of 40 degrees. In a follow-on effort, the Air Force Research Laboratory (AFRL), Wright-Patterson Air Force Base, Ohio, contracted with Boeing to fly AFRL's Reconfigurable Control for Tailless Fighter Aircraft (RESTORE) software as a demonstration of the adaptability of the neural-net algorithm to compensate for in-flight damage or malfunction of effectors, such as flaps, ailerons and rudders. Two RESTORE research flights were flown in December 1998, proving the viability of the software approach. The X-36 aircraft flown at the Dryden Flight Research Center in 1997 was a 28-percent scale representation of a theoretical advanced fighter aircraft. The Boeing Phantom Works (formerly McDonnell Douglas) in St. Louis, Missouri, built two of the vehicles in a cooperative agreement with the Ames Research Center, Moffett Field, California.

Powersail High Power Propulsion System Design Study

NASA Astrophysics Data System (ADS)

Gulczinski, Frank S., III

2000-11-01

A desire by the United States Air Force to exploit the space environment has led to a need for increased on-orbit electrical power availability. To enable this, the Air Force Research Laboratory Space Vehicles Directorate (AFRL/ VS) is developing Powersail: a two-phased program to demonstrate high power (100 kW to 1 MW) capability in space using a deployable, flexible solar array connected to the host spacecraft using a slack umbilical. The first phase will be a proof-of-concept demonstration at 50 kW, followed by the second phase, an operational system at full power. In support of this program, the AFRL propulsion Directorate's Spacecraft Propulsion Branch (AFRL/PRS ) at Edwards AFB has commissioned a design study of the Powersail High Power Propulsion System. The purpose of this study, the results of which are summarized in this paper, is to perform mission and design trades to identify potential full-power applications (both near-Earth and interplanetary) and the corresponding propulsion system requirements and design. The design study shall farther identify a suitable low power demonstration flight that maximizes risk reduction for the fully operational system. This propulsion system is expected to be threefold: (1) primary propulsion for moving the entire vehicle, (2) a propulsion unit that maintains the solar array position relative to the host spacecraft, and (3) control propulsion for maintaining proper orientation for the flexible solar array.

Design of a candidate flutter suppression control law for DAST ARW-2. [Drones for Aerodynamic and Structural Testing Aeroelastic Research Wing

NASA Technical Reports Server (NTRS)

Adams, W. M., Jr.; Tiffany, S. H.

1983-01-01

A control law is developed to suppress symmetric flutter for a mathematical model of an aeroelastic research vehicle. An implementable control law is attained by including modified LQG (linear quadratic Gaussian) design techniques, controller order reduction, and gain scheduling. An alternate (complementary) design approach is illustrated for one flight condition wherein nongradient-based constrained optimization techniques are applied to maximize controller robustness.

Experience Gained From Launch and Early Orbit Support of the Rossi X-Ray Timing Explorer (RXTE)

NASA Technical Reports Server (NTRS)

Fink, D. R.; Chapman, K. B.; Davis, W. S.; Hashmall, J. A.; Shulman, S. E.; Underwood, S. C.; Zsoldos, J. M.; Harman, R. R.

1996-01-01

this paper reports the results to date of early mission support provided by the personnel of the Goddard Space Flight Center Flight Dynamics Division (FDD) for the Rossi X-Ray Timing Explorer (RXTE) spacecraft. For this mission, the FDD supports onboard attitude determination and ephemeris propagation by supplying ground-based orbit and attitude solutions and calibration results. The first phase of that support was to provide launch window analyses. As the launch window was determined, acquisition attitudes were calculated and calibration slews were planned. postlaunch, these slews provided the basis for ground determined calibration. Ground determined calibration results are used to improve the accuracy of onboard solutions. The FDD is applying new calibration tools designed to facilitate use of the simultaneous, high-accuracy star observations from the two RXTE star trackers for ground attitude determination and calibration. An evaluation of the performance of these tools is presented. The FDD provides updates to the onboard star catalog based on preflight analysis and analysis of flight data. The in-flight results of the mission support in each area are summarized and compared with pre-mission expectations.

Active In-Flight Load Redistribution Utilizing Fiber-Optic Shape Sensing and Multiple Control Surfaces

NASA Technical Reports Server (NTRS)

Pena, Francisco; Martins, Benjamin L.; Richards, W. Lance

2018-01-01

Morphing wing technologies have gained research interest in recent years as technological advancements pave the way for such innovations. A key benefit of such a morphing wing concept is the ability of the wing to transition into an optimal configuration at multiple flight conditions. Such a morphing wing will have applications not only in drag reduction but also in flutter suppression and gust alleviation. By manipulating the wing geometry to match a given flight profile it is likely that the wing will yield increases in not just aerodynamic efficiency but also structural efficiency. These structurally efficient designs will likely rely on some type of structural sensing system which will ensure the wing maintains positive margins throughout its flight profile.

Adaptive Flight Control Design with Optimal Control Modification on an F-18 Aircraft Model

NASA Technical Reports Server (NTRS)

Burken, John J.; Nguyen, Nhan T.; Griffin, Brian J.

2010-01-01

In the presence of large uncertainties, a control system needs to be able to adapt rapidly to regain performance. Fast adaptation is referred to as the implementation of adaptive control with a large adaptive gain to reduce the tracking error rapidly; however, a large adaptive gain can lead to high-frequency oscillations which can adversely affect the robustness of an adaptive control law. A new adaptive control modification is presented that can achieve robust adaptation with a large adaptive gain without incurring high-frequency oscillations as with the standard model-reference adaptive control. The modification is based on the minimization of the Y2 norm of the tracking error, which is formulated as an optimal control problem. The optimality condition is used to derive the modification using the gradient method. The optimal control modification results in a stable adaptation and allows a large adaptive gain to be used for better tracking while providing sufficient robustness. A damping term (v) is added in the modification to increase damping as needed. Simulations were conducted on a damaged F-18 aircraft (McDonnell Douglas, now The Boeing Company, Chicago, Illinois) with both the standard baseline dynamic inversion controller and the adaptive optimal control modification technique. The results demonstrate the effectiveness of the proposed modification in tracking a reference model.

Flight control actuation system

NASA Technical Reports Server (NTRS)

Wingett, Paul T. (Inventor); Gaines, Louie T. (Inventor); Evans, Paul S. (Inventor); Kern, James I. (Inventor)

2004-01-01

A flight control actuation system comprises a controller, electromechanical actuator and a pneumatic actuator. During normal operation, only the electromechanical actuator is needed to operate a flight control surface. When the electromechanical actuator load level exceeds 40 amps positive, the controller activates the pneumatic actuator to offset electromechanical actuator loads to assist the manipulation of flight control surfaces. The assistance from the pneumatic load assist actuator enables the use of an electromechanical actuator that is smaller in size and mass, requires less power, needs less cooling processes, achieves high output forces and adapts to electrical current variations. The flight control actuation system is adapted for aircraft, spacecraft, missiles, and other flight vehicles, especially flight vehicles that are large in size and travel at high velocities.

Flight control actuation system

NASA Technical Reports Server (NTRS)

Wingett, Paul T. (Inventor); Gaines, Louie T. (Inventor); Evans, Paul S. (Inventor); Kern, James I. (Inventor)

2006-01-01

A flight control actuation system comprises a controller, electromechanical actuator and a pneumatic actuator. During normal operation, only the electromechanical actuator is needed to operate a flight control surface. When the electromechanical actuator load level exceeds 40 amps positive, the controller activates the pneumatic actuator to offset electromechanical actuator loads to assist the manipulation of flight control surfaces. The assistance from the pneumatic load assist actuator enables the use of an electromechanical actuator that is smaller in size and mass, requires less power, needs less cooling processes, achieves high output forces and adapts to electrical current variations. The flight control actuation system is adapted for aircraft, spacecraft, missiles, and other flight vehicles, especially flight vehicles that are large in size and travel at high velocities.

Intelligent Control for the BEES Flyer

NASA Technical Reports Server (NTRS)

Krishnakumar, K.; Gundy-Burlet, Karen; Aftosmis, Mike; Nemec, Marian; Limes, Greg; Berry, Misty; Logan, Michael

2004-01-01

This paper describes the effort to provide a preliminary capability analysis and a neural network based adaptive flight control system for the JPL-led BEES aircraft project. The BEES flyer was envisioned to be a small, autonomous platform with sensing and control systems mimicking those of biological systems for the purpose of scientific exploration on the surface of Mars. The platform is physically tightly constrained by the necessity of efficient packing within rockets for the trip to Mars. Given the physical constraints, the system is not an ideal configuration for aerodynamics or stability and control. The objectives of this effort are to evaluate the aerodynamics characteristics of the existing design, to make recommendaaons as to potential improvements and to provide a control system that stabilizes the existing aircraft for nominal flight and damaged conditions. Towards this several questions are raised and analyses are presented to arrive at answers to some of the questions raised. CART3D, a high-fidelity inviscid analysis package for conceptual and preliminary aerodynamic design, was used to compute a parametric set of solutions over the expected flight domain. Stability and control derivatives were extracted from the database and integrated with the neural flight control system. The Integrated Vehicle Modeling Environment (IVME) was also used for estimating aircraft geometric, inertial, and aerodynamic characteristics. A generic neural flight control system is used to provide adaptive control without the requirement for extensive gain scheduling or explicit system identification. The neural flight control system uses reference models to specify desired handling qualities in the roll, pitch, and yaw axes, and incorporates both pre-trained and on-line learning neural networks in the inverse model portion of the controller. Results are presented for the BEES aircraft in the subsonic regime for terrestrial and Martian environments.

Multicriteria Gain Tuning for Rotorcraft Flight Controls (also entitled The Development of the Conduit Advanced Control System Design and Evaluation Interface with a Case Study Application Fly by Wire Helicopter Design)

NASA Technical Reports Server (NTRS)

Biezad, Daniel

1997-01-01

Handling qualities analysis and control law design would seem to be naturally complimenting components of aircraft flight control system design, however these two closely coupled disciplines are often not well integrated in practice. Handling qualities engineers and control system engineers may work in separate groups within an aircraft company. Flight control system engineers and handling quality specialists may come from different backgrounds and schooling and are often not aware of the other group's research. Thus while the handling qualities specifications represent desired aircraft response characteristics, these are rarely incorporated directly in the control system design process. Instead modem control system design techniques are based on servo-loop robustness specifications, and simple representations of the desired control response. Comprehensive handling qualities analysis is often left until the end of the design cycle and performed as a check of the completed design for satisfactory performance. This can lead to costly redesign or less than satisfactory aircraft handling qualities when the flight testing phase is reached. The desire to integrate the fields of handling qualities and flight,control systems led to the development of the CONDUIT system. This tool facilitates control system designs that achieve desired handling quality requirements and servo-loop specifications in a single design process. With CONDUIT, the control system engineer is now able to directly design and control systems to meet the complete handling specifications. CONDUIT allows the designer to retain a preferred control law structure, but then tunes the system parameters to meet the handling quality requirements.

Description, characteristics and testing of the NASA airborne radar

NASA Technical Reports Server (NTRS)

Jones, W. R.; Altiz, O.; Schaffner, P.; Schrader, J. H.; Blume, H. J. C.

1991-01-01

Presented here is a description of a coherent radar scattermeter and its associated signal processing hardware, which have been specifically designed to detect microbursts and record their radar characteristics. Radar parameters, signal processing techniques and detection algorithms, all under computer control, combine to sense and process reflectivity, clutter, and microburst data. Also presented is the system's high density, high data rate recording system. This digital system is capable of recording many minutes of the in-phase and quadrature components and corresponding receiver gains of the scattered returns for selected spatial regions, as well as other aircraft and hardware related parameters of interest for post-flight analysis. Information is given in viewgraph form.

Supercruiser Arrow HS-8

NASA Technical Reports Server (NTRS)

Lord, Paul; Kao, Edward; Abobo, Joey B.; Collins, Todd A.; Ma, Leong; Murad, Adnan; Naran, Hitesh; Nguyen, Thuan P.; Nuon, Timithy I.; Thomas, Dimitri D.

1992-01-01

Technology in aeronautics has advanced dramatically since the last design of a production High Speed Civil Transport (HSCT) aircraft. Newly projected requirements call for a new High Speed Civil Transport aircraft with a range of approximately 550 nm and at least 275 passenger capacity. The aircraft must be affordable and marketable. The new HSCT must be able to sustain long-duration flights and to absorb the abuse of daily operation. The new aircraft must be safe and simple to fly and require a minimum amount of maintenance. This aircraft must meet FAA certification criteria of FAR Part 25 and environmental constraints. Several design configurations were examined and two designs were selected for further investigation. The first design employs the delta planform wings and conventional empennage layout. The other design uses a swing wing layout and conventional empennage. Other engineering challenges, including materials and propulsion are also discussed. At a cruise flight speed between Mach 2.2 and Mach 3.0, no current generation of materials can endure the thermal loading of supersonic flight and satisfy the stringent weight requirements. A new generation of lightweight composite materials must be developed for the HSCT. With the enforcement of stage 3 noise restrictions, these new engines must be able to propel the aircraft and satisfy the noise limit. The engine with the most promise is the variable cycle engine. At low subsonic speeds the engine operates like a turbofan engine, providing the most efficient performance. At higher speeds the variable cycle engine operates as a turbojet power plant. The two large engine manufacturers, General Electric and Pratt & Whitney in the United States, are combining forces to make the variable cycle engine a reality.

NASA Propulsion Concept Studies and Risk Reduction Activities for Resource Prospector Lander

NASA Technical Reports Server (NTRS)

Trinh, Huu P.; Williams, Hunter; Burnside, Chris

2015-01-01

The trade study has led to the selection of propulsion concept with the lowest cost and net lowest risk -Government-owned, flight qualified components -Meet mission requirements although the configuration is not optimized. Risk reduction activities have provided an opportunity -Implement design improvements while development with the early-test approach. -Gain knowledge on the operation and identify operation limit -Data to anchor analytical models for future flight designs; The propulsion system cold flow tests series have provided valuable data for future design. -The pressure surge from the system priming and waterhammer within component operation limits. -Enable to optimize the ullage volume to reduce the propellant tank mass; RS-34 hot fire tests have successfully demonstrated of using the engines for the RP mission -No degradation of performance due to extended storage life of the hardware. -Enable to operate the engine for RP flight mission scenarios, outside of the qualification regime. -Provide extended data for the thermal and GNC designs. Significant progress has been made on NASA propulsion concept design and risk reductions for Resource Prospector lander.

Upper-Stage Flight Experiment

NASA Technical Reports Server (NTRS)

Anderson, W. E.; Boxwell, R.; Crockett, D. V.; Ross, R.; Lewis, T.; McNeal, C.; Verdarame, K.

1999-01-01

For propulsion applications that require that the propellants are storable for long periods, have a high density impulse, and are environmentally clean and non-toxic, the best choice is a combination of high-concentration hydrogen peroxide (High Test Peroxide, or HTP) and a liquid hydrocarbon (LHC) fuel. The HTP/LHC combination is suitable for low-cost launch vehicles, space taxi and space maneuvering vehicles, and kick stages. Orbital Sciences Corporation is under contract with the NASA Marshall Space Flight Center in cooperation with the Air Force Research Lab to design, develop and demonstrate a new low-cost liquid upper stage based on HTP and JP-8. The Upper Stage Flight Experiment (USFE) focuses on key technologies necessary to demonstrate the operation of an inherently simple propulsion system with an innovative, state-of-the-art structure. Two key low-cost vehicle elements will be demonstrated - a 10,000 lbf thrust engine and an integrated composite tank structure. The suborbital flight test of the USFE is scheduled for 2001. Preceding the flight tests are two major series of ground tests at NASA Stennis Space Center and a subscale tank development program to identify compatible composite materials and to verify their compatibility over long periods of time. The ground tests include a thrust chamber development test series and an integrated stage test. This paper summarizes the results from the first phase of the thrust chamber development tests and the results to date from the tank material compatibility tests. Engine and tank configurations that meet the goals of the program are described.

Computational Modeling Using OpenSim to Simulate a Squat Exercise Motion

NASA Technical Reports Server (NTRS)

Gallo, C. A.; Thompson, W. K.; Lewandowski, B. E.; Humphreys, B. T.; Funk, J. H.; Funk, N. H.; Weaver, A. S.; Perusek, G. P.; Sheehan, C. C.; Mulugeta, L.

2015-01-01

Long duration space travel to destinations such as Mars or an asteroid will expose astronauts to extended periods of reduced gravity. Astronauts will use an exercise regime for the duration of the space flight to minimize the loss of bone density, muscle mass and aerobic capacity that occurs during exposure to a reduced gravity environment. Since the area available in the spacecraft for an exercise device is limited and gravity is not present to aid loading, compact resistance exercise device prototypes are being developed. Since it is difficult to rigorously test these proposed devices in space flight, computational modeling provides an estimation of the muscle forces, joint torques and joint loads during exercise to gain insight on the efficacy to protect the musculoskeletal health of astronauts.