Optimal starting conditions for the rendezvous maneuver: Analytical and computational approach

NASA Astrophysics Data System (ADS)

Ciarcia, Marco

The three-dimensional rendezvous between two spacecraft is considered: a target spacecraft on a circular orbit around the Earth and a chaser spacecraft initially on some elliptical orbit yet to be determined. The chaser spacecraft has variable mass, limited thrust, and its trajectory is governed by three controls, one determining the thrust magnitude and two determining the thrust direction. We seek the time history of the controls in such a way that the propellant mass required to execute the rendezvous maneuver is minimized. Two cases are considered: (i) time-to-rendezvous free and (ii) time-to-rendezvous given, respectively equivalent to (i) free angular travel and (ii) fixed angular travel for the target spacecraft. The above problem has been studied by several authors under the assumption that the initial separation coordinates and the initial separation velocities are given, hence known initial conditions for the chaser spacecraft. In this paper, it is assumed that both the initial separation coordinates and initial separation velocities are free except for the requirement that the initial chaser-to-target distance is given so as to prevent the occurrence of trivial solutions. Two approaches are employed: optimal control formulation (Part A) and mathematical programming formulation (Part B). In Part A, analyses are performed with the multiple-subarc sequential gradient-restoration algorithm for optimal control problems. They show that the fuel-optimal trajectory is zero-bang, namely it is characterized by two subarcs: a long coasting zero-thrust subarc followed by a short powered max-thrust braking subarc. While the thrust direction of the powered subarc is continuously variable for the optimal trajectory, its replacement with a constant (yet optimized) thrust direction produces a very efficient guidance trajectory. Indeed, for all values of the initial distance, the fuel required by the guidance trajectory is within less than one percent of the fuel required by the optimal trajectory. For the guidance trajectory, because of the replacement of the variable thrust direction of the powered subarc with a constant thrust direction, the optimal control problem degenerates into a mathematical programming problem with a relatively small number of degrees of freedom, more precisely: three for case (i) time-to-rendezvous free and two for case (ii) time-to-rendezvous given. In particular, we consider the rendezvous between the Space Shuttle (chaser) and the International Space Station (target). Once a given initial distance SS-to-ISS is preselected, the present work supplies not only the best initial conditions for the rendezvous trajectory, but simultaneously the corresponding final conditions for the ascent trajectory. In Part B, an analytical solution of the Clohessy-Wiltshire equations is presented (i) neglecting the change of the spacecraft mass due to the fuel consumption and (ii) and assuming that the thrust is finite, that is, the trajectory includes powered subarcs flown with max thrust and coasting subarc flown with zero thrust. Then, employing the found analytical solution, we study the rendezvous problem under the assumption that the initial separation coordinates and initial separation velocities are free except for the requirement that the initial chaser-to-target distance is given. The main contribution of Part B is the development of analytical solutions for the powered subarcs, an important extension of the analytical solutions already available for the coasting subarcs. One consequence is that the entire optimal trajectory can be described analytically. Another consequence is that the optimal control problems degenerate into mathematical programming problems. A further consequence is that, vis-a-vis the optimal control formulation, the mathematical programming formulation reduces the CPU time by a factor of order 1000. Key words. Space trajectories, rendezvous, optimization, guidance, optimal control, calculus of variations, Mayer problems, Bolza problems, transformation techniques, multiple-subarc sequential gradient-restoration algorithm.

Global Optimization of Low-Thrust Interplanetary Trajectories Subject to Operational Constraints

NASA Technical Reports Server (NTRS)

Englander, Jacob Aldo; Vavrina, Matthew; Hinckley, David

2016-01-01

Low-thrust electric propulsion provides many advantages for mission to difficult targets-Comets and asteroids-Mercury-Outer planets (with sufficient power supply)Low-thrust electric propulsion is characterized by high power requirements but also very high specific impulse (Isp), leading to very good mass fractions. Low-thrust trajectory design is a very different process from chemical trajectory.

Directed energy deflection laboratory measurements of common space based targets

NASA Astrophysics Data System (ADS)

Brashears, Travis; Lubin, Philip; Hughes, Gary B.; Meinhold, Peter; Batliner, Payton; Motta, Caio; Madajian, Jonathan; Mercer, Whitaker; Knowles, Patrick

2016-09-01

We report on laboratory studies of the effectiveness of directed energy planetary defense as a part of the DE-STAR (Directed Energy System for Targeting of Asteroids and exploRation) program. DE-STAR and DE-STARLITE are directed energy "stand-off" and "stand-on" programs, respectively. These systems consist of a modular array of kilowatt-class lasers powered by photovoltaics, and are capable of heating a spot on the surface of an asteroid to the point of vaporization. Mass ejection, as a plume of evaporated material, creates a reactionary thrust capable of diverting the asteroid's orbit. In a series of papers, we have developed a theoretical basis and described numerical simulations for determining the thrust produced by material evaporating from the surface of an asteroid. In the DESTAR concept, the asteroid itself is used as the deflection "propellant". This study presents results of experiments designed to measure the thrust created by evaporation from a laser directed energy spot. We constructed a vacuum chamber to simulate space conditions, and installed a torsion balance that holds a common space target sample. The sample is illuminated with a fiber array laser with flux levels up to 60 MW/m2 , which allows us to simulate a mission level flux but on a small scale. We use a separate laser as well as a position sensitive centroid detector to readout the angular motion of the torsion balance and can thus determine the thrust. We compare the measured thrust to the models. Our theoretical models indicate a coupling coefficient well in excess of 100 μN/Woptical, though we assume a more conservative value of 80 μN/Woptical and then degrade this with an optical "encircled energy" efficiency of 0.75 to 60 μN/Woptical in our deflection modeling. Our measurements discussed here yield about 45 μN/Wabsorbed as a reasonable lower limit to the thrust per optical watt absorbed. Results vary depending on the material tested and are limited to measurements of 1 axis, so further tests must be performed.

Magnetized Target Fusion Propulsion: Plasma Injectors for MTF Guns

NASA Technical Reports Server (NTRS)

Griffin, Steven T.

2003-01-01

To achieve increased payload size and decreased trip time for interplanetary travel, a low mass, high specific impulse, high thrust propulsion system is required. This suggests the need for research into fusion as a source of power and high temperature plasma. The plasma would be deflected by magnetic fields to provide thrust. Magnetized Target Fusion (MTF) research consists of several related investigations into these topics. These include the orientation and timing of the plasma guns and the convergence and interface development of the "pusher" plasma. Computer simulations of the gun as it relates to plasma initiation and repeatability are under investigation. One of the items under development is the plasma injector. This is a surface breakdown driven plasma generator designed to function at very low pressures. The performance, operating conditions and limitations of these injectors need to be determined.

Robust H ∞ Control for Spacecraft Rendezvous with a Noncooperative Target

PubMed Central

Wu, Shu-Nan; Zhou, Wen-Ya; Tan, Shu-Jun; Wu, Guo-Qiang

2013-01-01

The robust H ∞ control for spacecraft rendezvous with a noncooperative target is addressed in this paper. The relative motion of chaser and noncooperative target is firstly modeled as the uncertain system, which contains uncertain orbit parameter and mass. Then the H ∞ performance and finite time performance are proposed, and a robust H ∞ controller is developed to drive the chaser to rendezvous with the non-cooperative target in the presence of control input saturation, measurement error, and thrust error. The linear matrix inequality technology is used to derive the sufficient condition of the proposed controller. An illustrative example is finally provided to demonstrate the effectiveness of the controller. PMID:24027446

Characterization of Space Shuttle Reusable Rocket Motor Static Test Stand Thrust Measurements

NASA Technical Reports Server (NTRS)

Cook, Mart L.; Gruet, Laurent; Cash, Stephen F. (Technical Monitor)

2003-01-01

Space Shuttle Reusable Solid Rocket Motors (RSRM) are static tested at two ATK Thiokol Propulsion facilities in Utah, T-24 and T-97. The newer T-97 static test facility was recently upgraded to allow thrust measurement capability. All previous static test motor thrust measurements have been taken at T-24; data from these tests were used to characterize thrust parameters and requirement limits for flight motors. Validation of the new T-97 thrust measurement system is required prior to use for official RSRM performance assessments. Since thrust cannot be measured on RSRM flight motors, flight motor measured chamber pressure and a nominal thrust-to-pressure relationship (based on static test motor thrust and pressure measurements) are used to reconstruct flight motor performance. Historical static test and flight motor performance data are used in conjunction with production subscale test data to predict RSRM performance. The predicted motor performance is provided to support Space Shuttle trajectory and system loads analyses. Therefore, an accurate nominal thrust-to-pressure (F/P) relationship is critical for accurate RSRM flight motor performance and Space Shuttle analyses. Flight Support Motors (FSM) 7, 8, and 9 provided thrust data for the validation of the T-97 thrust measurement system. The T-97 thrust data were analyzed and compared to thrust previously measured at T-24 to verify measured thrust data and identify any test-stand bias. The T-97 FIP data were consistent and within the T-24 static test statistical family expectation. The FSMs 7-9 thrust data met all NASA contract requirements, and the test stand is now verified for future thrust measurements.

Explicit Low-Thrust Guidance for Reference Orbit Targeting

NASA Technical Reports Server (NTRS)

Lam, Try; Udwadia, Firdaus E.

2013-01-01

The problem of a low-thrust spacecraft controlled to a reference orbit is addressed in this paper. A simple and explicit low-thrust guidance scheme with constrained thrust magnitude is developed by combining the fundamental equations of motion for constrained systems from analytical dynamics with a Lyapunov-based method. Examples are given for a spacecraft controlled to a reference trajectory in the circular restricted three body problem.

Integrated Targeting and Guidance for Powered Planetary Descent

NASA Astrophysics Data System (ADS)

Azimov, Dilmurat M.; Bishop, Robert H.

2018-02-01

This paper presents an on-board guidance and targeting design that enables explicit state and thrust vector control and on-board targeting for planetary descent and landing. These capabilities are developed utilizing a new closed-form solution for the constant thrust arc of the braking phase of the powered descent trajectory. The key elements of proven targeting and guidance architectures, including braking and approach phase quartics, are employed. It is demonstrated that implementation of the proposed solution avoids numerical simulation iterations, thereby facilitating on-board execution of targeting procedures during the descent. It is shown that the shape of the braking phase constant thrust arc is highly dependent on initial mass and propulsion system parameters. The analytic solution process is explicit in terms of targeting and guidance parameters, while remaining generic with respect to planetary body and descent trajectory design. These features increase the feasibility of extending the proposed integrated targeting and guidance design to future cargo and robotic landing missions.

Integrated Targeting and Guidance for Powered Planetary Descent

NASA Astrophysics Data System (ADS)

Azimov, Dilmurat M.; Bishop, Robert H.

2018-06-01

This paper presents an on-board guidance and targeting design that enables explicit state and thrust vector control and on-board targeting for planetary descent and landing. These capabilities are developed utilizing a new closed-form solution for the constant thrust arc of the braking phase of the powered descent trajectory. The key elements of proven targeting and guidance architectures, including braking and approach phase quartics, are employed. It is demonstrated that implementation of the proposed solution avoids numerical simulation iterations, thereby facilitating on-board execution of targeting procedures during the descent. It is shown that the shape of the braking phase constant thrust arc is highly dependent on initial mass and propulsion system parameters. The analytic solution process is explicit in terms of targeting and guidance parameters, while remaining generic with respect to planetary body and descent trajectory design. These features increase the feasibility of extending the proposed integrated targeting and guidance design to future cargo and robotic landing missions.

Thrust stand evaluation of engine performance improvement algorithms in an F-15 airplane

NASA Technical Reports Server (NTRS)

Conners, Timothy R.

1992-01-01

An investigation is underway to determine the benefits of a new propulsion system optimization algorithm in an F-15 airplane. The performance seeking control (PSC) algorithm optimizes the quasi-steady-state performance of an F100 derivative turbofan engine for several modes of operation. The PSC algorithm uses an onboard software engine model that calculates thrust, stall margin, and other unmeasured variables for use in the optimization. As part of the PSC test program, the F-15 aircraft was operated on a horizontal thrust stand. Thrust was measured with highly accurate load cells. The measured thrust was compared to onboard model estimates and to results from posttest performance programs. Thrust changes using the various PSC modes were recorded. Those results were compared to benefits using the less complex highly integrated digital electronic control (HIDEC) algorithm. The PSC maximum thrust mode increased intermediate power thrust by 10 percent. The PSC engine model did very well at estimating measured thrust and closely followed the transients during optimization. Quantitative results from the evaluation of the algorithms and performance calculation models are included with emphasis on measured thrust results. The report presents a description of the PSC system and a discussion of factors affecting the accuracy of the thrust stand load measurements.

PEG Enhancement for EM1 and EM2+ Missions

NASA Technical Reports Server (NTRS)

Von der Porten, Paul; Ahmad, Naeem; Hawkins, Matt

2018-01-01

NASA is currently building the Space Launch System (SLS) Block-1 launch vehicle for the Exploration Mission 1 (EM-1) test flight. The next evolution of SLS, the Block-1B Exploration Mission 2 (EM-2), is currently being designed. The Block-1 and Block-1B vehicles will use the Powered Explicit Guidance (PEG) algorithm. Due to the relatively low thrust-to-weight ratio of the Exploration Upper Stage (EUS), certain enhancements to the Block-1 PEG algorithm are needed to perform Block-1B missions. In order to accommodate mission design for EM-2 and beyond, PEG has been significantly improved since its use on the Space Shuttle program. The current version of PEG has the ability to switch to different targets during Core Stage (CS) or EUS flight, and can automatically reconfigure for a single Engine Out (EO) scenario, loss of communication with the Launch Abort System (LAS), and Inertial Navigation System (INS) failure. The Thrust Factor (TF) algorithm uses measured state information in addition to a priori parameters, providing PEG with an improved estimate of propulsion information. This provides robustness against unknown or undetected engine failures. A loft parameter input allows LAS jettison while maximizing payload mass. The current PEG algorithm is now able to handle various classes of missions with burn arcs much longer than were seen in the shuttle program. These missions include targeting a circular LEO orbit with a low-thrust, long-burn-duration upper stage, targeting a highly eccentric Trans-Lunar Injection (TLI) orbit, targeting a disposal orbit using the low-thrust Reaction Control System (RCS), and targeting a hyperbolic orbit. This paper will describe the design and implementation of the TF algorithm, the strategy to handle EO in various flight regimes, algorithms to cover off-nominal conditions, and other enhancements to the Block-1 PEG algorithm. This paper illustrates challenges posed by the Block-1B vehicle, and results show that the improved PEG algorithm is capable for use on the SLS Block 1-B vehicle as part of the Guidance, Navigation, and Control System.

Human Mars Ascent Vehicle Performance Sensitivities

NASA Technical Reports Server (NTRS)

Polsgrove, Tara P.; Thomas, Herbert D.

2016-01-01

Human Mars mission architecture studies have shown that the ascent vehicle mass drives performance requirements for the descent and in-space transportation elements. Understanding the sensitivity of Mars ascent vehicle (MAV) mass to various mission and vehicle design choices enables overall transportation system optimization. This paper presents the results of a variety of sensitivity trades affecting MAV performance including: landing site latitude, target orbit, initial thrust to weight ratio, staging options, specific impulse, propellant type and engine design.

Analytical study of nozzle performance for nuclear thermal rockets

NASA Technical Reports Server (NTRS)

Davidian, Kenneth O.; Kacynski, Kenneth J.

1991-01-01

Nuclear propulsion has been identified as one of the key technologies needed for human exploration of the Moon and Mars. The Nuclear Thermal Rocket (NTR) uses a nuclear reactor to heat hydrogen to a high temperature followed by expansion through a conventional convergent-divergent nozzle. A parametric study of NTR nozzles was performed using the Rocket Engine Design Expert System (REDES) at the NASA Lewis Research Center. The REDES used the JANNAF standard rigorous methodology to determine nozzle performance over a range of chamber temperatures, chamber pressures, thrust levels, and different nozzle configurations. A design condition was set by fixing the propulsion system exit radius at five meters and throat radius was varied to achieve a target thrust level. An adiabatic wall was assumed for the nozzle, and its length was assumed to be 80 percent of a 15 degree cone. The results conclude that although the performance of the NTR, based on infinite reaction rates, looks promising at low chamber pressures, finite rate chemical reactions will cause the actual performance to be considerably lower. Parameters which have a major influence on the delivered specific impulse value include the chamber temperature and the chamber pressures in the high thrust domain. Other parameters, such as 2-D and boundary layer effects, kinetic rates, and number of nozzles, affect the deliverable performance of an NTR nozzle to a lesser degree. For a single nozzle, maximum performance of 930 seconds and 1030 seconds occur at chamber temperatures of 2700 and 3100 K, respectively.

Optimal thrust level for orbit insertion

NASA Astrophysics Data System (ADS)

Cerf, Max

2017-07-01

The minimum-fuel orbital transfer is analyzed in the case of a launcher upper stage using a constantly thrusting engine. The thrust level is assumed to be constant and its value is optimized together with the thrust direction. A closed-loop solution for the thrust direction is derived from the extremal analysis for a planar orbital transfer. The optimal control problem reduces to two unknowns, namely the thrust level and the final time. Guessing and propagating the costates is no longer necessary and the optimal trajectory is easily found from a rough initialization. On the other hand the initial costates are assessed analytically from the initial conditions and they can be used as initial guess for transfers at different thrust levels. The method is exemplified on a launcher upper stage targeting a geostationary transfer orbit.

Program manual for HILTOP, a heliocentric interplanetary low thrust trajectory optimization program. Part 1: User's guide

NASA Technical Reports Server (NTRS)

Mann, F. I.; Horsewood, J. L.

1974-01-01

A performance-analysis computer program, that was developed explicitly to generate optimum electric propulsion trajectory data for missions of interest in the exploration of the solar system is presented. The program was primarily designed to evaluate the performance capabilities of electric propulsion systems, and in the simulation of a wide variety of interplanetary missions. A numerical integration of the two-body, three-dimensional equations of motion and the Euler-Lagrange equations was used in the program. Transversality conditions which permit the rapid generation of converged maximum-payload trajectory data, and the optimization of numerous other performance indices for which no transversality conditions exist are included. The ability to simulate constrained optimum solutions, including trajectories having specified propulsion time and constant thrust cone angle, is also in the program. The program was designed to handle multiple-target missions with various types of encounters, such as rendezvous, stopover, orbital capture, and flyby. Performance requirements for a variety of launch vehicles can be determined.

Evaluation of various thrust calculation techniques on an F404 engine

NASA Technical Reports Server (NTRS)

Ray, Ronald J.

1990-01-01

In support of performance testing of the X-29A aircraft at the NASA-Ames, various thrust calculation techniques were developed and evaluated for use on the F404-GE-400 engine. The engine was thrust calibrated at NASA-Lewis. Results from these tests were used to correct the manufacturer's in-flight thrust program to more accurately calculate thrust for the specific test engine. Data from these tests were also used to develop an independent, simplified thrust calculation technique for real-time thrust calculation. Comparisons were also made to thrust values predicted by the engine specification model. Results indicate uninstalled gross thrust accuracies on the order of 1 to 4 percent for the various in-flight thrust methods. The various thrust calculations are described and their usage, uncertainty, and measured accuracies are explained. In addition, the advantages of a real-time thrust algorithm for flight test use and the importance of an accurate thrust calculation to the aircraft performance analysis are described. Finally, actual data obtained from flight test are presented.

Pulsed Electric Propulsion Thrust Stand Calibration Method

NASA Technical Reports Server (NTRS)

Wong, Andrea R.; Polzin, Kurt A.; Pearson, J. Boise

2011-01-01

The evaluation of the performance of any propulsion device requires the accurate measurement of thrust. While chemical rocket thrust is typically measured using a load cell, the low thrust levels associated with electric propulsion (EP) systems necessitate the use of much more sensitive measurement techniques. The design and development of electric propulsion thrust stands that employ a conventional hanging pendulum arm connected to a balance mechanism consisting of a secondary arm and variable linkage have been reported in recent publications by Polzin et al. These works focused on performing steady-state thrust measurements and employed a static analysis of the thrust stand response. In the present work, we present a calibration method and data that will permit pulsed thrust measurements using the Variable Amplitude Hanging Pendulum with Extended Range (VAHPER) thrust stand. Pulsed thrust measurements are challenging in general because the pulsed thrust (impulse bit) occurs over a short timescale (typically 1 micros to 1 millisecond) and cannot be resolved directly. Consequently, the imparted impulse bit must be inferred through observation of the change in thrust stand motion effected by the pulse. Pulsed thrust measurements have typically only consisted of single-shot operation. In the present work, we discuss repetition-rate pulsed thruster operation and describe a method to perform these measurements. The thrust stand response can be modeled as a spring-mass-damper system with a repetitive delta forcing function to represent the impulsive action of the thruster.

Thrust Performance Evaluation of a Turbofan Engine Based on Exergetic Approach and Thrust Management in Aircraft

NASA Astrophysics Data System (ADS)

Yalcin, Enver

2017-05-01

The environmental parameters such as temperature and air pressure which are changing depending on altitudes are effective on thrust and fuel consumption of aircraft engines. In flights with long routes, thrust management function in airplane information system has a structure that ensures altitude and performance management. This study focused on thrust changes throughout all flight were examined by taking into consideration their energy and exergy performances for fuel consumption of an aircraft engine used in flight with long route were taken as reference. The energetic and exergetic performance evaluations were made under the various altitude conditions. The thrust changes for different altitude conditions were obtained to be at 86.53 % in descending direction and at 142.58 % in ascending direction while the energy and exergy efficiency changes for the referenced engine were found to be at 80.77 % and 84.45 %, respectively. The results revealed here can be helpful to manage thrust and reduce fuel consumption, but engine performance will be in accordance with operation requirements.

Simplified procedures for correlation of experimentally measured and predicted thrust chamber performance

NASA Technical Reports Server (NTRS)

Powell, W. B.

1973-01-01

Thrust chamber performance is evaluated in terms of an analytical model incorporating all the loss processes that occur in a real rocket motor. The important loss processes in the real thrust chamber were identified, and a methodology and recommended procedure for predicting real thrust chamber vacuum specific impulse were developed. Simplified equations for the calculation of vacuum specific impulse are developed to relate the delivered performance (both vacuum specific impulse and characteristic velocity) to the ideal performance as degraded by the losses corresponding to a specified list of loss processes. These simplified equations enable the various performance loss components, and the corresponding efficiencies, to be quantified separately (except that interaction effects are arbitrarily assigned in the process). The loss and efficiency expressions presented can be used to evaluate experimentally measured thrust chamber performance, to direct development effort into the areas most likely to yield improvements in performance, and as a basis to predict performance of related thrust chamber configurations.

The results of a low-speed wind tunnel test to investigate the effects of the Refan JT8D engine target thrust reverser on the stability and control characteristics of the Boeing 727-200 airplane

NASA Technical Reports Server (NTRS)

Kupcis, E. A.

1974-01-01

The effects of the Refan JT8D side engine target thrust reverser on the stability and control characteristics of the Boeing 727-200 airplane were investigated using the Boeing-Vertol 20 x 20 ft Low-Speed Wind Tunnel. A powered model of the 727-200 was tested in groud effect in the landing configuration. The Refan target reverser configuration was evaluated relative to the basic production 727 airplane with its clamshell-deflector door thrust reverser design. The Refan configuration had slightly improved directional control characteristics relative to the basic airplane. Clocking the Refan thrust reversers 20 degrees outboard to direct the reverser flow away from the vertical tail, had little effect on directional control. However, clocking them 20 degrees inboard resulted in a complete loss of rudder effectiveness for speeds greater than 90 knots. Variations in Refan reverser lip/fence geometry had a minor effect on directional control.

Measurement of Impulsive Thrust from a Closed Radio Frequency Cavity in Vacuum

NASA Technical Reports Server (NTRS)

White, Harold; March, Paul; Lawrence, James; Vera, Jerry; Sylvester, Andre; Brady, David; Bailey, Paul

2016-01-01

A vacuum test campaign evaluating the impulsive thrust performance of a tapered RF test article excited in the TM212 mode at 1,937 megahertz (MHz) has been completed. The test campaign consisted of a forward thrust phase and reverse thrust phase at less than 8 x 10(exp -6) Torr vacuum with power scans at 40 watts, 60 watts, and 80 watts. The test campaign included a null thrust test effort to identify any mundane sources of impulsive thrust, however none were identified. Thrust data from forward, reverse, and null suggests that the system is consistently performing with a thrust to power ratio of 1.2 +/- 0.1 mN/kW.

Thrust distribution for attitude control in a variable thrust propulsion system with four ACS nozzles

NASA Astrophysics Data System (ADS)

Lim, Yeerang; Lee, Wonsuk; Bang, Hyochoong; Lee, Hosung

2017-04-01

A thrust distribution approach is proposed in this paper for a variable thrust solid propulsion system with an attitude control system (ACS) that uses a reduced number of nozzles for a three-axis attitude maneuver. Although a conventional variable thrust solid propulsion system needs six ACS nozzles, this paper proposes a thrust system with four ACS nozzles to reduce the complexity and mass of the system. The performance of the new system was analyzed with numerical simulations, and the results show that the performance of the system with four ACS nozzles was similar to the original system while the mass of the whole system was simultaneously reduced. Moreover, a feasibility analysis was performed to determine whether a thrust system with three ACS nozzles is possible.

Experimental Study of an Axisymmetric Dual Throat Fluidic Thrust Vectoring Nozzle for Supersonic Aircraft Application

NASA Technical Reports Server (NTRS)

Flamm, Jeffrey D.; Deere, Karen A.; Mason, Mary L.; Berrier, Bobby L.; Johnson, Stuart K.

2007-01-01

An axisymmetric version of the Dual Throat Nozzle concept with a variable expansion ratio has been studied to determine the impacts on thrust vectoring and nozzle performance. The nozzle design, applicable to a supersonic aircraft, was guided using the unsteady Reynolds-averaged Navier-Stokes computational fluid dynamics code, PAB3D. The axisymmetric Dual Throat Nozzle concept was tested statically in the Jet Exit Test Facility at the NASA Langley Research Center. The nozzle geometric design variables included circumferential span of injection, cavity length, cavity convergence angle, and nozzle expansion ratio for conditions corresponding to take-off and landing, mid climb and cruise. Internal nozzle performance and thrust vectoring performance was determined for nozzle pressure ratios up to 10 with secondary injection rates up to 10 percent of the primary flow rate. The 60 degree span of injection generally performed better than the 90 degree span of injection using an equivalent injection area and number of holes, in agreement with computational results. For injection rates less than 7 percent, thrust vector angle for the 60 degree span of injection was 1.5 to 2 degrees higher than the 90 degree span of injection. Decreasing cavity length improved thrust ratio and discharge coefficient, but decreased thrust vector angle and thrust vectoring efficiency. Increasing cavity convergence angle from 20 to 30 degrees increased thrust vector angle by 1 degree over the range of injection rates tested, but adversely affected system thrust ratio and discharge coefficient. The dual throat nozzle concept generated the best thrust vectoring performance with an expansion ratio of 1.0 (a cavity in between two equal minimum areas). The variable expansion ratio geometry did not provide the expected improvements in discharge coefficient and system thrust ratio throughout the flight envelope of typical a supersonic aircraft. At mid-climb and cruise conditions, the variable geometry design compromised thrust vector angle achieved, but some thrust vector control would be available, potentially for aircraft trim. The fixed area, expansion ratio of 1.0, Dual Throat Nozzle provided the best overall compromise for thrust vectoring and nozzle internal performance over the range of NPR tested compared to the variable geometry Dual Throat Nozzle.

Reverse thrust performance of the QCSEE variable pitch turbofan engine

NASA Technical Reports Server (NTRS)

Samanich, N. E.; Reemsnyder, D. C.; Blodmer, H. E.

1980-01-01

Results of steady state reverse and forward to reverse thrust transient performance tests are presented. The original quiet, clean, short haul, experimental engine four segment variable fan nozzle was retested in reverse and compared with a continuous, 30 deg half angle conical exlet. Data indicated that the significantly more stable, higher pressure recovery flow with the fixed 30 deg exlet resulted in lower engine vibrations, lower fan blade stress, and approximately a 20 percent improvement in reverse thrust. Objective reverse thrust of 35 percent of takeoff thrust was reached. Thrust response of less than 1.5 sec was achieved for the approach and the takeoff to reverse thrust transients.

Automated low-thrust guidance for the orbital maneuvering vehicle

NASA Technical Reports Server (NTRS)

Rose, Richard E.; Schmeichel, Harry; Shortwell, Charles P.; Werner, Ronald A.

1988-01-01

This paper describes the highly autonomous OMV Guidance Navigation and Control system. Emphasis is placed on a key feature of the design, the low thrust guidance algorithm. The two guidance modes, orbit change guidance and rendezvous guidance, are discussed in detail. It is shown how OMV will automatically transfer from its initial orbit to an arbitrary target orbit and reach a specified rendezvous position relative to the target vehicle.

Performance characteristics of two multiaxis thrust-vectoring nozzles at Mach numbers up to 1.28

NASA Technical Reports Server (NTRS)

Wing, David J.; Capone, Francis J.

1993-01-01

The thrust-vectoring axisymmetric (VA) nozzle and a spherical convergent flap (SCF) thrust-vectoring nozzle were tested along with a baseline nonvectoring axisymmetric (NVA) nozzle in the Langley 16-Foot Transonic Tunnel at Mach numbers from 0 to 1.28 and nozzle pressure ratios from 1 to 8. Test parameters included geometric yaw vector angle and unvectored divergent flap length. No pitch vectoring was studied. Nozzle drag, thrust minus drag, yaw thrust vector angle, discharge coefficient, and static thrust performance were measured and analyzed, as well as external static pressure distributions. The NVA nozzle and the VA nozzle displayed higher static thrust performance than the SCF nozzle throughout the nozzle pressure ratio (NPR) range tested. The NVA nozzle had higher overall thrust minus drag than the other nozzles throughout the NPR and Mach number ranges tested. The SCF nozzle had the lowest jet-on nozzle drag of the three nozzles throughout the test conditions. The SCF nozzle provided yaw thrust angles that were equal to the geometric angle and constant with NPR. The VA nozzle achieved yaw thrust vector angles that were significantly higher than the geometric angle but not constant with NPR. Nozzle drag generally increased with increases in thrust vectoring for all the nozzles tested.

Low Thrust Cis-Lunar Transfers Using a 40 kW-Class Solar Electric Propulsion Spacecraft

NASA Technical Reports Server (NTRS)

Mcguire, Melissa L.; Burke, Laura M.; Mccarty, Steven L.; Hack, Kurt J.; Whitley, Ryan J.; Davis, Diane C.; Ocampo, Cesar

2017-01-01

This paper captures trajectory analysis of a representative low thrust, high power Solar Electric Propulsion (SEP) vehicle to move a mass around cis-lunar space in the range of 20 to 40 kW power to the Electric Propulsion (EP) system. These cis-lunar transfers depart from a selected Near Rectilinear Halo Orbit (NRHO) and target other cis-lunar orbits. The NRHO cannot be characterized in the classical two-body dynamics more familiar in the human spaceflight community, and the use of low thrust orbit transfers provides unique analysis challenges. Among the target orbit destinations documented in this paper are transfers between a Southern and Northern NRHO, transfers between the NRHO and a Distant Retrograde Orbit (DRO) and a transfer between the NRHO and two different Earth Moon Lagrange Point 2 (EML2) Halo orbits. Because many different NRHOs and EML2 halo orbits exist, simplifying assumptions rely on previous analysis of orbits that meet current abort and communication requirements for human mission planning. Investigation is done into the sensitivities of these low thrust transfers to EP system power. Additionally, the impact of the Thrust to Weight ratio of these low thrust SEP systems and the ability to transit between these unique orbits are investigated.

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.

Evaluation of an Outer Loop Retrofit Architecture for Intelligent Turbofan Engine Thrust Control

NASA Technical Reports Server (NTRS)

Litt, Jonathan S.; Sowers, T. Shane

2006-01-01

The thrust control capability of a retrofit architecture for intelligent turbofan engine control and diagnostics is evaluated. The focus of the study is on the portion of the hierarchical architecture that performs thrust estimation and outer loop thrust control. The inner loop controls fan speed so the outer loop automatically adjusts the engine's fan speed command to maintain thrust at the desired level, based on pilot input, even as the engine deteriorates with use. The thrust estimation accuracy is assessed under nominal and deteriorated conditions at multiple operating points, and the closed loop thrust control performance is studied, all in a complex real-time nonlinear turbofan engine simulation test bed. The estimation capability, thrust response, and robustness to uncertainty in the form of engine degradation are evaluated.

Computational Study of an Axisymmetric Dual Throat Fluidic Thrust Vectoring Nozzle for a Supersonic Aircraft Application

NASA Technical Reports Server (NTRS)

Deere, Karen A.; Flamm, Jeffrey D.; Berrier, Bobby L.; Johnson, Stuart K.

2007-01-01

A computational investigation of an axisymmetric Dual Throat Nozzle concept has been conducted. This fluidic thrust-vectoring nozzle was designed with a recessed cavity to enhance the throat shifting technique for improved thrust vectoring. The structured-grid, unsteady Reynolds- Averaged Navier-Stokes flow solver PAB3D was used to guide the nozzle design and analyze performance. Nozzle design variables included extent of circumferential injection, cavity divergence angle, cavity length, and cavity convergence angle. Internal nozzle performance (wind-off conditions) and thrust vector angles were computed for several configurations over a range of nozzle pressure ratios from 1.89 to 10, with the fluidic injection flow rate equal to zero and up to 4 percent of the primary flow rate. The effect of a variable expansion ratio on nozzle performance over a range of freestream Mach numbers up to 2 was investigated. Results indicated that a 60 circumferential injection was a good compromise between large thrust vector angles and efficient internal nozzle performance. A cavity divergence angle greater than 10 was detrimental to thrust vector angle. Shortening the cavity length improved internal nozzle performance with a small penalty to thrust vector angle. Contrary to expectations, a variable expansion ratio did not improve thrust efficiency at the flight conditions investigated.

Static Thrust and Vectoring Performance of a Spherical Convergent Flap Nozzle with a Nonrectangular Divergent Duct

NASA Technical Reports Server (NTRS)

Wing, David J.

1998-01-01

The static internal performance of a multiaxis-thrust-vectoring, spherical convergent flap (SCF) nozzle with a non-rectangular divergent duct was obtained in the model preparation area of the Langley 16-Foot Transonic Tunnel. Duct cross sections of hexagonal and bowtie shapes were tested. Additional geometric parameters included throat area (power setting), pitch flap deflection angle, and yaw gimbal angle. Nozzle pressure ratio was varied from 2 to 12 for dry power configurations and from 2 to 6 for afterburning power configurations. Approximately a 1-percent loss in thrust efficiency from SCF nozzles with a rectangular divergent duct was incurred as a result of internal oblique shocks in the flow field. The internal oblique shocks were the result of cross flow generated by the vee-shaped geometric throat. The hexagonal and bowtie nozzles had mirror-imaged flow fields and therefore similar thrust performance. Thrust vectoring was not hampered by the three-dimensional internal geometry of the nozzles. Flow visualization indicates pitch thrust-vector angles larger than 10' may be achievable with minimal adverse effect on or a possible gain in resultant thrust efficiency as compared with the performance at a pitch thrust-vector angle of 10 deg.

Multiphysics Nuclear Thermal Rocket Thrust Chamber Analysis

NASA Technical Reports Server (NTRS)

Wang, Ten-See

2005-01-01

The objective of this effort is t o develop an efficient and accurate thermo-fluid computational methodology to predict environments for hypothetical thrust chamber design and analysis. The current task scope is to perform multidimensional, multiphysics analysis of thrust performance and heat transfer analysis for a hypothetical solid-core, nuclear thermal engine including thrust chamber and nozzle. The multiphysics aspects of the model include: real fluid dynamics, chemical reactivity, turbulent flow, and conjugate heat transfer. The model will be designed to identify thermal, fluid, and hydrogen environments in all flow paths and materials. This model would then be used to perform non- nuclear reproduction of the flow element failures demonstrated in the Rover/NERVA testing, investigate performance of specific configurations and assess potential issues and enhancements. A two-pronged approach will be employed in this effort: a detailed analysis of a multi-channel, flow-element, and global modeling of the entire thrust chamber assembly with a porosity modeling technique. It is expected that the detailed analysis of a single flow element would provide detailed fluid, thermal, and hydrogen environments for stress analysis, while the global thrust chamber assembly analysis would promote understanding of the effects of hydrogen dissociation and heat transfer on thrust performance. These modeling activities will be validated as much as possible by testing performed by other related efforts.

Experimental thrust performance of a high-area-ratio rocket nozzle

NASA Technical Reports Server (NTRS)

Pavli, Albert J.; Kacynski, Kenneth J.; Smith, Tamara A.

1987-01-01

An experimental investigation was conducted to determine the thrust performance attainable from high-area-ratio rocket nozzles. A modified Rao-contoured nozzle with an expansion area of 1030 was test fired with hydrogen-oxygen propellants at altitude conditions. The nozzle was also tested as a truncated nozzle, at an expansion area ratio of 428. Thrust coefficient and thrust coefficient efficiency values are presented for each configuration at various propellant mixture ratios (oxygen/fuel). Several procedural techniques were developed permitting improved measurement of nozzle performance. The more significant of these were correcting the thrust for the aneroid effects, determining the effective chamber pressure, and referencing differential pressure transducers to a vacuum reference tank.

Experimental thrust performance of a high area-ratio rocket nozzle

NASA Technical Reports Server (NTRS)

Pavli, A. J.; Kacynski, K. J.; Smith, T. A.

1986-01-01

An experimental investigation was conducted to determine the thrust performance attainable from high-area-ratio rocket nozzles. A modified Rao-contoured nozzle with an expansion area of 1030 was test fired with hydrogen-oxygen propellants at altitude conditions. The nozzle was also tested as a truncated nozzle, at an expansion area ratio of 428. Thrust coefficient and thrust coefficient efficiency values are presented for each configuration at various propellant mixture ratios (oxygen/fuel). Several procedural techniques were developed permitting improved measurement of nozzle performance. The more significant of these were correcting the thrust for the aneroid effects, determining the effective chamber pressure, and referencing differential pressure transducers to a vacuum reference tank.

Rapid prototype fabrication processes for high-performance thrust cells

NASA Technical Reports Server (NTRS)

Hunt, K.; Chwiedor, T.; Diab, J.; Williams, R.

1994-01-01

The Thrust Cell Technologies Program (Air Force Phillips Laboratory Contract No. F04611-92-C-0050) is currently being performed by Rocketdyne to demonstrate advanced materials and fabrication technologies which can be utilized to produce low-cost, high-performance thrust cells for launch and space transportation rocket engines. Under Phase 2 of the Thrust Cell Technologies Program (TCTP), rapid prototyping and investment casting techniques are being employed to fabricate a 12,000-lbf thrust class combustion chamber for delivery and hot-fire testing at Phillips Lab. The integrated process of investment casting directly from rapid prototype patterns dramatically reduces design-to-delivery cycle time, and greatly enhances design flexibility over conventionally processed cast or machined parts.

Influence of Structural Parameters on the Performance of Vortex Valve Variable-Thrust Solid Rocket Motor

NASA Astrophysics Data System (ADS)

Wei, Xianggeng; Li, Jiang; He, Guoqiang

2017-04-01

The vortex valve solid variable thrust motor is a new solid motor which can achieve Vehicle system trajectory optimization and motor energy management. Numerical calculation was performed to investigate the influence of vortex chamber diameter, vortex chamber shape, and vortex chamber height of the vortex valve solid variable thrust motor on modulation performance. The test results verified that the calculation results are consistent with laboratory results with a maximum error of 9.5%. The research drew the following major conclusions: the optimal modulation performance was achieved in a cylindrical vortex chamber, increasing the vortex chamber diameter improved the modulation performance of the vortex valve solid variable thrust motor, optimal modulation performance could be achieved when the height of the vortex chamber is half of the vortex chamber outlet diameter, and the hot gas control flow could result in an enhancement of modulation performance. The results can provide the basis for establishing the design method of the vortex valve solid variable thrust motor.

Static internal performance of single expansion-ramp nozzles with thrust vectoring and reversing

NASA Technical Reports Server (NTRS)

Re, R. J.; Berrier, B. L.

1982-01-01

The effects of geometric design parameters on the internal performance of nonaxisymmetric single expansion-ramp nozzles were investigated at nozzle pressure ratios up to approximately 10. Forward-flight (cruise), vectored-thrust, and reversed-thrust nozzle operating modes were investigated.

High-speed engine/component performance assessment using exergy and thrust-based methods

NASA Technical Reports Server (NTRS)

Riggins, D. W.

1996-01-01

This investigation summarizes a comparative study of two high-speed engine performance assessment techniques based on energy (available work) and thrust-potential (thrust availability). Simple flow-fields utilizing Rayleigh heat addition and one-dimensional flow with friction are used to demonstrate the fundamental inability of conventional energy techniques to predict engine component performance, aid in component design, or accurately assess flow losses. The use of the thrust-based method on these same examples demonstrates its ability to yield useful information in all these categories. Energy and thrust are related and discussed from the stand-point of their fundamental thermodynamic and fluid dynamic definitions in order to explain the differences in information obtained using the two methods. The conventional definition of energy is shown to include work which is inherently unavailable to an aerospace Brayton engine. An engine-based energy is then developed which accurately accounts for this inherently unavailable work; performance parameters based on this quantity are then shown to yield design and loss information equivalent to the thrust-based method.

A piloted simulation investigation of the normal load factor and longitudinal thrust required for air-to-air acquisition and tracking

NASA Technical Reports Server (NTRS)

Whalley, Matthew S.

1993-01-01

A piloted simulation study was performed by the U.S. Army Aeroflighydynamics Directorate to develop insight into the maneuverability requirements for aggressive helicopter maneuvering tasks such as air-to-air combat. Both a conventional helicopter and a helicopter with auxiliary thrust were examined. The aircraft parameters of interest were the normal and longitudinal load factor envelopes. Of particular interest were the mission performance and handling qualities tradeoffs with the parameters of interest. Two air-to-air acquisition and tracking tasks and a return-to-cover task were performed to assess mission performance. Results indicate that without auxiliary thrust, the ownship normal load factor capability needs to match that of the adversary in order to provide satisfactory handling qualities. Auxiliary thrust provides significant handling qualities advantages and can be substituted to some extent for normal load factor capability. Auxiliary thrust levels as low as 0.2 thrust/weight can provide significant handling qualities advantages.

Selected Performance Measurements of the F-15 Active Axisymmetric Thrust-vectoring Nozzle

NASA Technical Reports Server (NTRS)

Orme, John S.; Sims, Robert L.

1998-01-01

Flight tests recently completed at the NASA Dryden Flight Research Center evaluated performance of a hydromechanically vectored axisymmetric nozzle onboard the F-15 ACTIVE. A flight-test technique whereby strain gages installed onto engine mounts provided for the direct measurement of thrust and vector forces has proven to be extremely valuable. Flow turning and thrust efficiency, as well as nozzle static pressure distributions were measured and analyzed. This report presents results from testing at an altitude of 30,000 ft and a speed of Mach 0.9. Flow turning and thrust efficiency were found to be significantly different than predicted, and moreover, varied substantially with power setting and pitch vector angle. Results of an in-flight comparison of the direct thrust measurement technique and an engine simulation fell within the expected uncertainty bands. Overall nozzle performance at this flight condition demonstrated the F100-PW-229 thrust-vectoring nozzles to be highly capable and efficient.

Performance characteristics of a wedge nozzle installed on an F-18 propulsion wind tunnel model

NASA Technical Reports Server (NTRS)

Petit, J. E.; Capone, F. J.

1979-01-01

The results of two-dimensional wedge non-axisymmetric nozzle (2D-AIN) tests to determine its performance relative to the baseline axisymmetric nozzle using an F-18 jet effects wind tunnel model are presented. Configurations and test conditions simulated forward thrust-minus drag, thrust vectoring/induced lift, and thrust reversing flight conditions from Mach .6 to 1.20 and attack angles up to 10 degrees. Results of the model test program indicate that non-axisymmetric nozzles can be installed on a twin engine fighter aircraft model with equivalent thrust minus drag performance as the baseline axisymmetric nozzles. Thrust vectoring capability of the non-axisymmetric nozzles provided significant jet-induced lift on the nozzle/aftbody and horizontal tail surfaces. Thrust reversing panels deployed from the 2D-AIN centerbody wedge were very effective for static and inflight operation

Selected Performance Measurements of the F-15 ACTIVE Axisymmetric Thrust-Vectoring Nozzle

NASA Technical Reports Server (NTRS)

Orme, John S.; Sims, Robert L.

1999-01-01

Flight tests recently completed at the NASA Dryden Flight Research Center evaluated performance of a hydromechanically vectored axisymmetric nozzle onboard the F-15 ACTIVE. A flight-test technique whereby strain gages installed onto engine mounts provided for the direct measurement of thrust and vector forces has proven to be extremely valuable. Flow turning and thrust efficiency, as well as nozzle static pressure distributions were measured and analyzed. This report presents results from testing at an altitude of 30,000 ft and a speed of Mach 0.9. Flow turning and thrust efficiency were found to be significantly different than predicted, and moreover, varied substantially with power setting and pitch vector angle. Results of an in-flight comparison of the direct thrust measurement technique and an engine simulation fell within the expected uncertainty bands. Overall nozzle performance at this flight condition demonstrated the F100-PW-229 thrust-vectoring nozzles to be highly capable and efficient.

Internal performance characteristics of thrust-vectored axisymmetric ejector nozzles

NASA Technical Reports Server (NTRS)

Lamb, Milton

1995-01-01

A series of thrust-vectored axisymmetric ejector nozzles were designed and experimentally tested for internal performance and pumping characteristics at the Langley research center. This study indicated that discontinuities in the performance occurred at low primary nozzle pressure ratios and that these discontinuities were mitigated by decreasing expansion area ratio. The addition of secondary flow increased the performance of the nozzles. The mid-to-high range of secondary flow provided the most overall improvements, and the greatest improvements were seen for the largest ejector area ratio. Thrust vectoring the ejector nozzles caused a reduction in performance and discharge coefficient. With or without secondary flow, the vectored ejector nozzles produced thrust vector angles that were equivalent to or greater than the geometric turning angle. With or without secondary flow, spacing ratio (ejector passage symmetry) had little effect on performance (gross thrust ratio), discharge coefficient, or thrust vector angle. For the unvectored ejectors, a small amount of secondary flow was sufficient to reduce the pressure levels on the shroud to provide cooling, but for the vectored ejector nozzles, a larger amount of secondary air was required to reduce the pressure levels to provide cooling.

Trajectory design for a rendezvous mission to Earth's Trojan asteroid 2010 TK7

NASA Astrophysics Data System (ADS)

Lei, Hanlun; Xu, Bo; Zhang, Lei

2017-12-01

In this paper a rendezvous mission to the Earth's Trojan asteroid 2010 TK7 is proposed, and preliminary transfer trajectories are designed. Due to the high inclination (∼ 20.9°) of the target asteroid relative to the ecliptic plane, direct transfers usually require large amounts of fuel consumption, which is beyond the capacity of current technology. As gravity assist technique could effectively change the inclination of spacecraft's trajectory, it is adopted to reduce the launch energy and rendezvous velocity maneuver. In practical computation, impulsive and low-thrust, gravity-assisted trajectories are considered. Among all the trajectories computed, the low-thrust gravity-assisted trajectory with Venus-Earth-Venus (V-E-V) swingby sequence performs the best in terms of propellant mass. For a spacecraft with initial mass of 800 kg , propellant mass of the best trajectory is 36.74 kg . Numerical results indicate that both the impulsive and low-thrust, gravity-assisted trajectories corresponding to V-E-V sequence could satisfy mission constraints, and can be applied to practical rendezvous mission.

Tests of a D vented thrust deflecting nozzle behind a simulated turbofan engine

NASA Technical Reports Server (NTRS)

Watson, T. L.

1982-01-01

A D vented thrust deflecting nozzle applicable to subsonic V/STOL aircraft was tested behind a simulated turbofan engine in the verticle thrust stand. Nozzle thrust, fan operating characteristics, nozzle entrance conditions, and static pressures were measured. Nozzle performance was measured for variations in exit area and thrust deflection angle. Six core nozzle configurations, the effect of core exit axial location, mismatched core and fan stream nozzle pressure ratios, and yaw vane presence were evaluated. Core nozzle configuration affected performance at normal and engine out operating conditions. Highest vectored nozzle performance resulted for a given exit area when core and fan stream pressure were equal. Its is concluded that high nozzle performance can be maintained at both normal and engine out conditions through control of the nozzle entrance Mach number with a variable exit area.

Extended performance solar electric propulsion thrust system study. Volume 2: Baseline thrust system

NASA Technical Reports Server (NTRS)

Poeschel, R. L.; Hawthorne, E. I.

1977-01-01

Several thrust system design concepts were evaluated and compared using the specifications of the most advanced 30- cm engineering model thruster as the technology base. Emphasis was placed on relatively high-power missions (60 to 100 kW) such as a Halley's comet rendezvous. The extensions in thruster performance required for the Halley's comet mission were defined and alternative thrust system concepts were designed in sufficient detail for comparing mass, efficiency, reliability, structure, and thermal characteristics. Confirmation testing and analysis of thruster and power-processing components were performed, and the feasibility of satisfying extended performance requirements was verified. A baseline design was selected from the alternatives considered, and the design analysis and documentation were refined. The baseline thrust system design features modular construction, conventional power processing, and a concentractor solar array concept and is designed to interface with the space shuttle.

Rotor Hover Performance and Flowfield Measurements with Untwisted and Highly-Twisted Blades

NASA Technical Reports Server (NTRS)

Ramasamy, Manikandan; Gold, Nili P.; Bhagwat, Mahendra J.

2010-01-01

The flowfield and performance characteristics of highly-twisted blades were analyzed at various thrust conditions to improve the fundamental understanding relating the wake effects on rotor performance. Similar measurements made using untwisted blades served as the baseline case. Twisted blades are known to give better hover performance than untwisted blades at high thrust coefficients typical of those found in full-scale rotors. However, the present experiments were conducted at sufficiently low thrust (beginning from zero thrust), where the untwisted blades showed identical, if not better, performance when compared with the highly-twisted blades. The flowfield measurements showed some key wake differences between the two rotors, as well. These observations when combined with simple blade element momentum theory (also called annular disk momentum theory) helped further the understanding of rotor performance characteristics.

Static performance and noise tests on a thrust reverser for an augmentor wing aircraft

NASA Technical Reports Server (NTRS)

Harkonen, D. L.; Marrs, C. C.; Okeefe, J. V.

1974-01-01

A 1/3 scale model static test program was conducted to measure the noise levels and reverse thrust performance characteristics of wing-mounted thrust reverser that could be used on an advanced augmentor wing airplane. The configuration tested represents only the most fundamental designs where installation and packaging restraints are not considered. The thrust reverser performance is presented in terms of horizontal, vertical, and resultant effectiveness ratios and the reverser noise is compared on the basis of peak perceived noise level (PNL) and one-third octave band data (OASPL). From an analysis of the model force and acoustic data, an assessment is made on the stopping distance versus noise for a 90,900 kg (200,000 lb) airplane using this type of thrust reverser.

Cold seeps and splay faults on Nankai margin

NASA Astrophysics Data System (ADS)

Henry, P.; Ashi, J.; Tsunogai, U.; Toki, T.; Kuramoto, S.; Kinoshita, M.; Lallemant, S. J.

2003-04-01

Cold seeps (bacterial mats, specific fauna, authigenic carbonates) are common on the Nankai margin and considered as evidence for seepage of methane bearing fluids. Camera and submersible surveys performed over the years have shown that cold seeps are generally associated with active faults. One question is whether part of the fluids expelled originate from the seismogenic zone and migrate along splay faults to the seafloor. The localisation of most cold seeps on the hanging wall of major thrusts may, however, be interpreted in various ways: (a) footwall compaction and diffuse flow (b) fluid channelling along the fault zone at depths and diffuse flow near the seafloor (c) erosion and channelling along permeable strata. In 2002, new observations and sampling were performed with submersible and ROV (1) on major thrusts along the boundary between the Kumano forearc basin domain and the accretionary wedge domain, (2) on a fault affecting the forearc (Kodaiba fault), (3) on mud volcanoes in the Kumano basin. In area (1) tsunami and seismic inversions indicate that the targeted thrusts are in the slip zone of the To-Nankai 1944 earthquakes. In this area, the largest seep zone, continuous over at least 2 km, coincides with the termination of a thrust trace, indicating local fluid channelling along the edge of the fault zone. Kodaiba fault is part of another splay fault system, which has both thrusting and strike-slip components and terminates westward into an en-echelon fold system. Strong seepage activity with abundant carbonates was found on a fold at the fault termination. One mud volcano, rooted in one of the en-echelon fold, has exceptionally high seepage activity compared with the others and thick carbonate crusts. These observations suggest that fluid expulsion along fault zones is most active at fault terminations and may be enhanced during fault initiation. Preliminary geochemical results indicate signatures differ between seep sites and suggests that the two fault systems tap in different sources.

Effects of internal yaw-vectoring devices on the static performance of a pitch-vectoring nonaxisymmetric convergent-divergent nozzle

NASA Technical Reports Server (NTRS)

Asbury, Scott C.

1993-01-01

An investigation was conducted in the static test facility of the Langley 16-Foot Transonic Tunnel to evaluate the internal performance of a nonaxisymmetric convergent divergent nozzle designed to have simultaneous pitch and yaw thrust vectoring capability. This concept utilized divergent flap deflection for thrust vectoring in the pitch plane and flow-turning deflectors installed within the divergent flaps for yaw thrust vectoring. Modifications consisting of reducing the sidewall length and deflecting the sidewall outboard were investigated as means to increase yaw-vectoring performance. This investigation studied the effects of multiaxis (pitch and yaw) thrust vectoring on nozzle internal performance characteristics. All tests were conducted with no external flow, and nozzle pressure ratio was varied from 2.0 to approximately 13.0. The results indicate that this nozzle concept can successfully generate multiaxis thrust vectoring. Deflection of the divergent flaps produced resultant pitch vector angles that, although dependent on nozzle pressure ratio, were nearly equal to the geometric pitch vector angle. Losses in resultant thrust due to pitch vectoring were small or negligible. The yaw deflectors produced resultant yaw vector angles up to 21 degrees that were controllable by varying yaw deflector rotation. However, yaw deflector rotation resulted in significant losses in thrust ratios and, in some cases, nozzle discharge coefficient. Either of the sidewall modifications generally reduced these losses and increased maximum resultant yaw vector angle. During multiaxis (simultaneous pitch and yaw) thrust vectoring, little or no cross coupling between the thrust vectoring processes was observed.

Static internal performance characteristics of two thrust reverser concepts for axisymmetric nozzles

NASA Technical Reports Server (NTRS)

Leavitt, L. D.; Re, R. J.

1982-01-01

The statis performance of two axisymmetric nozzle thrust reverser concepts was investigated. A rotating vane thrust reverser represented a concept in which reversing is accomplished upstream of the nozzle throat, and a three door reverser concept provided reversing downstream of the nozzle throat. Nozzle pressure ratio was varied from 2.0 to approximately 6.0. The results of this investigation indicate that both the rotating vane and three door reverser concepts were effective static thrust spoilers with the landing approach nozzle geometry and were capable of providing at least a 50 percent reversal of static thrust when fully deployed with the ground roll nozzle geometry.

Research on axial thrust of the waterjet pump based on CFD under cavitation conditions

NASA Astrophysics Data System (ADS)

Shen, Z. H.; Pan, Z. Y.

2015-01-01

Based on RANS equations, performance of a contra-rotating axial-flow waterjet pump without hydrodynamic cavitation state had been obtained combined with shear stress transport turbulence model. Its cavitation hydrodynamic performance was calculated and analysed with mixture homogeneous flow cavitation model based on Rayleigh-Plesset equations. The results shows that the cavitation causes axial thrust of waterjet pump to drop. Furthermore, axial thrust and head cavitation characteristic curve is similar. However, the drop point of the axial thrust is postponed by 5.1% comparing with one of head, and the critical point of the axial thrust is postponed by 2.6%.

Space Storable Propellant Performance Gas/Liquid Like-Doublet Injector Characterization

NASA Technical Reports Server (NTRS)

Falk, A. Y.

1972-01-01

A 30-month applied research program was conducted, encompassing an analytical, design, and experimental effort to relate injector design parameters to simultaneous attainment of high performance and component (injector/thrust chamber) compatibility for gas/liquid space-storable propellants. The gas/liquid propellant combination selected for study was FLOX (82.6% F2)/ambient temperature gaseous methane. The injector pattern characterized was the like-(self)-impinging doublet. Program effort was apportioned into four basic technical tasks: injector and thrust chamber design, injector and thrust chamber fabrication, performance evaluation testing, and data evaluation and reporting. Analytical parametric combustion analyses and cold flow distribution and atomization experiments were conducted with injector segment models to support design of injector/thrust chamber combinations for hot fire evaluation. Hot fire tests were conducted to: (1) optimize performance of the injector core elements, and (2) provide design criteria for the outer zone elements so that injector/thrust chamber compatibility could be achieved with only minimal performance losses.

A system study for the application of microcomputers to research flight test techniques

NASA Technical Reports Server (NTRS)

Smyth, R. K.

1983-01-01

The onboard simulator is a three degree of freedom aircraft behavior simulator which provides parameters used by the interception procedure. These parameters can be used for verifying closed loop performance before flight. The air to air intercept mode is a software package integrated in the simulation process that generates a target motion and performs a tracking procedure that predicts the most likely next target position, for a defined time step. This procedure also updates relative position parameters and gives adequate fire commands. A microcomputer based on an aircraft spin warning system periodically samples the assymetric thrust and yaw rate of an airplane and then issues voice synthesized warnings and /or suggests to the ilot how to respond to the situation.

Static internal performance evaluation of several thrust reversing concepts for 2D-CD nozzles

NASA Technical Reports Server (NTRS)

Rowe, R. K.; Duss, D. J.; Leavitt, L. D.

1984-01-01

Recent performance testing of the two-dimensional convergent-divergent (2D-CD) nozzle has established the concept as a viable alternative to the axisymmetric nozzle for advanced technology aircraft. This type of exhaust system also offers potential integration and performance advantages in the areas of thrust reversing and vectoring over axi-symmetric nozzles. These advantages include the practical integration of thrust reversers which operate not only to reduce landing roll but also operate in-flight for enhanced maneuvering and thrust spoiling. To date there is a very limited data base available from which criteria can be developed for the design and evaluation of this type of thrust reverser system. For this reason, a static scale model test was conducted in which five different thrust reverser designs were evaluated. Each of the five models had varying performance/integration requirements which dictated the five different designs. Some of the parameters investigated in this test included; variable angle external cascade vanes, fixed angle internal cascade vanes, variable position inner doors, external slider doors and internal slider valves. In addition, normal force and yawing moment generation was investigated using the thrust reverser system. Selected results from this test will be presented and discussed in this paper.

Extended performance solar electric propulsion thrust system study. Volume 3: Tradeoff studies of alternate thrust system configurations

NASA Technical Reports Server (NTRS)

Hawthorne, E. I.

1977-01-01

Several thrust system design concepts were evaluated and compared using the specifications of the most advanced 30 cm engineering model thruster as the technology base. Emphasis was placed on relatively high power missions. The extensions in thruster performance required for the Halley's comet mission were defined and alternative thrust system concepts were designed in sufficient detail for comparing mass, efficiency, reliability, structure, and thermal characteristics. Confirmation testing and analysis of thruster and power-processing components were performed. A baseline design was selected from the alternatives considered, and the design analysis and documentation were refined. A program development plan was formulated that outlines the work structure considered necessary for developing, qualifying, and fabricating the flight hardware for the baseline thrust system within the time frame of a project to rendezvous with Halley's comet. An assessment was made of the costs and risks associated with a baseline thrust system as provided to the mission project under this plan. Critical procurements and interfaces were identified and defined.

Development of Composite Technologies for the European Next Generation Launcher

NASA Astrophysics Data System (ADS)

Fatemi, Javad; van der Bas, Finn

2014-06-01

In the frame of the European Space Agency's Future Launchers Preparatory Programme (FLPP), in conjunction with national Research and Technology programs, Dutch Space has undertaken the development of composite technologies for application in the Europe's next generation launcher, Ariane 6. The efforts have focused on development of a Carbon Fibre Reinforced Plastic (CFRP) Engine Thrust Frame (ETF) for the upper-stage of Ariane6 launcher. These new technologies are expected to improve performance and to lower cost of development and exploitation of the launcher. Although the first targeted application is the thrust frame, the developed technologies are set to be generic in the sense that they can be applied to other structures of the launcher, e.g. inter-stage structures.This paper addresses the design, analysis, manufacturing and testing activities related to the composite technology developments.

Aerospace Laser Ignition/Ablation Variable High Precision Thruster

NASA Technical Reports Server (NTRS)

Campbell, Jonathan W. (Inventor); Edwards, David L. (Inventor); Campbell, Jason J. (Inventor)

2015-01-01

A laser ignition/ablation propulsion system that captures the advantages of both liquid and solid propulsion. A reel system is used to move a propellant tape containing a plurality of propellant material targets through an ignition chamber. When a propellant target is in the ignition chamber, a laser beam from a laser positioned above the ignition chamber strikes the propellant target, igniting the propellant material and resulting in a thrust impulse. The propellant tape is advanced, carrying another propellant target into the ignition chamber. The propellant tape and ignition chamber are designed to ensure that each ignition event is isolated from the remaining propellant targets. Thrust and specific impulse may by precisely controlled by varying the synchronized propellant tape/laser speed. The laser ignition/ablation propulsion system may be scaled for use in small and large applications.

Static performance of vectoring/reversing non-axisymmetric nozzles

NASA Technical Reports Server (NTRS)

Willard, C. M.; Capone, F. J.; Konarski, M.; Stevens, H. L.

1977-01-01

An experimental program sponsored by the Air Force Flight Dynamics Laboratory is currently in progress to determine the internal and installed performance characteristics of five different thrust vectoring/reversing non-axisymmetric nozzle concepts for tactical fighter aircraft applications. Internal performance characteristics for the five non-axisymmetric nozzles and an advanced technology axisymmetric baseline nozzle were determined in static tests conducted in January 1977 at the NASA-Langley Research Center. The non-axisymmetric nozzle models were tested at thrust deflection angles of up to 30 degrees from horizontal at throat areas associated with both dry and afterburning power. In addition, dry power reverse thrust geometries were tested for three of the concepts. The best designs demonstrated internal performance levels essentially equivalent to the baseline axisymmetric nozzle at unvectored conditions. The best designs also gave minimum performance losses due to vectoring, and reverse thrust levels up to 50% of maximum dry power forward thrust. The installed performance characteristics will be established based on wind tunnel testing to be conducted at Arnold Engineering Development Center in the fall of 1977.

Investigation of Thrust Augmentation and Acoustic Performance by Ejectors on PDE

NASA Astrophysics Data System (ADS)

Xu, Gui-yang; Weng, Chun-sheng; Li, Ning; Huang, Xiao-long

2016-04-01

Thrust augmentation and acoustic performance of a Pulse Detonation Engine (PDE) with ejector system is experimentally investigated. For these tests the LEjector/DEjector is varied from 1.18 to 4 and the axial placement of the ejector relative to the PDE exhaust is varied from an x/DPDE of -3 to 3. Results from the tests show that the optimum LEjector/DEjector based on thrust augmentation and Overall Sound Pressure Level (OASPL) is found to be 2.61. The divergent ejector performed the best based on thrust augmentation, while the reduction effect for OASPL and Peak Sound Pressure Level (PSPL) at 60° is most prominent for the convergent ejector. The optimum axial position based on thrust augmentation is determined to be x/DPDE = 2, while, x/DPDE = 0 based on OASPL and PSPL.

CAVITATION SOUNDS DURING CERVICOTHORACIC SPINAL MANIPULATION.

PubMed

Dunning, James; Mourad, Firas; Zingoni, Andrea; Iorio, Raffaele; Perreault, Thomas; Zacharko, Noah; de Las Peñas, César Fernández; Butts, Raymond; Cleland, Joshua A

2017-08-01

No study has previously investigated the side, duration or number of audible cavitation sounds during high-velocity low-amplitude (HVLA) thrust manipulation to the cervicothoracic spine. The primary purpose was to determine which side of the spine cavitates during cervicothoracic junction (CTJ) HVLA thrust manipulation. Secondary aims were to calculate the average number of cavitations, the duration of cervicothoracic thrust manipulation, and the duration of a single cavitation. Quasi-experimental study. Thirty-two patients with upper trapezius myalgia received two cervicothoracic HVLA thrust manipulations targeting the right and left T1-2 articulation, respectively. Two high sampling rate accelerometers were secured bilaterally 25 mm lateral to midline of the T1-2 interspace. For each manipulation, two audio signals were extracted using Short-Time Fourier Transformation (STFT) and singularly processed via spectrogram calculation in order to evaluate the frequency content and number of instantaneous energy bursts of both signals over time for each side of the CTJ. Unilateral cavitation sounds were detected in 53 (91.4%) of 58 cervicothoracic HVLA thrust manipulations and bilateral cavitation sounds were detected in just five (8.6%) of the 58 thrust manipulations; that is, cavitation was significantly (p<0.001) more likely to occur unilaterally than bilaterally. In addition, cavitation was significantly (p<0.0001) more likely to occur on the side contralateral to the clinician's short-lever applicator. The mean number of audible cavitations per manipulation was 4.35 (95% CI 2.88, 5.76). The mean duration of a single manipulation was 60.77 ms (95% CI 28.25, 97.42) and the mean duration of a single audible cavitation was 4.13 ms (95% CI 0.82, 7.46). In addition to single-peak and multi-peak energy bursts, spectrogram analysis also demonstrated high frequency sounds, low frequency sounds, and sounds of multiple frequencies for all 58 manipulations. Cavitation was significantly more likely to occur unilaterally, and on the side contralateral to the short-lever applicator contact, during cervicothoracic HVLA thrust manipulation. Clinicians should expect multiple cavitation sounds when performing HVLA thrust manipulation to the CTJ. Due to the presence of multi-peak energy bursts and sounds of multiple frequencies, the cavitation hypothesis (i.e. intra-articular gas bubble collapse) alone appears unable to explain all of the audible sounds during HVLA thrust manipulation, and the possibility remains that several phenomena may be occurring simultaneously. 2b.

Laser Plasma Microthruster Performance Evaluation

NASA Astrophysics Data System (ADS)

Luke, James R.; Phipps, Claude R.

2003-05-01

The micro laser plasma thruster (μLPT) is a sub-kilogram thruster that is capable of meeting the Air Force requirements for the Attitude Control System on a 100-kg class small satellite. The μLPT uses one or more 4W diode lasers to ablate a solid fuel, producing a jet of hot gas or plasma which creates thrust with a high thrust/power ratio. A pre-prototype continuous thrust experiment has been constructed and tested. The continuous thrust experiment uses a 505 mm long continuous loop fuel tape, which consists of a black laser-absorbing fuel material on a transparent plastic substrate. When the laser is operated continuously, the exhaust plume and thrust vector are steered in the direction of the tape motion. Thrust steering can be avoided by pulsing the laser. A torsion pendulum thrust stand has been constructed and calibrated. Many fuel materials and substrates have been tested. Best performance from a non-energetic fuel material was obtained with black polyvinyl chloride (PVC), which produced an average of 70 μN thrust and coupling coefficient (Cm) of 190 μN/W. A proprietary energetic material was also tested, in which the laser initiates a non-propagating detonation. This material produced 500 μN of thrust.

Extended performance solar electric propulsion thrust system study. Volume 1: Executive summary

NASA Technical Reports Server (NTRS)

Poeschel, R. L.; Hawthorne, E. I.

1977-01-01

Several thrust system design concepts were evaluated and compared using the specifications of the most advanced 30 cm engineering model thruster as the technology base. The extensions in thruster performance required for the Halley's comet mission were defined and alternative thrust system concepts were designed. Confirmation testing and analysis of thruster and power-processing components were performed, and the feasibility of satisfying extended performance requirements was verified. A baseline design was selected from the alternatives considered, and the design analysis and documentation were refined. A program development plan was formulated that outlines the work structure considered necessary for developing, qualifying, and fabricating the flight hardware for the baseline thrust system within the time frame of a project to rendezvous with Halley's comet. An assessment was made of the costs and risks associated with a baseline thrust system as provided to the mission project under this plan. Critical procurements and interfaces were identified and defined. Results are presented.

Detailed Design of a Pulsed Plasma Thrust Stand

NASA Astrophysics Data System (ADS)

Verbin, Andrew J.

This thesis gives a detailed design process for a pulsed type thruster. The thrust stand designed in this paper is for a Pulsed Plasma Thruster built by Sun Devil Satellite Laboratory, a student organization at Arizona State University. The thrust stand uses a torsional beam rotating to record displacement. This information, along with impulse-momentum theorem is applied to find the impulse bit of the thruster, which varies largely from other designs which focus on using the natural dynamics their fixtures. The target impulse to record on this fixture was estimated to be 275 muN-s of impulse. Through calibration and experimentation, the fixture is capable of recording an impulse of 332 muN-s +/- 14.81 muN-s, close to the target impulse. The error due to noise was characterized and evaluated to be under 5% which is deemed to be acceptable.

Evaluation of a ducted-fan power plant designed for high output and good cruise fuel economy

NASA Technical Reports Server (NTRS)

Behun, M; Rom, F E; Hensley, R V

1950-01-01

Theoretical analysis of performance of a ducted-fan power plant designed both for high-output, high-altitude operation at low supersonic Mach numbers and for good fuel economy at lower fight speeds is presented. Performance of ducted fan is compared with performance (with and without tail-pipe burner) of two hypothetical turbojet engines. At maximum power, the ducted fan has propulsive thrust per unit of frontal area between thrusts obtained by turbojet engines with and without tail-pipe burners. At cruise, the ducted fan obtains lowest thrust specific fuel consumption. For equal maximum thrusts, the ducted fan obtains cruising flight duration and range appreciably greater than turbojet engines.

Investigation of two-dimensional wedge exhaust nozzles for advanced aircraft

NASA Technical Reports Server (NTRS)

Maiden, D. L.; Petit, J. E.

1975-01-01

Two-dimensional wedge nozzle performance characteristics were investigated in a series of wind-tunnel tests. An isolated single-engine/nozzle model was used to study the effects of internal expansion area ratio, aftbody cowl boattail angle, and wedge length. An integrated twin-engine/nozzle model, tested with and without empenage surfaces, included cruise, acceleration, thrust vectoring and thrust reversing nozzle operating modes. Results indicate that the thrust-minus-aftbody drag performance of the twin two-dimensional nozzle integration is significantly higher, for speeds greater than Mach 0.8, than the performance achieved with twin axisymmetric nozzle installations. Significant jet-induced lift was obtained on an aft-mounted lifting surface using a cambered wedge center body to vector thrust. The thrust reversing capabilities of reverser panels installed on the two-dimensional wedge center body were very effective for static or in-flight operation.

Aerodynamic Performance of a 0.27-Scale Model of an AH-64 Helicopter with Baseline and Alternate Rotor Blade Sets

NASA Technical Reports Server (NTRS)

Kelley, Henry L.

1990-01-01

Performance of a 27 percent scale model rotor designed for the AH-64 helicopter (alternate rotor) was measured in hover and forward flight and compared against and AH-64 baseline rotor model. Thrust, rotor tip Mach number, advance ratio, and ground proximity were varied. In hover, at a nominal thrust coefficient of 0.0064, the power savings was about 6.4 percent for the alternate rotor compared to the baseline. The corresponding thrust increase at this condition was approx. 4.5 percent which represents an equivalent full scale increase in lift capability of about 660 lbs. Comparable results were noted in forward flight except for the high thrust, high speed cases investigated where the baseline rotor was slightly superior. Reduced performance at the higher thrusts and speeds was likely due to Reynolds number effects and blade elasticity differences.

Interplanetary Trajectory Design for the Asteroid Robotic Redirect Mission Alternate Approach Trade Study

NASA Technical Reports Server (NTRS)

Merrill, Raymond Gabriel; Qu, Min; Vavrina, Matthew A.; Englander, Jacob A.; Jones, Christopher A.

2014-01-01

This paper presents mission performance analysis methods and results for the Asteroid Robotic Redirect Mission (ARRM) option to capture a free standing boulder on the surface of a 100 m or larger NEA. It details the optimization and design of heliocentric low-thrust trajectories to asteroid targets for the ARRM solar electric propulsion spacecraft. Extensive searches were conducted to determine asteroid targets with large pick-up mass potential and potential observation opportunities. Interplanetary trajectory approximations were developed in method based tools for Itokawa, Bennu, 1999 JU3, and 2008 EV5 and were validated by end-to-end integrated trajectories.

Study of Required Thrust Profile Determination of a Three Stages Small Launch Vehicle

NASA Astrophysics Data System (ADS)

Fariz, A.; Sasongko, R. A.; Poetro, R. E.

2018-04-01

The effect of solid rocket motor specifications, i.e. specific impulse and mass flow rate, and coast time on the thrust profile of three stages small launch vehicle is studied. Solid rocket motor specifications are collected from various small launch vehicle that had ever been in operation phase, and also from previous study. Comparison of orbital parameters shows that the radius of apocenter targeted can be approached using one combination of solid rocket motor specifications and appropriate coast time. However, the launch vehicle designed is failed to achieve the targeted orbit nor injecting the satellite to any orbit.

A Performance and Plume Comparison of Xenon and Krypton Propellant on the SPT-100

DTIC Science & Technology

2012-07-02

HET (1.35 kW), performance measurements were made using an inverted pendulum thrust stand. The plume was also characterized by a Faraday probe and RPA...performance reduction for the case of the flight model SPT-100 HET (1.35 kW), per- formance measurements were made using an inverted pendulum thrust stand...where high thrust at reduced specific impulse is advantageous, such as orbit raising missions. Bismuth’s main drawback is that the metal must be

The Dawn of Vesta Science

NASA Technical Reports Server (NTRS)

Garner, Charles E.; Rayman, Marc D.; Brophy, John R.; Mikes, Steven C.

2011-01-01

The Dawn mission is part of NASA's Discovery Program and has as its goal the scientific exploration of the two most massive main-belt asteroids, Vesta and Ceres. The Dawn spacecraft was launched from the Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H-9.5 (Delta-II Heavy) rocket that placed the 1218-kg spacecraft onto an Earth-escape trajectory. On-board the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory which will provide a total ?V of 11.3 km/s for the heliocentric transfer to Vesta, orbit capture at Vesta, transfer between Vesta science orbits, departure and escape from Vesta, heliocentric transfer to Ceres, orbit capture at Ceres, and transfer between Ceres science orbits. Fullpower thrusting from December 2007 through October 2008 was used to successfully target a Mars gravity assist flyby in February 2009 that provided an additional ?V of 2.6 km/s. Deterministic thrusting for the heliocentric transfer to Vesta resumed in June 2009 and concluded with orbit capture at Vesta on July 16, 2011. An additional 210 hours of IPS thrusting was used to enter the first Vesta science orbit, called Survey orbit, on August 3, 2011 at an altitude of approximately 2,735 km. To date the IPS has been operated for 23,621 hours, consumed approximately 252 kg of xenon, and provided a delta-V of approximately 6.7 km/s. The IPS performance characteristics are very close to the expected performance based on analysis and testing performed pre-launch. The only significant issue in over the almost four years of IPS operations in flight was the temporary failure of a valve driver board in the Digital Control and Interface Unit-1 (DCIU-1), resulting in a loss of thrust of approximately 29 hours. Thrusting operations resumed after switching to DCIU-2, and power cycling conducted after orbit capture restored DCIU-1 to full functionality. After about three weeks of Survey orbit operations the IPS will be used to transfer the spacecraft to the other planned science orbit altitudes. After approximately one year of science operations the IPS will be used for escape from Vesta and then for thrusting to Ceres with a planned arrival date at Ceres in February 2015. This paper provides an overview of Dawn's mission objectives and the results of Dawn IPS mission operations through the start of science operations at Vesta.

Use of magnetic sails for advanced exploration missions

NASA Technical Reports Server (NTRS)

Andrews, Dana G.; Zubrin, Robert M.

1990-01-01

The magnetic sail, or magsail, is a field effect device which interacts with the ambient solar wind or interstellar medium over a considerable volume of space to generate drag and lift forces. Two theories describing the method of thrust generation are analyzed and data results are presented. The techniques for maintaining superconductor temperatures in interplanetary space are analyzed and low risk options presented. Comparisons are presented showing mission performance differences between currently proposed spacecraft using chemical and electric propulsion systems, and a Magsail propelled spacecraft capable of generating an average thrust of 250 Newtons at a radius of one A.U. The magsail also provides unique capabilities for interstellar missions, in that at relativistic speeds the magnetic field would ionize and deflect the interstellar medium producing a large drag force. This would make it an ideal brake for decelerating a spacecraft from relativistic speeds and then maneuvering within the target star system.

Performance of a 100 kW class applied field MPD thruster

NASA Technical Reports Server (NTRS)

Mantenieks, Maris A.; Sovey, James S.; Myers, Roger M.; Haag, Thomas W.; Raitano, Paul; Parkes, James E.

1989-01-01

Performance of a 100 kW, applied field magnetoplasmadynamic (MPD) thruster was evaluated and sensitivities of discharge characteristics to arc current, mass flow rate, and applied magnetic field were investigated. Thermal efficiencies as high as 60 percent, thrust efficiencies up to 21 percent, and specific impulses of up to 1150 s were attained with argon propellant. Thrust levels up to 2.5 N were directly measured with an inverted pendulum thrust stand at discharge input powers up to 57 kW. It was observed that thrust increased monotonically with the product of arc current and magnet current.

Statistical error model for a solar electric propulsion thrust subsystem

NASA Technical Reports Server (NTRS)

Bantell, M. H.

1973-01-01

The solar electric propulsion thrust subsystem statistical error model was developed as a tool for investigating the effects of thrust subsystem parameter uncertainties on navigation accuracy. The model is currently being used to evaluate the impact of electric engine parameter uncertainties on navigation system performance for a baseline mission to Encke's Comet in the 1980s. The data given represent the next generation in statistical error modeling for low-thrust applications. Principal improvements include the representation of thrust uncertainties and random process modeling in terms of random parametric variations in the thrust vector process for a multi-engine configuration.

Parametric Model of an Aerospike Rocket Engine

NASA Technical Reports Server (NTRS)

Korte, J. J.

2000-01-01

A suite of computer codes was assembled to simulate the performance of an aerospike engine and to generate the engine input for the Program to Optimize Simulated Trajectories. First an engine simulator module was developed that predicts the aerospike engine performance for a given mixture ratio, power level, thrust vectoring level, and altitude. This module was then used to rapidly generate the aerospike engine performance tables for axial thrust, normal thrust, pitching moment, and specific thrust. Parametric engine geometry was defined for use with the engine simulator module. The parametric model was also integrated into the iSIGHTI multidisciplinary framework so that alternate designs could be determined. The computer codes were used to support in-house conceptual studies of reusable launch vehicle designs.

Parametric Model of an Aerospike Rocket Engine

NASA Technical Reports Server (NTRS)

Korte, J. J.

2000-01-01

A suite of computer codes was assembled to simulate the performance of an aerospike engine and to generate the engine input for the Program to Optimize Simulated Trajectories. First an engine simulator module was developed that predicts the aerospike engine performance for a given mixture ratio, power level, thrust vectoring level, and altitude. This module was then used to rapidly generate the aerospike engine performance tables for axial thrust, normal thrust, pitching moment, and specific thrust. Parametric engine geometry was defined for use with the engine simulator module. The parametric model was also integrated into the iSIGHT multidisciplinary framework so that alternate designs could be determined. The computer codes were used to support in-house conceptual studies of reusable launch vehicle designs.

Supersonic wings with significant leading-edge thrust at cruise

NASA Technical Reports Server (NTRS)

Robins, A. W.; Carlson, H. W.; Mack, R. J.

1980-01-01

Experimental/theoretical correlations are presented which show that significant levels of leading-edge thrust are possible at supersonic speeds for certain planforms having the geometry to support the theoretical thrust-distribution potential. The new analytical process employed provides not only the level of leading-edge thrust attainable but also the spanwise distribution of both it and that component of full theoretical thrust which acts as vortex lift. Significantly improved aerodynamic performance in the moderate supersonic speed regime is indicated.

Some tests on small-scale rectangular throat ejector. [thrust augmentation for V/STOL aircraft

NASA Technical Reports Server (NTRS)

Dean, W. N., Jr.; Franke, M. E.

1979-01-01

A small scale rectangular throat ejector with plane slot nozzles and a fixed throat area was tested to determine the effects of diffuser sidewall length, diffuser area ratio, and sidewall nozzle position on thrust and mass augmentation. The thrust augmentation ratio varied from approximately 0.9 to 1.1. Although the ejector did not have good thrust augmentation performance, the effects of the parameters studied are believed to indicate probable trends in thrust augmenting ejectors.

Over-the-wing model thrust reverser noise tests

NASA Technical Reports Server (NTRS)

Goodykoontz, J.; Gutierrez, O.

1977-01-01

Static acoustic tests were conducted on a 1/12 scale model over-the-wing target type thrust reverser. The model configuration simulates a design that is applicable to the over-the-wing short-haul advanced technology engine. Aerodynamic screening tests of a variety of reverser designs identified configurations that satisfied a reverse thrust requirement of 35 percent of forward thrust at a nozzle pressure ratio of 1.29. The variations in the reverser configuration included, blocker door angle, blocker door lip angle and shape, and side skirt shape. Acoustic data are presented and compared for the various configurations. The model data scaled to a single full size engine show that peak free field perceived noise (PN) levels at a 152.4 meter sideline distance range from 98 to 104 PNdb.

Multiphysics Computational Analysis of a Solid-Core Nuclear Thermal Engine Thrust Chamber

NASA Technical Reports Server (NTRS)

Wang, Ten-See; Canabal, Francisco; Cheng, Gary; Chen, Yen-Sen

2007-01-01

The objective of this effort is to develop an efficient and accurate computational heat transfer methodology to predict thermal, fluid, and hydrogen environments for a hypothetical solid-core, nuclear thermal engine - the Small Engine. In addition, the effects of power profile and hydrogen conversion on heat transfer efficiency and thrust performance were also investigated. The computational methodology is based on an unstructured-grid, pressure-based, all speeds, chemically reacting, computational fluid dynamics platform, while formulations of conjugate heat transfer were implemented to describe the heat transfer from solid to hydrogen inside the solid-core reactor. The computational domain covers the entire thrust chamber so that the afore-mentioned heat transfer effects impact the thrust performance directly. The result shows that the computed core-exit gas temperature, specific impulse, and core pressure drop agree well with those of design data for the Small Engine. Finite-rate chemistry is very important in predicting the proper energy balance as naturally occurring hydrogen decomposition is endothermic. Locally strong hydrogen conversion associated with centralized power profile gives poor heat transfer efficiency and lower thrust performance. On the other hand, uniform hydrogen conversion associated with a more uniform radial power profile achieves higher heat transfer efficiency, and higher thrust performance.

Design of Launch Abort System Thrust Profile and Concept of Operations

NASA Technical Reports Server (NTRS)

Litton, Daniel; O'Keefe, Stephen A.; Winski, Richard G.; Davidson, John B.

2008-01-01

This paper describes how the Abort Motor thrust profile has been tailored and how optimizing the Concept of Operations on the Launch Abort System (LAS) of the Orion Crew Exploration Vehicle (CEV) aides in getting the crew safely away from a failed Crew Launch Vehicle (CLV). Unlike the passive nature of the Apollo system, the Orion Launch Abort Vehicle will be actively controlled, giving the program a more robust abort system with a higher probability of crew survival for an abort at all points throughout the CLV trajectory. By optimizing the concept of operations and thrust profile the Orion program will be able to take full advantage of the active Orion LAS. Discussion will involve an overview of the development of the abort motor thrust profile and the current abort concept of operations as well as their effects on the performance of LAS aborts. Pad Abort (for performance) and Maximum Drag (for separation from the Launch Vehicle) are the two points that dictate the required thrust and shape of the thrust profile. The results in this paper show that 95% success of all performance requirements is not currently met for Pad Abort. Future improvements to the current parachute sequence and other potential changes will mitigate the current problems, and meet abort performance requirements.

Exoskeletal Engine Concept: Feasibility Studies for Medium and Small Thrust Engines

NASA Technical Reports Server (NTRS)

Halliwell, Ian

2001-01-01

The exoskeletal engine concept is one in which the shafts and disks are eliminated and are replaced by rotating casings that support the blades in spanwise compression. Omission of the shafts and disks leads to an open channel at the engine centerline. This has immense potential for reduced jet noise and for the accomodation of an alternative form of thruster for use in a combined cycle. The use of ceramic composite materials has the potential for significantly reduced weight as well as higher working temperatures without cooling air. The exoskeletal configuration is also a natural stepping-stone to complete counter-rotating turbomachinery. Ultimately this will lead to reductions in weight, length, parts count and improved efficiency. The feasibility studies are in three parts. Part I-Systems and Component Requirements addressed the mechanical aspects of components from a functionality perspective. This effort laid the groundwork for preliminary design studies. Although important, it is not felt to be particularly original, and has therefore not been included in the current overview. Part 2-Preliminary Design Studies turned to some of the cycle and performance issues inherent in an exoskeletal configuration and some initial attempts at preliminary design of turbomachinery were described. Twin-spoon and single-spool 25.800-lbf-thrust turbofans were used as reference vehicles in a mid-size commercial subsonic category in addition to a single-spool 5,000-lbf-thrust turbofan that represented a general aviation application. The exoskeletal engine, with its open centerline, has tremendous potential for noise suppression and some preliminary analysis was done which began to quantify the benefits. Part 3-Additional Preliminary Design Studies revisited the design of single-spool 25,800-lbf-thrust turbofan configurations, but in addition to the original FPR = 1.6 and BPR = 5.1 reference engine, two additional configurations used FPR = 2.4 and BPR = 3.0 and FPR = 3.2 and BPR = 2.0 were investigated. The single-spool 5,000-lbf-thrust turbofan was refined and the small engine study was extended to include a 2,000-lbf-thrust turbojet. More attention was paid to optimizing the turbomachinery. Turbine cooling flows were eliminated, in keeping with the use of uncooled CMC material in exoskeletal engines. The turbine performance parameters moved much closer to the nominal target values, demonstrating the great benefits to the cycle of uncooled turbines.

Exoskeletal Engine Concept: Feasibility Studies for Medium and Small Thrust Engines

NASA Technical Reports Server (NTRS)

Halliwell, Ian

2001-01-01

The exoskeletal engine concept is one in which the shafts and disks are eliminated and are replaced by rotating casings that support the blades in spanwise compression. Omission of the shafts and disks leads to an open channel at the engine centerline. This has immense potential for reduced jet noise and for the accommodation of an alternative form of thruster for use in a combined cycle. The use of ceramic composite materials has the potential for significantly reduced weight as well as higher working temperatures without cooling air. The exoskeletal configuration is also a natural stepping-stone to complete counter-rotating turbomachinery. Ultimately this will lead to reductions in weight, length, parts count and improved efficiency. The feasibility studies are in three parts. Part 1: Systems and Component Requirements addressed the mechanical aspects of components from a functionality perspective. This effort laid the groundwork for preliminary design studies. Although important, it is not felt to be particularly original, and has therefore not been included in the current overview. Part 2: Preliminary Design Studies turned to some of the cycle and performance issues inherent in an exoskeletal configuration and some initial attempts at preliminary design of turbomachinery were described. Twin-spoon and single-spool 25,800-lbf-thrust turbofans were used as reference vehicles in a mid-size commercial subsonic category in addition to a single-spool 5,000-lbf-thrust turbofan that represented a general aviation application. The exoskeletal engine, with its open centerline, has tremendous potential for noise suppression and some preliminary analysis was done which began to quantify the benefits. Part 3: Additional Preliminary Design Studies revisited the design of single-spool 25,800-lbf-thrust turbofan configurations, but in addition to the original FPR = 1.6 and BPR = 5.1 reference engine. two additional configurations used FPR = 2.4 and BPR = 3.0 and FPR = 3.2 and BPR = 2.0 were investigated. The single-spool 5.000-lbf-thrust turbofan was refined and the small engine study was extended to include a 2,000-lbf-thrust turbojet. More attention was paid to optimizing the turbomachinery. Turbine cooling flows were eliminated, in keeping with the use of uncooled CMC materials in exoskeletal engines. The turbine performance parameters moved much closer to the nominal target values, demonstrating the great benefits to the cycle of uncooled turbines.

Magnetized Target Fusion

NASA Technical Reports Server (NTRS)

Griffin, Steven T.

2002-01-01

Magnetized target fusion (MTF) is under consideration as a means of building a low mass, high specific impulse, and high thrust propulsion system for interplanetary travel. This unique combination is the result of the generation of a high temperature plasma by the nuclear fusion process. This plasma can then be deflected by magnetic fields to provide thrust. Fusion is initiated by a small traction of the energy generated in the magnetic coils due to the plasma's compression of the magnetic field. The power gain from a fusion reaction is such that inefficiencies due to thermal neutrons and coil losses can be overcome. Since the fusion reaction products are directly used for propulsion and the power to initiate the reaction is directly obtained from the thrust generation, no massive power supply for energy conversion is required. The result should be a low engine mass, high specific impulse and high thrust system. The key is to successfully initiate fusion as a proof-of-principle for this application. Currently MSFC is implementing MTF proof-of-principle experiments. This involves many technical details and ancillary investigations. Of these, selected pertinent issues include the properties, orientation and timing of the plasma guns and the convergence and interface development of the "pusher" plasma. Computer simulations of the target plasma's behavior under compression and the convergence and mixing of the gun plasma are under investigation. This work is to focus on the gun characterization and development as it relates to plasma initiation and repeatability.

Subsonic Performance of Ejector Systems

NASA Astrophysics Data System (ADS)

Weil, Samuel

Combined cycle engines combining scramjets with turbo jets or rockets can provide efficient hypersonic flight. Ejectors have the potential to increase the thrust and efficiency of combined cycle engines near static conditions. A computer code was developed to support the design of a small-scale, turbine-based combined cycle demonstrator with an ejector, built around a commercially available turbojet engine. This code was used to analyze the performance of an ejector system built around a micro-turbojet. With the use of a simple ejector, net thrust increases as large as 20% over the base engine were predicted. Additionally the specific fuel consumption was lowered by 10%. Increasing the secondary to primary area ratio of the ejector lead to significant improvements in static thrust, specific fuel consumption (SFC), and propulsive efficiency. Further ejector performance improvements can be achieved by using a diffuser. Ejector performance drops off rapidly with increasing Mach number. The ejector has lower thrust and higher SFC than the turbojet core at Mach numbers above 0.2. When the nozzle chokes a significant drop in ejector performance is seen. When a diffuser is used, higher Mach numbers lead to choking in the mixer and a shock in the nozzle causing a significant decrease in ejector performance. Evaluation of different turbo jets shows that ejector performance depends significantly on the properties of the turbojet. Static thrust and SFC improvements can be achieved with increasing ejector area for all engines, but size of increase and change in performance at higher Mach numbers depend heavily on the turbojet. The use of an ejector in a turbine based combined cycle configuration also increases performance at static conditions with a thrust increase of 5% and SFC decrease of 5% for the tested configuration.

SEP thrust subsystem performance sensitivity analysis

NASA Technical Reports Server (NTRS)

Atkins, K. L.; Sauer, C. G., Jr.; Kerrisk, D. J.

1973-01-01

This is a two-part report on solar electric propulsion (SEP) performance sensitivity analysis. The first part describes the preliminary analysis of the SEP thrust system performance for an Encke rendezvous mission. A detailed description of thrust subsystem hardware tolerances on mission performance is included together with nominal spacecraft parameters based on these tolerances. The second part describes the method of analysis and graphical techniques used in generating the data for Part 1. Included is a description of both the trajectory program used and the additional software developed for this analysis. Part 2 also includes a comprehensive description of the use of the graphical techniques employed in this performance analysis.

Powered Descent Guidance with General Thrust-Pointing Constraints

NASA Technical Reports Server (NTRS)

Carson, John M., III; Acikmese, Behcet; Blackmore, Lars

2013-01-01

The Powered Descent Guidance (PDG) algorithm and software for generating Mars pinpoint or precision landing guidance profiles has been enhanced to incorporate thrust-pointing constraints. Pointing constraints would typically be needed for onboard sensor and navigation systems that have specific field-of-view requirements to generate valid ground proximity and terrain-relative state measurements. The original PDG algorithm was designed to enforce both control and state constraints, including maximum and minimum thrust bounds, avoidance of the ground or descent within a glide slope cone, and maximum speed limits. The thrust-bound and thrust-pointing constraints within PDG are non-convex, which in general requires nonlinear optimization methods to generate solutions. The short duration of Mars powered descent requires guaranteed PDG convergence to a solution within a finite time; however, nonlinear optimization methods have no guarantees of convergence to the global optimal or convergence within finite computation time. A lossless convexification developed for the original PDG algorithm relaxed the non-convex thrust bound constraints. This relaxation was theoretically proven to provide valid and optimal solutions for the original, non-convex problem within a convex framework. As with the thrust bound constraint, a relaxation of the thrust-pointing constraint also provides a lossless convexification that ensures the enhanced relaxed PDG algorithm remains convex and retains validity for the original nonconvex problem. The enhanced PDG algorithm provides guidance profiles for pinpoint and precision landing that minimize fuel usage, minimize landing error to the target, and ensure satisfaction of all position and control constraints, including thrust bounds and now thrust-pointing constraints.

Preliminary supersonic flight test evaluation of performance seeking control

NASA Technical Reports Server (NTRS)

Orme, John S.; Gilyard, Glenn B.

1993-01-01

Digital flight and engine control, powerful onboard computers, and sophisticated controls techniques may improve aircraft performance by maximizing fuel efficiency, maximizing thrust, and extending engine life. An adaptive performance seeking control system for optimizing the quasi-steady state performance of an F-15 aircraft was developed and flight tested. This system has three optimization modes: minimum fuel, maximum thrust, and minimum fan turbine inlet temperature. Tests of the minimum fuel and fan turbine inlet temperature modes were performed at a constant thrust. Supersonic single-engine flight tests of the three modes were conducted using varied after burning power settings. At supersonic conditions, the performance seeking control law optimizes the integrated airframe, inlet, and engine. At subsonic conditions, only the engine is optimized. Supersonic flight tests showed improvements in thrust of 9 percent, increases in fuel savings of 8 percent, and reductions of up to 85 deg R in turbine temperatures for all three modes. The supersonic performance seeking control structure is described and preliminary results of supersonic performance seeking control tests are given. These findings have implications for improving performance of civilian and military aircraft.

Subsonic flight test evaluation of a performance seeking control algorithm on an F-15 airplane

NASA Technical Reports Server (NTRS)

Gilyard, Glenn B.; Orme, John S.

1992-01-01

The subsonic flight test evaluation phase of the NASA F-15 (powered by F 100 engines) performance seeking control program was completed for single-engine operation at part- and military-power settings. The subsonic performance seeking control algorithm optimizes the quasi-steady-state performance of the propulsion system for three modes of operation. The minimum fuel flow mode minimizes fuel consumption. The minimum thrust mode maximizes thrust at military power. Decreases in thrust-specific fuel consumption of 1 to 2 percent were measured in the minimum fuel flow mode; these fuel savings are significant, especially for supersonic cruise aircraft. Decreases of up to approximately 100 degree R in fan turbine inlet temperature were measured in the minimum temperature mode. Temperature reductions of this magnitude would more than double turbine life if inlet temperature was the only life factor. Measured thrust increases of up to approximately 15 percent in the maximum thrust mode cause substantial increases in aircraft acceleration. The system dynamics of the closed-loop algorithm operation were good. The subsonic flight phase has validated the performance seeking control technology, which can significantly benefit the next generation of fighter and transport aircraft.

Experimental study on the thrust modulation performance of powdered magnesium and CO2 bipropellant engine

NASA Astrophysics Data System (ADS)

Li, Chao; Hu, Chunbo; Zhu, Xiaofei; Hu, Jiaming; Li, Yue; Hu, Xu

2018-06-01

Powdered Mg and CO2 bipropellant engine providing a practical demonstration of in situ resource utilization (ISRU) for Mars Sample Return (MSR) mission seems to be feasible by current investigations. However, essential functions of the engine to satisfy the complicated ballistics requirements such as thrust modulation and multiple pulse have not been established yet. The aim of this experimental study is to evaluate the engine's thrust modulation feasibility and to investigate its thrust modulation characteristics. A powdered Mg and CO2 bipropellant engine construction aiming to achieve thrust modulation ability was proposed. A mass flow rate calibration experiment to evaluate the gas-solid mass flow rate regulating performance was conducted before fire tests. Fire test result shows that the engine achieved successful ignition as well as self-sustaining combustion; Thrust modulation of the engine is feasible, detail thrust estimating result of the test shows that maximum thrust is 135.91 N and the minimum is 5.65 N with a 22.11 thrust modulation ratio, moreover, the transportation period is quick and the thrust modulation ratio is adjustable. At the same time, the powder feed system reaches a two-step flow rate regulating with a modulation ratio of 4.5-5. What' more, caused by the uneven engine working conditions, there is an obvious difference in combustion efficiency value, maximum combustion efficiency of the powdered Mg and CO2 bipropellant engine is 80.20%.

Task 12 data dump (phase 2) OME integrated thrust chamber test report

NASA Technical Reports Server (NTRS)

Tobin, R. D.; Pauckert, R. P.

1974-01-01

The characteristics and performance of the orbit maneuvering engine for the space shuttle are discussed. Emphasis is placed on the regeneratively cooled thrust chamber of the engine. Tests were conducted to determine engine operating parameters during the start, shutdown, and restart. Characteristics of the integrated thrust chamber and the performance and thermal conditions for blowdown operation without supplementary boundary layer cooling were investigated. The results of the test program are presented.

A microNewton thrust stand for average thrust measurement of pulsed microthruster.

PubMed

Zhou, Wei-Jing; Hong, Yan-Ji; Chang, Hao

2013-12-01

A torsional thrust stand has been developed for the study of the average thrust for microNewton pulsed thrusters. The main body of the thrust stand mainly consists of a torsional balance, a pair of flexural pivots, a capacitive displacement sensor, a calibration assembly, and an eddy current damper. The behavior of the stand was thoroughly studied. The principle of thrust measurement was analyzed. The average thrust is determined as a function of the average equilibrium angle displacement of the balance and the spring stiffness. The thrust stand has a load capacity up to 10 kg, and it can theoretically measure the force up to 609.6 μN with a resolution of 24.4 nN. The static calibrations were performed based on the calibration assembly composed of the multiturn coil and the permanent magnet. The calibration results demonstrated good repeatability (less than 0.68% FSO) and good linearity (less than 0.88% FSO). The assembly of the multiturn coil and the permanent magnet was also used as an exciter to simulate the microthruster to further research the performance of the thrust stand. Three sets of force pulses at 17, 33.5, and 55 Hz with the same amplitude and pulse width were tested. The repeatability error at each frequency was 7.04%, 1.78%, and 5.08%, respectively.

Large-Scale Wind-Tunnel Tests of Exhaust Ingestion Due to Thrust Reversal on a Four-Engine Jet Transport during Ground Roll

NASA Technical Reports Server (NTRS)

Tolhurst, William H., Jr.; Hickey, David H.; Aoyagi, Kiyoshi

1961-01-01

Wind-tunnel tests have been conducted on a large-scale model of a swept-wing jet transport type airplane to study the factors affecting exhaust gas ingestion into the engine inlets when thrust reversal is used during ground roll. The model was equipped with four small jet engines mounted in nacelles beneath the wing. The tests included studies of both cascade and target type reversers. The data obtained included the free-stream velocity at the occurrence of exhaust gas ingestion in the outboard engine and the increment of drag due to thrust reversal for various modifications of thrust reverser configuration. Motion picture films of smoke flow studies were also obtained to supplement the data. The results show that the free-stream velocity at which ingestion occurred in the outboard engines could be reduced considerably, by simple modifications to the reversers, without reducing the effective drag due to reversed thrust.

Aircraft Engine Thrust Estimator Design Based on GSA-LSSVM

NASA Astrophysics Data System (ADS)

Sheng, Hanlin; Zhang, Tianhong

2017-08-01

In view of the necessity of highly precise and reliable thrust estimator to achieve direct thrust control of aircraft engine, based on support vector regression (SVR), as well as least square support vector machine (LSSVM) and a new optimization algorithm - gravitational search algorithm (GSA), by performing integrated modelling and parameter optimization, a GSA-LSSVM-based thrust estimator design solution is proposed. The results show that compared to particle swarm optimization (PSO) algorithm, GSA can find unknown optimization parameter better and enables the model developed with better prediction and generalization ability. The model can better predict aircraft engine thrust and thus fulfills the need of direct thrust control of aircraft engine.

CAVITATION SOUNDS DURING CERVICOTHORACIC SPINAL MANIPULATION

PubMed Central

Mourad, Firas; Zingoni, Andrea; Iorio, Raffaele; Perreault, Thomas; Zacharko, Noah; de las Peñas, César Fernández; Butts, Raymond; Cleland, Joshua A.

2017-01-01

Background No study has previously investigated the side, duration or number of audible cavitation sounds during high-velocity low-amplitude (HVLA) thrust manipulation to the cervicothoracic spine. Purpose The primary purpose was to determine which side of the spine cavitates during cervicothoracic junction (CTJ) HVLA thrust manipulation. Secondary aims were to calculate the average number of cavitations, the duration of cervicothoracic thrust manipulation, and the duration of a single cavitation. Study Design Quasi-experimental study Methods Thirty-two patients with upper trapezius myalgia received two cervicothoracic HVLA thrust manipulations targeting the right and left T1-2 articulation, respectively. Two high sampling rate accelerometers were secured bilaterally 25 mm lateral to midline of the T1-2 interspace. For each manipulation, two audio signals were extracted using Short-Time Fourier Transformation (STFT) and singularly processed via spectrogram calculation in order to evaluate the frequency content and number of instantaneous energy bursts of both signals over time for each side of the CTJ. Result Unilateral cavitation sounds were detected in 53 (91.4%) of 58 cervicothoracic HVLA thrust manipulations and bilateral cavitation sounds were detected in just five (8.6%) of the 58 thrust manipulations; that is, cavitation was significantly (p<0.001) more likely to occur unilaterally than bilaterally. In addition, cavitation was significantly (p<0.0001) more likely to occur on the side contralateral to the clinician's short-lever applicator. The mean number of audible cavitations per manipulation was 4.35 (95% CI 2.88, 5.76). The mean duration of a single manipulation was 60.77 ms (95% CI 28.25, 97.42) and the mean duration of a single audible cavitation was 4.13 ms (95% CI 0.82, 7.46). In addition to single-peak and multi-peak energy bursts, spectrogram analysis also demonstrated high frequency sounds, low frequency sounds, and sounds of multiple frequencies for all 58 manipulations. Discussion Cavitation was significantly more likely to occur unilaterally, and on the side contralateral to the short-lever applicator contact, during cervicothoracic HVLA thrust manipulation. Clinicians should expect multiple cavitation sounds when performing HVLA thrust manipulation to the CTJ. Due to the presence of multi-peak energy bursts and sounds of multiple frequencies, the cavitation hypothesis (i.e. intra-articular gas bubble collapse) alone appears unable to explain all of the audible sounds during HVLA thrust manipulation, and the possibility remains that several phenomena may be occurring simultaneously. Level of Evidence 2b PMID:28900571

Space shuttle orbit maneuvering engine reusable thrust chamber program

NASA Technical Reports Server (NTRS)

Senneff, J. M.

1975-01-01

Reusable thrust chamber and injector concepts were evaluated for the space shuttle orbit maneuvering engine (OME). Parametric engine calculations were carried out by computer program for N2O4/amine, LOX/amine and LOX/hydrocarbon propellant combinations for engines incorporating regenerative cooled and insulated columbium thrust chambers. The calculation methods are described including the fuel vortex film cooling method of combustion gas temperature control, and performance prediction. A method of acceptance of a regeneratively cooled heat rejection reduction using a silicone oil additive was also demonstrated by heated tube heat transfer testing. Regeneratively cooled thrust chamber operation was also demonstrated where the injector was characterized for the OME application with a channel wall regenerative thrust chamber. Bomb stability testing of the demonstration chambers/injectors demonstrated recovery for the nominal design of acoustic cavities. Cavity geometry changes were also evaluated to assess their damping margin. Performance and combustion stability was demonstrated of the originally developed 10 inch diameter combustion pattern operating in an 8 inch diameter thrust chamber.

A Retro-Fit Control Architecture to Maintain Engine Performance With Usage

NASA Technical Reports Server (NTRS)

Litt, Jonathan S.; Sowers, T. Shane; Garg, Sanjay

2007-01-01

An outer loop retrofit engine control architecture is presented which modifies fan speed command to obtain a desired thrust based on throttle position. This maintains the throttle-to-thrust relationship in the presence of engine degradation, which has the effect of changing the engine s thrust output for a given fan speed. Such an approach can minimize thrust asymmetry in multi-engine aircraft, and reduce pilot workload. The outer loop control is demonstrated under various levels of engine deterioration using a standard deterioration profile as well as an atypical profile. It is evaluated across various transients covering a wide operating range. The modified fan speed command still utilizes the standard engine control logic so all original life and operability limits remain in place. In all cases it is shown that with the outer loop thrust control in place, the deteriorated engine is able to match the thrust performance of a new engine up to the limits the controller will allow.

Evaluating the dynamic response of in-flight thrust calculation techniques during throttle transients

NASA Technical Reports Server (NTRS)

Ray, Ronald J.

1994-01-01

New flight test maneuvers and analysis techniques for evaluating the dynamic response of in-flight thrust models during throttle transients have been developed and validated. The approach is based on the aircraft and engine performance relationship between thrust and drag. Two flight test maneuvers, a throttle step and a throttle frequency sweep, were developed and used in the study. Graphical analysis techniques, including a frequency domain analysis method, were also developed and evaluated. They provide quantitative and qualitative results. Four thrust calculation methods were used to demonstrate and validate the test technique. Flight test applications on two high-performance aircraft confirmed the test methods as valid and accurate. These maneuvers and analysis techniques were easy to implement and use. Flight test results indicate the analysis techniques can identify the combined effects of model error and instrumentation response limitations on the calculated thrust value. The methods developed in this report provide an accurate approach for evaluating, validating, or comparing thrust calculation methods for dynamic flight applications.

Extended performance solar electric propulsion thrust system study. Volume 4: Thruster technology evaluation

NASA Technical Reports Server (NTRS)

Poeschel, R. L.; Hawthorne, E. I.; Weisman, Y. C.; Frisman, M.; Benson, G. C.; Mcgrath, R. J.; Martinelli, R. M.; Linsenbardt, T. L.; Beattie, J. R.

1977-01-01

Several thrust system design concepts were evaluated and compared using the specifications of the most advanced 30 cm engineering model thruster as the technology base. Emphasis was placed on relatively high power missions (60 to 100 kW) such as a Halley's comet rendezvous. The extensions in thruster performance required for the Halley's comet mission were defined and alternative thrust system concepts were designed in sufficient detail for comparing mass, efficiency, reliability, structure, and thermal characteristics. Confirmation testing and analysis of thruster and power processing components were performed, and the feasibility of satisfying extended performance requirements was verified. A baseline design was selected from the alternatives considered, and the design analysis and documentation were refined. The baseline thrust system design features modular construction, conventional power processing, and a concentrator solar array concept and is designed to interface with the Space Shuttle.

Predicted performance of an integrated modular engine system

NASA Technical Reports Server (NTRS)

Binder, Michael; Felder, James L.

1993-01-01

Space vehicle propulsion systems are traditionally comprised of a cluster of discrete engines, each with its own set of turbopumps, valves, and a thrust chamber. The Integrated Modular Engine (IME) concept proposes a vehicle propulsion system comprised of multiple turbopumps, valves, and thrust chambers which are all interconnected. The IME concept has potential advantages in fault-tolerance, weight, and operational efficiency compared with the traditional clustered engine configuration. The purpose of this study is to examine the steady-state performance of an IME system with various components removed to simulate fault conditions. An IME configuration for a hydrogen/oxygen expander cycle propulsion system with four sets of turbopumps and eight thrust chambers has been modeled using the Rocket Engine Transient Simulator (ROCETS) program. The nominal steady-state performance is simulated, as well as turbopump thrust chamber and duct failures. The impact of component failures on system performance is discussed in the context of the system's fault tolerant capabilities.

Space shuttle orbit maneuvering engine reusable thrust chamber program

NASA Technical Reports Server (NTRS)

Senneff, J. M.

1975-01-01

The feasibility of potential reusable thrust chamber concepts is studied. Propellant condidates were examined and analytically combined with potential cooling schemes. A data base of engine data which would assist in a configuration selection was produced. The data base verification was performed by the demonstration of a thrust chamber of a selected coolant scheme design. A full scale insulated columbium thrust chamber was used for propellant coolant configurations. Combustion stability of the injectors and a reduced size thrust chamber were experimentally verified as proof of concept demonstrations of the design and study results.

Supersonic wings with significant leading-edge thrust at cruise

NASA Technical Reports Server (NTRS)

Robins, A. W.; Carlson, H. W.; Mack, R. J.

1980-01-01

Experimental/theoretical correlations are presented which show that significant levels of leading edge thrust are possible at supersonic speeds for certain planforms which match the theoretical thrust distribution potential with the supporting airfoil geometry. The analytical process employed spanwise distribution of both it and/or that component of full theoretical thrust which acts as vortex lift. Significantly improved aerodynamic performance in the moderate supersonic speed regime is indicated.

Lateral dampers for thrust bearings

NASA Technical Reports Server (NTRS)

Hibner, D. H.; Szafir, D. R.

1985-01-01

The development of lateral damping schemes for thrust bearings was examined, ranking their applicability to various engine classes, selecting the best concept for each engine class and performing an in-depth evaluation. Five major engine classes were considered: large transport, military, small general aviation, turboshaft, and non-manrated. Damper concepts developed for evaluation were: curved beam, constrained and unconstrained elastomer, hybrid boost bearing, hydraulic thrust piston, conical squeeze film, and rolling element thrust face.

A Study on Aircraft Engine Control Systems for Integrated Flight and Propulsion Control

NASA Astrophysics Data System (ADS)

Yamane, Hideaki; Matsunaga, Yasushi; Kusakawa, Takeshi; Yasui, Hisako

The Integrated Flight and Propulsion Control (IFPC) for a highly maneuverable aircraft and a fighter-class engine with pitch/yaw thrust vectoring is described. Of the two IFPC functions the aircraft maneuver control utilizes the thrust vectoring based on aerodynamic control surfaces/thrust vectoring control allocation specified by the Integrated Control Unit (ICU) of a FADEC (Full Authority Digital Electronic Control) system. On the other hand in the Performance Seeking Control (PSC) the ICU identifies engine's various characteristic changes, optimizes manipulated variables and finally adjusts engine control parameters in cooperation with the Engine Control Unit (ECU). It is shown by hardware-in-the-loop simulation that the thrust vectoring can enhance aircraft maneuverability/agility and that the PSC can improve engine performance parameters such as SFC (specific fuel consumption), thrust and gas temperature.

Design Enhancements of the Two-Dimensional, Dual Throat Fluidic Thrust Vectoring Nozzle Concept

NASA Technical Reports Server (NTRS)

Flamm, Jeffrey D.; Deere, Karen A.; Mason, Mary L.; Berrier, Bobby L.; Johnson, Stuart K.

2006-01-01

A Dual Throat Nozzle fluidic thrust vectoring technique that achieves higher thrust-vectoring efficiencies than other fluidic techniques, without sacrificing thrust efficiency has been developed at NASA Langley Research Center. The nozzle concept was designed with the aid of the structured-grid, Reynolds-averaged Navier-Stokes computational fluidic dynamics code PAB3D. This new concept combines the thrust efficiency of sonic-plane skewing with increased thrust-vectoring efficiencies obtained by maximizing pressure differentials in a separated cavity located downstream of the nozzle throat. By injecting secondary flow asymmetrically at the upstream minimum area, a new aerodynamic minimum area is formed downstream of the geometric minimum and the sonic line is skewed, thus vectoring the exhaust flow. The nozzle was tested in the NASA Langley Research Center Jet Exit Test Facility. Internal nozzle performance characteristics were defined for nozzle pressure ratios up to 10, with a range of secondary injection flow rates up to 10 percent of the primary flow rate. Most of the data included in this paper shows the effect of secondary injection rate at a nozzle pressure ratio of 4. The effects of modifying cavity divergence angle, convergence angle and cavity shape on internal nozzle performance were investigated, as were effects of injection geometry, hole or slot. In agreement with computationally predicted data, experimental data verified that decreasing cavity divergence angle had a negative impact and increasing cavity convergence angle had a positive impact on thrust vector angle and thrust efficiency. A curved cavity apex provided improved thrust ratios at some injection rates. However, overall nozzle performance suffered with no secondary injection. Injection holes were more efficient than the injection slot over the range of injection rates, but the slot generated larger thrust vector angles for injection rates less than 4 percent of the primary flow rate.

Thrust Augmentation of a Turbojet Engine at Simulated Flight Conditions by Introduction of a Water-Alcohol Mixture into the Compressor

NASA Technical Reports Server (NTRS)

Useller, James W.; Auble, Carmon M.; Harvey, Ray W., Sr.

1952-01-01

An investigation was conducted at simulated high-altitude flight conditions to evaluate the use of compressor evaporative cooling as a means of turbojet-engine thrust augmentation. Comparison of the performance of the engine with water-alcohol injection at the compressor inlet, at the sixth stage of the compressor, and at the sixth and ninth stages was made. From consideration of the thrust increases achieved, the interstage injection of the coolant was considered more desirable preferred over the combined sixth- and ninth-stage injection because of its relative simplicity. A maximum augmented net-thrust ratio of 1.106 and a maximum augmented jet-thrust ratio of 1.062 were obtained at an augmented liquid ratio of 2.98 and an engine-inlet temperature of 80 F. At lower inlet temperatures (-40 to 40 F), the maximum augmented net-thrust ratios ranged from 1.040 to 1.076 and the maximum augmented jet-thrust ratios ranged from 1.027 to 1.048, depending upon the inlet temperature. The relatively small increase in performance at the lower inlet-air temperatures can be partially attributed to the inadequate evaporation of the water-alcohol mixture, but the more significant limitation was believed to be caused by the negative influence of the liquid coolant on engine- component performance. In general, it is concluded that the effectiveness of the injection of a coolant into the compressor as a means of thrust augmentation is considerably influenced by the design characteristics of the components of the engine being used.

A Performance Comparison of Xenon and Krypton Propellant on an SPT-100 Hall Thruster (Preprint)

DTIC Science & Technology

2011-08-10

plume data from electrostatic probes. This paper presents the results of performance measurements made using an inverted pendulum thrust stand. Krypton...inverted pendulum thrust stand. Krypton operating conditions were tested over a large range of operating powers from 800 W to 3.9 kW. Analysis of how...advantages for missions where high thrust at reduced specific impulse is advantageous, primarily for orbit raising missions. Bismuth’s main drawback is

Real-time in-flight thrust calculation on a digital electronic engine control-equipped F100 engine in an F-15 airplane

NASA Technical Reports Server (NTRS)

Ray, R. J.; Myers, L. P.

1984-01-01

Computer algorithms which calculate in-flight engine and aircraft performance real-time are discussed. The first step was completed with the implementation of a real-time thrust calculation program on a digital electronic engine control (DEEC) equiped F100 engine in an F-15 aircraft. The in-flight thrust modifications that allow calculations to be performed in real-time, to compare results to predictions, are presented.

Comparison of Performance Predictions for New Low-Thrust Trajectory Tools

NASA Technical Reports Server (NTRS)

Polsgrove, Tara; Kos, Larry; Hopkins, Randall; Crane, Tracie

2006-01-01

Several low thrust trajectory optimization tools have been developed over the last 3% years by the Low Thrust Trajectory Tools development team. This toolset includes both low-medium fidelity and high fidelity tools which allow the analyst to quickly research a wide mission trade space and perform advanced mission design. These tools were tested using a set of reference trajectories that exercised each tool s unique capabilities. This paper compares the performance predictions of the various tools against several of the reference trajectories. The intent is to verify agreement between the high fidelity tools and to quantify the performance prediction differences between tools of different fidelity levels.

The kinematic determinants of anuran swimming performance: an inverse and forward dynamics approach.

PubMed

Richards, Christopher T

2008-10-01

The aims of this study were to explore the hydrodynamic mechanism of Xenopus laevis swimming and to describe how hind limb kinematics shift to control swimming performance. Kinematics of the joints, feet and body were obtained from high speed video of X. laevis frogs (N=4) during swimming over a range of speeds. A blade element approach was used to estimate thrust produced by both translational and rotational components of foot velocity. Peak thrust from the feet ranged from 0.09 to 0.69 N across speeds ranging from 0.28 to 1.2 m s(-1). Among 23 swimming strokes, net thrust impulse from rotational foot motion was significantly higher than net translational thrust impulse, ranging from 6.1 to 29.3 N ms, compared with a range of -7.0 to 4.1 N ms from foot translation. Additionally, X. laevis kinematics were used as a basis for a forward dynamic anuran swimming model. Input joint kinematics were modulated to independently vary the magnitudes of foot translational and rotational velocity. Simulations predicted that maximum swimming velocity (among all of the kinematics patterns tested) requires that maximal translational and maximal rotational foot velocity act in phase. However, consistent with experimental kinematics, translational and rotational motion contributed unequally to total thrust. The simulation powered purely by foot translation reached a lower peak stroke velocity than the pure rotational case (0.38 vs 0.54 m s(-1)). In all simulations, thrust from the foot was positive for the first half of the power stroke, but negative for the second half. Pure translational foot motion caused greater negative thrust (70% of peak positive thrust) compared with pure rotational simulation (35% peak positive thrust) suggesting that translational motion is propulsive only in the early stages of joint extension. Later in the power stroke, thrust produced by foot rotation overcomes negative thrust (due to translation). Hydrodynamic analysis from X. laevis as well as forward dynamics give insight into the differential roles of translational and rotational foot motion in the aquatic propulsion of anurans, providing a mechanistic link between joint kinematics and swimming performance.

Why do airlines want and use thrust reversers? A compilation of airline industry responses to a survey regarding the use of thrust reversers on commercial transport airplanes

NASA Technical Reports Server (NTRS)

Yetter, Jeffrey A.

1995-01-01

Although thrust reversers are used for only a fraction of the airplane operating time, their impact on nacelle design, weight, airplane cruise performance, and overall airplane operating and maintenance expenses is significant. Why then do the airlines want and use thrust reversers? In an effort to understand the airlines need for thrust reversers, a survey of the airline industry was made to determine why and under what situations thrust reversers are currently used or thought to be needed. The survey was intended to help establish the cost/benefits trades for the use of thrust reversers and airline opinion regarding alternative deceleration devices. A compilation and summary of the responses given to the survey questionnaire is presented.

Combined high and low-thrust geostationary orbit insertion with radiation constraint

NASA Astrophysics Data System (ADS)

Macdonald, Malcolm; Owens, Steven Robert

2018-01-01

The sequential use of an electric propulsion system is considered in combination with a high-thrust propulsion system for application to the propellant-optimal Geostationary Orbit insertion problem, whilst considering both temporal and radiation flux constraints. Such usage is found to offer a combined propellant mass saving when compared with an equivalent high-thrust only transfer. This propellant mass saving is seen to increase as the allowable transfer duration is increased, and as the thrust from the low-thrust system is increased, assuming constant specific impulse. It was found that the required plane change maneuver is most propellant-efficiently performed by the high-thrust system. The propellant optimal trajectory incurs a significantly increased electron flux when compared to an equivalent high-thrust only transfer. However, the electron flux can be reduced to a similar order of magnitude by increasing the high-thrust propellant consumption, whilst still delivering an improved mass fraction.

Static internal performance including thrust vectoring and reversing of two-dimensional convergent-divergent nozzles

NASA Technical Reports Server (NTRS)

Re, R. J.; Leavitt, L. D.

1984-01-01

The effects of geometric design parameters on two dimensional convergent-divergent nozzles were investigated at nozzle pressure ratios up to 12 in the static test facility. Forward flight (dry and afterburning power settings), vectored-thrust (afterburning power setting), and reverse-thrust (dry power setting) nozzles were investigated. The nozzles had thrust vector angles from 0 deg to 20.26 deg, throat aspect ratios of 3.696 to 7.612, throat radii from sharp to 2.738 cm, expansion ratios from 1.089 to 1.797, and various sidewall lengths. The results indicate that unvectored two dimensional convergent-divergent nozzles have static internal performance comparable to axisymmetric nozzles with similar expansion ratios.

Preliminary flight evaluation of an engine performance optimization algorithm

NASA Technical Reports Server (NTRS)

Lambert, H. H.; Gilyard, G. B.; Chisholm, J. D.; Kerr, L. J.

1991-01-01

A performance seeking control (PSC) algorithm has undergone initial flight test evaluation in subsonic operation of a PW 1128 engined F-15. This algorithm is designed to optimize the quasi-steady performance of an engine for three primary modes: (1) minimum fuel consumption; (2) minimum fan turbine inlet temperature (FTIT); and (3) maximum thrust. The flight test results have verified a thrust specific fuel consumption reduction of 1 pct., up to 100 R decreases in FTIT, and increases of as much as 12 pct. in maximum thrust. PSC technology promises to be of value in next generation tactical and transport aircraft.

Performance of Low-Power Pulsed Arcjets

NASA Technical Reports Server (NTRS)

Burton, Rodney L.

1995-01-01

The Electric Propulsion Laboratory at UIUC has in place all the capability and diagnostics required for performance testing of low power pulsed and DC arcjets. The UIUC thrust stand is operating with excellent accuracy and sensitivity at very low thrust levels. An important aspect of the experimental setup is the use of a PID controller to maintain a constant thruster position, which reduces hysterisis effects. Electrical noise from the arcjet induces some noise into the thrust signal, but this does not affect the measurement.

Static internal performance of an axisymmetric nozzle with multiaxis thrust-vectoring capability

NASA Technical Reports Server (NTRS)

Carson, George T., Jr.; Capone, Francis J.

1991-01-01

An investigation was conducted in the static test facility of the Langley 16 Foot Transonic Tunnel in order to determine the internal performance characteristics of a multiaxis thrust vectoring axisymmetric nozzle. Thrust vectoring for this nozzle was achieved by deflection of only the divergent section of this nozzle. The effects of nozzle power setting and divergent flap length were studied at nozzle deflection angles of 0 to 30 at nozzle pressure ratios up to 8.0.

A Method of Efficient Inclination Changes for Low-thrust Spacecraft

NASA Technical Reports Server (NTRS)

Falck, Robert; Gefert, Leon

2002-01-01

The evolution of low-thrust propulsion technologies has reached a point where such systems have become an economical option for many space missions. The development of efficient, low trip time control laws has received an increasing amount of attention in recent years, though few studies have examined the subject of inclination changing maneuvers in detail. A method for performing economical inclination changes through the use of an efficiency factor is derived front Lagrange's planetary equations. The efficiency factor can be used to regulate propellant expenditure at the expense of trip time. Such a method can be used for discontinuous-thrust transfers that offer reduced propellant masses and trip-times in comparison to continuous thrust transfers, while utilizing thrusters that operate at a lower specific impulse. Performance comparisons of transfers utilizing this approach with continuous-thrust transfers are generated through trajectory simulation and are presented in this paper.

Low-thrust chemical rocket engine study

NASA Technical Reports Server (NTRS)

Shoji, J. M.

1981-01-01

An analytical study evaluating thrust chamber cooling engine cycles and preliminary engine design for low thrust chemical rocket engines for orbit transfer vehicles is described. Oxygen/hydrogen, oxygen/methane, and oxygen/RP-1 engines with thrust levels from 444.8 N to 13345 N, and chamber pressures from 13.8 N/sq cm to 689.5 N/sq cm were evaluated. The physical and thermodynamic properties of the propellant theoretical performance data, and transport properties are documented. The thrust chamber cooling limits for regenerative/radiation and film/radiation cooling are defined and parametric heat transfer data presented. A conceptual evaluation of a number of engine cycles was performed and a 2224.1 N oxygen/hydrogen engine cycle configuration and a 2224.1 N oxygen/methane configuration chosen for preliminary engine design. Updated parametric engine data, engine design drawings, and an assessment of technology required are presented.

Thrust Augmentation with Mixer/Ejector Systems

NASA Technical Reports Server (NTRS)

Presz, Walter M., Jr.; Reynolds, Gary; Hunter, Craig

2002-01-01

Older commercial aircraft often exceed FAA (Federal Aviation Administration) sideline noise regulations. The major problem is the jet noise associated with the high exhaust velocities of the low bypass ratio engines on such aircraft. Mixer/ejector exhaust systems can provide a simple means of reducing the jet noise on these aircraft by mixing cool ambient air with the high velocity engine gases before they are exhausted to ambient. This paper presents new information on thrust performance predictions, and thrust augmentation capabilities of mixer/ejectors. Results are presented from the recent development program of the patented Alternating Lobe Mixer Ejector Concept (ALMEC) suppressor system for the Gulfstream GII, GIIB and GIII aircraft. Mixer/ejector performance procedures are presented which include classical control volume analyses, compound compressible flow theory, lobed nozzle loss correlations and state of the art computational fluid dynamic predictions. The mixer/ejector thrust predictions are compared to subscale wind tunnel test model data and actual aircraft flight test measurements. The results demonstrate that a properly designed mixer/ejector noise suppressor can increase effective engine bypass ratio and generate large thrust gains at takeoff conditions with little or no thrust loss at cruise conditions. The cruise performance obtained for such noise suppressor systems is shown to be a strong function of installation effects on the aircraft.

Liquid-hydrogen rocket engine development at Aerojet, 1944 - 1950

NASA Technical Reports Server (NTRS)

Osborn, G. H.; Gordon, R.; Coplen, H. L.; James, G. S.

1977-01-01

This program demonstrated the feasibility of virtually all the components in present-day, high-energy, liquid-rocket engines. Transpiration and film-cooled thrust chambers were successfully operated. The first liquid-hydrogen tests of the coaxial injector was conducted and the first pump to successfully produce high pressures in pumping liquid hydrogen was tested. A 1,000-lb-thrust gaseous propellant and a 3,000-lb-thrust liquid-propellant thrust chamber were operated satisfactorily. Also, the first tests were conducted to evaluate the effects of jet overexpansion and separation on performance of rocket thrust chambers with hydrogen-oxygen propellants.

Thrust Stand for Vertically Oriented Electric Propulsion Performance Evaluation

NASA Technical Reports Server (NTRS)

Moeller, Trevor; Polzin, Kurt A.

2010-01-01

A variation of a hanging pendulum thrust stand capable of measuring the performance of an electric thruster operating in the vertical orientation is presented. The vertical orientation of the thruster dictates that the thruster must be horizontally offset from the pendulum pivot arm, necessitating the use of a counterweight system to provide a neutrally-stable system. Motion of the pendulum arm is transferred through a balance mechanism to a secondary arm on which deflection is measured. A non-contact light-based transducer is used to measure displacement of the secondary beam. The members experience very little friction, rotating on twisting torsional pivots with oscillatory motion attenuated by a passive, eddy current damper. Displacement is calibrated using an in situ thrust calibration system. Thermal management and self-leveling systems are incorporated to mitigate thermal and mechanical drifts. Gravitational restoring force and torsional spring constants associated with flexure pivots provide restoring moments. An analysis of the design indicates that the thrust measurement range spans roughly four decades, with the stand capable of measuring thrust up to 12 N for a 200 kg thruster and up to approximately 800 mN for a 10 kg thruster. Data obtained from calibration tests performed using a 26.8 lbm simulated thruster indicated a resolution of 1 mN on 100 mN-level thrusts, while those tests conducted on 200 lbm thruster yielded a resolution of roughly 2.5 micro at thrust levels of 0.5 N and greater.

Thrust stand for vertically oriented electric propulsion performance evaluation

DOE Office of Scientific and Technical Information (OSTI.GOV)

Moeller, Trevor; Polzin, Kurt A.

A variation of a hanging pendulum thrust stand capable of measuring the performance of an electric thruster operating in the vertical orientation is presented. The vertical orientation of the thruster dictates that the thruster must be horizontally offset from the pendulum pivot arm, necessitating the use of a counterweight system to provide a neutrally stable system. Motion of the pendulum arm is transferred through a balance mechanism to a secondary arm on which deflection is measured. A noncontact light-based transducer is used to measure displacement of the secondary beam. The members experience very little friction, rotating on twisting torsional pivotsmore » with oscillatory motion attenuated by a passive, eddy-current damper. Displacement is calibrated using an in situ thrust calibration system. Thermal management and self-leveling systems are incorporated to mitigate thermal and mechanical drifts. Gravitational force and torsional spring constants associated with flexure pivots provide restoring moments. An analysis of the design indicates that the thrust measurement range spans roughly four decades, with the stand capable of measuring thrust up to 12 N for a 200 kg thruster and up to approximately 800 mN for a 10 kg thruster. Data obtained from calibration tests performed using a 26.8 lbm simulated thruster indicated a resolution of 1 mN on 100 mN level thrusts, while those tests conducted on a 200 lbm thruster yielded a resolution of roughly 2.5 mN at thrust levels of 0.5 N and greater.« less

Performance of Simple Gas Foil Thrust Bearings in Air

NASA Technical Reports Server (NTRS)

Bruckner, Robert J.

2012-01-01

Foil bearings are self-acting hydrodynamics devices used to support high speed rotating machinery. The advantages that they offer to process fluid lubricated machines include: high rotational speed capability, no auxiliary lubrication system, non-contacting high speed operation, and improved damping as compared to rigid hydrodynamic bearings. NASA has had a sporadic research program in this technology for almost 6 decades. Advances in the technology and understanding of foil journal bearings have enabled several new commercial products in recent years. These products include oil-free turbochargers for both heavy trucks and automobiles, high speed electric motors, microturbines for distributed power generation, and turbojet engines. However, the foil thrust bearing has not received a complimentary level of research and therefore has become the weak link of oil-free turbomachinery. In an effort to both provide machine designers with basic performance parameters and to elucidate the underlying physics of foil thrust bearings, NASA Glenn Research Center has completed an effort to experimentally measure the performance of simple gas foil thrust bearing in air. The database includes simple bump foil supported thrust bearings with full geometry and manufacturing techniques available to the user. Test conditions consist of air at ambient pressure and temperatures up to 500 C and rotational speeds to 55,000 rpm. A complete set of axial load, frictional torque, and rotational speed is presented for two different compliant sub-structures and inter-pad gaps. Data obtained from commercially available foil thrust bearings both with and without active cooling is presented for comparison. A significant observation made possible by this data set is the speed-load capacity characteristic of foil thrust bearings. Whereas for the foil journal bearing the load capacity increases linearly with rotational speed, the foil thrust bearing operates in the hydrodynamic high speed limit. In this case, the load capacity is constant and in fact often decreases with speed if other factors such as thermal conditions and runner distortions are permitted to dominate the bearing performance.

Advanced Manufacturing Technologies (AMT): Additive Manufactured Hot Fire Planning and Testing in GRC Cell 32 Project

NASA Technical Reports Server (NTRS)

Fikes, John C.

2014-01-01

The objective of this project is to hot fire test an additively manufactured thrust chamber assembly TCA (injector and thrust chamber). GRC will install the additively manufactured Inconel 625 injector, two additively manufactured (SLM) water cooled Cu-Cr thrust chamber barrels and one additively manufactured (SLM) water cooled Cu-Cr thrust chamber nozzle on the test stand in Cell 32 and perform hot fire testing of the integrated TCA.

Comparison of Performance and Component Frontal Areas of Hypothetical Two-spool and One-spool Turbojet Engines

NASA Technical Reports Server (NTRS)

Dugan, James F , Jr

1956-01-01

For constant-mechanical-speed operation, the two-spool thrust values are as great as or greater than the one-spool thrust values over the entire flight range considered, while the specific fuel consumption for the two engines agrees within 1 percent. The maximum difference in thrust occurs at Mach 2.8 in the stratosphere, where the two-spool thrust advantage is about 9 percent for operation with the after burning.

Static Performance of a Wing-Mounted Thrust Reverser Concept

NASA Technical Reports Server (NTRS)

Asbury, Scott C.; Yetter, Jeffrey A.

1998-01-01

An experimental investigation was conducted in the Jet-Exit Test Facility at NASA Langley Research Center to study the static aerodynamic performance of a wing-mounted thrust reverser concept applicable to subsonic transport aircraft. This innovative engine powered thrust reverser system is designed to utilize wing-mounted flow deflectors to produce aircraft deceleration forces. Testing was conducted using a 7.9%-scale exhaust system model with a fan-to-core bypass ratio of approximately 9.0, a supercritical left-hand wing section attached via a pylon, and wing-mounted flow deflectors attached to the wing section. Geometric variations of key design parameters investigated for the wing-mounted thrust reverser concept included flow deflector angle and chord length, deflector edge fences, and the yaw mount angle of the deflector system (normal to the engine centerline or parallel to the wing trailing edge). All tests were conducted with no external flow and high pressure air was used to simulate core and fan engine exhaust flows. Test results indicate that the wing-mounted thrust reverser concept can achieve overall thrust reverser effectiveness levels competitive with (parallel mount), or better than (normal mount) a conventional cascade thrust reverser system. By removing the thrust reverser system from the nacelle, the wing-mounted concept offers the nacelle designer more options for improving nacelle aero dynamics and propulsion-airframe integration, simplifying nacelle structural designs, reducing nacelle weight, and improving engine maintenance access.

Characterization of in-flight performance of ion propulsion systems

NASA Astrophysics Data System (ADS)

Sovey, James S.; Rawlin, Vincent K.

1993-06-01

In-flight measurements of ion propulsion performance, ground test calibrations, and diagnostic performance measurements were reviewed. It was found that accelerometers provided the most accurate in-flight thrust measurements compared with four other methods that were surveyed. An experiment has also demonstrated that pre-flight alignment of the thrust vector was sufficiently accurate so that gimbal adjustments and use of attitude control thrusters were not required to counter disturbance torques caused by thrust vector misalignment. The effects of facility background pressure, facility enhanced charge-exchange reactions, and contamination on ground-based performance measurements are also discussed. Vacuum facility pressures for inert-gas ion thruster life tests and flight qualification tests will have to be less than 2 mPa to ensure accurate performance measurements.

Characterization of in-flight performance of ion propulsion systems

NASA Technical Reports Server (NTRS)

Sovey, James S.; Rawlin, Vincent K.

1993-01-01

In-flight measurements of ion propulsion performance, ground test calibrations, and diagnostic performance measurements were reviewed. It was found that accelerometers provided the most accurate in-flight thrust measurements compared with four other methods that were surveyed. An experiment has also demonstrated that pre-flight alignment of the thrust vector was sufficiently accurate so that gimbal adjustments and use of attitude control thrusters were not required to counter disturbance torques caused by thrust vector misalignment. The effects of facility background pressure, facility enhanced charge-exchange reactions, and contamination on ground-based performance measurements are also discussed. Vacuum facility pressures for inert-gas ion thruster life tests and flight qualification tests will have to be less than 2 mPa to ensure accurate performance measurements.

Dielectric Barrier Discharge (DBD) Plasma Actuators Thrust-Measurement Methodology Incorporating New Anti-Thrust Hypothesis

NASA Technical Reports Server (NTRS)

Ashpis, David E.; Laun, Matthew C.

2014-01-01

We discuss thrust measurements of Dielectric Barrier Discharge (DBD) plasma actuators devices used for aerodynamic active flow control. After a review of our experience with conventional thrust measurement and significant non-repeatability of the results, we devised a suspended actuator test setup, and now present a methodology of thrust measurements with decreased uncertainty. The methodology consists of frequency scans at constant voltages. The procedure consists of increasing the frequency in a step-wise fashion from several Hz to the maximum frequency of several kHz, followed by frequency decrease back down to the start frequency of several Hz. This sequence is performed first at the highest voltage of interest, then repeated at lower voltages. The data in the descending frequency direction is more consistent and selected for reporting. Sample results show strong dependence of thrust on humidity which also affects the consistency and fluctuations of the measurements. We also observed negative values of thrust or "anti-thrust", at low frequencies between 4 Hz and up to 64 Hz. The anti-thrust is proportional to the mean-squared voltage and is frequency independent. Departures from the parabolic anti-thrust curve are correlated with appearance of visible plasma discharges. We propose the anti-thrust hypothesis. It states that the measured thrust is a sum of plasma thrust and anti-thrust, and assumes that the anti-thrust exists at all frequencies and voltages. The anti-thrust depends on actuator geometry and materials and on the test installation. It enables the separation of the plasma thrust from the measured total thrust. This approach enables more meaningful comparisons between actuators at different installations and laboratories. The dependence on test installation was validated by surrounding the actuator with a large diameter, grounded, metal sleeve.

The 30-centimeter ion thrust subsystem design manual

NASA Technical Reports Server (NTRS)

1979-01-01

The principal characteristics of the 30-centimeter ion propulsion thrust subsystem technology that was developed to satisfy the propulsion needs of future planetary and early orbital missions are described. Functional requirements and descriptions, interface and performance requirements, and physical characteristics of the hardware are described at the thrust subsystem, BIMOD engine system, and component level.

Acoustic and aerodynamic performance of a variable-pitch 1.83-meter-(6-ft) diameter 1.20-pressure-ratio fan stage (QF-9)

NASA Technical Reports Server (NTRS)

Glaser, F. W.; Woodward, R. P.; Lucas, J. G.

1977-01-01

Far field noise data and related aerodynamic performance are presented for a variable pitch fan stage having characteristics suitable for low noise, STOL engine application. However, no acoustic suppression material was used in the flow passages. The fan was externally driven by an electric motor. Tests were made at several forward thrust rotor blade pitch angles and one for reverse thrust. Fan speed was varied from 60 to 120 percent of takeoff (design) speed, and exhaust nozzles having areas 92 to 105 percent of design were tested. The fan noise level was at a minimum at the design rotor blade pitch angles of 64 deg for takeoff thrust and at 57 deg for approach (50 percent takeoff thrust). Perceived noise along a 152.4-m sideline reached 100.1 PNdb for the takeoff (design) configuration for a stage pressure ratio of 1.17 and thrust of 57,600 N. For reverse thrust the PNL values were 4 to 5 PNdb above the takeoff values at comparable fan speeds.

Thrust stand evaluation of engine performance improvement algorithms in an F-15 airplane

NASA Technical Reports Server (NTRS)

Conners, Timothy R.

1992-01-01

Results are presented from the evaluation of the performance seeking control (PSC) optimization algorithm developed by Smith et al. (1990) for F-15 aircraft, which optimizes the quasi-steady-state performance of an F100 derivative turbofan engine for several modes of operation. The PSC algorithm uses onboard software engine model that calculates thrust, stall margin, and other unmeasured variables for use in the optimization. Comparisons are presented between the load cell measurements, PSC onboard model thrust calculations, and posttest state variable model computations. Actual performance improvements using the PSC algorithm are presented for its various modes. The results of using PSC algorithm are compared with similar test case results using the HIDEC algorithm.

Internal Performance Evaluation of a Two Position Divergent Shroud Ejector

NASA Technical Reports Server (NTRS)

Mihaloew, James R.; Stofan, Andrew J.

1960-01-01

A two-position divergent shroud ejector was investigated in an unheated quiescent-air facility over a range of operational variables applicable to a Mach 2.5 aircraft. The performance data are shown in terms of hypothetical engine operating conditions to illustrate variations of performance with Mach number. The overall thrust performance was reasonably good, with ejector thrust ratios ranging from 0.97 to 0.98 for all conditions except that corresponding to acceleration with afterburning through the transonic flight Mach number region from 0.9 to 1.1, where the ejector thrust ratio decreased to as low as 0.945 for an ejector corrected weight-flow ratio of 0.105.

Experimental results for a two-dimensional supersonic inlet used as a thrust deflecting nozzle

NASA Technical Reports Server (NTRS)

Johns, Albert L.; Burstadt, Paul L.

1984-01-01

Nearly all supersonic V/STOL aircraft concepts are dependent on the thrust deflecting capability of a nozzle. In one unique concept, referred to as the reverse flow dual fan, not only is there a thrust deflecting nozzle for the fan and core engine exit flow, but because of the way the propulsion system operates during vertical takeoff and landing, the supersonic inlet is also used as a thrust deflecting nozzle. This paper presents results of an experimental study to evaluate the performance of a supersonic inlet used as a thrust deflecting nozzle for this reverse flow dual fan concept. Results are presented in terms of nozzle thrust coefficient and thrust vector angle for a number of inlet/nozzle configurations. Flow visualization and nozzle exit flow survey results are also shown.

Direct measurement of the impulse in a magnetic thrust chamber system for laser fusion rocket

DOE Office of Scientific and Technical Information (OSTI.GOV)

Maeno, Akihiro; Yamamoto, Naoji; Nakashima, Hideki

2011-08-15

An experiment is conducted to measure an impulse for demonstrating a magnetic thrust chamber system for laser fusion rocket. The impulse is produced by the interaction between plasma and magnetic field. In the experiment, the system consists of plasma and neodymium permanent magnets. The plasma is created by a single-beam laser aiming at a polystyrene spherical target. The impulse is 1.5 to 2.2 {mu}Ns by means of a pendulum thrust stand, when the laser energy is 0.7 J. Without magnetic field, the measured impulse is found to be zero. These results indicate that the system for generating impulse is working.

Effect of varying internal geometry on the static performance of rectangular thrust-reverser ports

NASA Technical Reports Server (NTRS)

Re, Richard J.; Mason, Mary L.

1987-01-01

An investigation has been conducted to evaluate the effects of several geometric parameters on the internal performance of rectangular thrust-reverser ports for nonaxisymmetric nozzles. Internal geometry was varied with a test apparatus which simulated a forward-flight nozzle with a single, fully deployed reverser port. The test apparatus was designed to simulate thrust reversal (conceptually) either in the convergent section of the nozzle or in the constant-area duct just upstream of the nozzle. The main geometric parameters investigated were port angle, port corner radius, port location, and internal flow blocker angle. For all reverser port geometries, the port opening had an aspect ratio (throat width to throat height) of 6.1 and had a constant passage area from the geometric port throat to the exit. Reverser-port internal performance and thrust-vector angles computed from force-balance measurements are presented.

Internal performance of two nozzles utilizing gimbal concepts for thrust vectoring

NASA Technical Reports Server (NTRS)

Berrier, Bobby L.; Taylor, John G.

1990-01-01

The internal performance of an axisymmetric convergent-divergent nozzle and a nonaxisymmetric convergent-divergent nozzle, both of which utilized a gimbal type mechanism for thrust vectoring was evaluated in the Static Test Facility of the Langley 16-Foot Transonic Tunnel. The nonaxisymmetric nozzle used the gimbal concept for yaw thrust vectoring only; pitch thrust vectoring was accomplished by simultaneous deflection of the upper and lower divergent flaps. The model geometric parameters investigated were pitch vector angle for the axisymmetric nozzle and pitch vector angle, yaw vector angle, nozzle throat aspect ratio, and nozzle expansion ratio for the nonaxisymmetric nozzle. All tests were conducted with no external flow, and nozzle pressure ratio was varied from 2.0 to approximately 12.0.

Measurements versus Predictions for a Hybrid (Hydrostatic Plus Hydrodynamic) Thrust Bearing for a Range of Orifice Diameters

DTIC Science & Technology

2010-05-01

Type of Lubrication for a Tilting Pad Thrust Bearing ,” ASME Journal of Lubrication Technology, 96 Ser F (1), pp. 22-27. [9] Gregory, R.S., 1974...1986, “Measurements of Maximum Temperature in Tilting - Pad Thrust Bearings ,” Technical Preprints - Presented at the ASLE 41st Annual Meeting. (ASLE...Safar [7] provides a modified Reynolds number analysis on hydrostatic thrust bearing performance parameters including the effects of tilt . Finally, San

Comparison Between Predictions and Measurements of Performance Characteristics for an Eight Pocket Hybrid (Combination Hydrostatic and Hydrodynamic) Thrust Bearing

DTIC Science & Technology

2008-12-03

STB stainless steel housing. Drive End Thrust Collar Free End Splash Guard Slave Thrust Bearing Face STB Flange Figure 12: Exploded view of slave...Bulk Flow Analysis of Hybrid Thrust Bearing for Process Fluid Applications.” ASME Trans., Journal of Tribology , January, Vol. 122 (1), pp. 170-180...ASME Journal of Tribology , Vol. 123, pp. 768-776 [5] San Andrés, L., (2006), “MEEN 626 Notes 2 Derivation of the Classical Reynolds Equation for

Static internal performance of a two-dimensional convergent nozzle with thrust-vectoring capability up to 60 deg

NASA Technical Reports Server (NTRS)

Leavitt, L. D.

1985-01-01

An investigation was conducted at wind-off conditions in the static-test facility of the Langley 16-Foot Transonic Tunnel to determine the internal performance characteristics of a two-dimensional convergent nozzle with a thrust-vectoring capability up to 60 deg. Vectoring was accomplished by a downward rotation of a hinged upper convergent flap and a corresponding rotation of a center-pivoted lower convergent flap. The effects of geometric thrust-vector angle and upper-rotating-flap geometry on internal nozzle performance characteristics were investigated. Nozzle pressure ratio was varied from 1.0 (jet off) to approximately 5.0.

Maximum thrust mode evaluation

NASA Technical Reports Server (NTRS)

Orme, John S.; Nobbs, Steven G.

1995-01-01

Measured reductions in acceleration times which resulted from the application of the F-15 performance seeking control (PSC) maximum thrust mode during the dual-engine test phase is presented as a function of power setting and flight condition. Data were collected at altitudes of 30,000 and 45,000 feet at military and maximum afterburning power settings. The time savings for the supersonic acceleration is less than at subsonic Mach numbers because of the increased modeling and control complexity. In addition, the propulsion system was designed to be optimized at the mid supersonic Mach number range. Recall that even though the engine is at maximum afterburner, PSC does not trim the afterburner for the maximum thrust mode. Subsonically at military power, time to accelerate from Mach 0.6 to 0.95 was cut by between 6 and 8 percent with a single engine application of PSC, and over 14 percent when both engines were optimized. At maximum afterburner, the level of thrust increases were similar in magnitude to the military power results, but because of higher thrust levels at maximum afterburner and higher aircraft drag at supersonic Mach numbers the percentage thrust increase and time to accelerate was less than for the supersonic accelerations. Savings in time to accelerate supersonically at maximum afterburner ranged from 4 to 7 percent. In general, the maximum thrust mode has performed well, demonstrating significant thrust increases at military and maximum afterburner power. Increases of up to 15 percent at typical combat-type flight conditions were identified. Thrust increases of this magnitude could be useful in a combat situation.

An analysis of the effect of aeroassist maneuvers on orbital transfer vehicle performance

NASA Technical Reports Server (NTRS)

Murphy, Gregory O.; Suit, William T.

1987-01-01

This paper summarizes a Langley Research Summer Scholars (LARSS) research project (Summer 1986) dealing with the topic of the effectiveness of aeroassist maneuvers to accomplish a change in the orbital inclination of an Orbital Transfer Vehicle (OTV). This task was subject to OTV design constraints, chief of which were the axial acceleration and the aerodynamic heating rate limits of the OTV. The use of vehicle thrust to replace lost kinetic energy and, thereby, to increase the maximum possible change in orbital inclination was investigated. A relation between time in the hover orbit and payload to LEO was established. The amount of plane change possible during this type of maneuver was checked for several runs and a possible thrusting procedure to increase the plane change and still get to LEO was suggested. Finally, the sensitivity of various target parameters to controllable independent variables was established, trades between the amount of control allowed, and payload to LEO suggested.

Improved Propulsion Modeling for Low-Thrust Trajectory Optimization

NASA Technical Reports Server (NTRS)

Knittel, Jeremy M.; Englander, Jacob A.; Ozimek, Martin T.; Atchison, Justin A.; Gould, Julian J.

2017-01-01

Low-thrust trajectory design is tightly coupled with spacecraft systems design. In particular, the propulsion and power characteristics of a low-thrust spacecraft are major drivers in the design of the optimal trajectory. Accurate modeling of the power and propulsion behavior is essential for meaningful low-thrust trajectory optimization. In this work, we discuss new techniques to improve the accuracy of propulsion modeling in low-thrust trajectory optimization while maintaining the smooth derivatives that are necessary for a gradient-based optimizer. The resulting model is significantly more realistic than the industry standard and performs well inside an optimizer. A variety of deep-space trajectory examples are presented.

Thrust augmentation nozzle (TAN) concept for rocket engine booster applications

NASA Astrophysics Data System (ADS)

Forde, Scott; Bulman, Mel; Neill, Todd

2006-07-01

Aerojet used the patented thrust augmented nozzle (TAN) concept to validate a unique means of increasing sea-level thrust in a liquid rocket booster engine. We have used knowledge gained from hypersonic Scramjet research to inject propellants into the supersonic region of the rocket engine nozzle to significantly increase sea-level thrust without significantly impacting specific impulse. The TAN concept overcomes conventional engine limitations by injecting propellants and combusting in an annular region in the divergent section of the nozzle. This injection of propellants at moderate pressures allows for obtaining high thrust at takeoff without overexpansion thrust losses. The main chamber is operated at a constant pressure while maintaining a constant head rise and flow rate of the main propellant pumps. Recent hot-fire tests have validated the design approach and thrust augmentation ratios. Calculations of nozzle performance and wall pressures were made using computational fluid dynamics analyses with and without thrust augmentation flow, resulting in good agreement between calculated and measured quantities including augmentation thrust. This paper describes the TAN concept, the test setup, test results, and calculation results.

Assessing the Mirror Fusion Propulsion System (MFPS) Concept as Applied to Outer-Solar-System (OSS) Missions

NASA Astrophysics Data System (ADS)

Carpenter, Scott A.; Deveny, Marc E.; Schulze, Norman R.; Gatti, Raymond C.; Peters, Micheal B.

1994-07-01

In this paper, we strive to achieve three goals: (1) to describe a continuous-thrusting space-fusion-propulsion engine called the Mirror Fusion Propulsion System (MFPS), (2) to describe MFPS' ability to accomplish two candidate outer-solar-system (OSS) missions using various levels of advanced technology identified in the laboratory, and (3) to describe some interesting safety features of MFPS that include continuous mission-abort capability, magnetic-field-shielding against solar particle events (SPE), and performance of in-orbit characterization of the target body's natural resources (prior to human landings) using fusion-neutrons, x-rays, and possibly the neutralized thrust beam. The first OSS mission discussed is a mission to the Saturnian system, primarily exploration and resource- characterization driven, with emphasis on minimizing the Earth-to-Saturn and return-trip flight times. The other OSS mission discussed is an economically-driven mission to Uranus, stopping first to perform in-orbit resource characterization of the major moons of Uranus prior to human landing, and then returning to earth with a payload consisting of 3He (removed from the Uranian atmosphere or extracted from the Uranian moons) to be used in a future earth-based fusion-power industry.

NASA Double Asteroid Redirection Test (Dart) Trajectory Validation and Robustness

NASA Technical Reports Server (NTRS)

Sarli, Bruno V.; Ozimek, Martin T.; Atchison, Justin A.; Englander, Jacob A.; Barbee, Brent W.

2017-01-01

The Double Asteroid Redirection Test (DART) mission will be the first to test the concept of a kinetic impactor. Several studies have been made on asteroid redirection and impact mitigation, however, to this date no mission tested the proposed concepts. An impact study on a representative body allows the measurement of the effects on the target's orbit and physical structure. With this goal, DART's objective is to verify the effectiveness of the kinetic impact concept for planetary defense. The spacecraft uses solar electric propulsion to escape Earth, flyby (138971) 2001 CB21 for impart rehearsal, and impact the secondary body of the (65803) Didymos system. This work focuses on the interplanetary trajectory design part of the mission with the validation of the baseline trajectory, performance comparison to other mission objectives, and assessment of the baseline robustness to missed thrust events. Results show a good performance of the selected trajectory for different mission objectives: latest possible escape date, maximum kinetic energy on impact, shortest possible time of flight, and use of an Earth swing-by. The baseline trajectory was shown to be robust to a missed thrust with 1% of fuel margin being enough to recover the mission for failures of more than 14 days.

Recent advances in low-thrust propulsion technology

NASA Technical Reports Server (NTRS)

Stone, James R.

1988-01-01

The NASA low-thrust propulsion technology program is aimed at providing high performance options to a broad class of near-term and future missions. Major emphases of the program are on storable and hydrogen/oxygen low-thrust chemical, low-power (auxiliary) electrothermal, and high-power electric propulsion. This paper represents the major accomplishments of the program and discusses their impact.

Performance seeking control: Program overview and future directions

NASA Technical Reports Server (NTRS)

Gilyard, Glenn B.; Orme, John S.

1993-01-01

A flight test evaluation of the performance-seeking control (PSC) algorithm on the NASA F-15 highly integrated digital electronic control research aircraft was conducted for single-engine operation at subsonic and supersonic speeds. The model-based PSC system was developed with three optimization modes: minimum fuel flow at constant thrust, minimum turbine temperature at constant thrust, and maximum thrust at maximum dry and full afterburner throttle settings. Subsonic and supersonic flight testing were conducted at the NASA Dryden Flight Research Facility covering the three PSC optimization modes and over the full throttle range. Flight results show substantial benefits. In the maximum thrust mode, thrust increased up to 15 percent at subsonic and 10 percent at supersonic flight conditions. The minimum fan turbine inlet temperature mode reduced temperatures by more than 100 F at high altitudes. The minimum fuel flow mode results decreased fuel consumption up to 2 percent in the subsonic regime and almost 10 percent supersonically. These results demonstrate that PSC technology can benefit the next generation of fighter or transport aircraft. NASA Dryden is developing an adaptive aircraft performance technology system that is measurement based and uses feedback to ensure optimality. This program will address the technical weaknesses identified in the PSC program and will increase performance gains.

Internal performance characteristics of short convergent-divergent exhaust nozzles designed by the method of characteristics

NASA Technical Reports Server (NTRS)

Krull, H George; Beale, William T

1956-01-01

Internal performance data on a short exhaust nozzle designed by the method of characteristics were obtained over a range of pressure ratios from 1.5 to 22. The peak thrust coefficient was not affected by a shortened divergent section, but it occurred at lower pressure ratios due to reduction in expansion ratio. This nozzle contour based on characteristics solution gave higher thrust coefficients than a conical convergent-divergent nozzle of equivalent length. Abrupt-inlet sections permitted a reduction in nozzle length without a thrust-coefficient reduction.

Direct thrust measurements and modelling of a radio-frequency expanding plasma thruster

DOE Office of Scientific and Technical Information (OSTI.GOV)

Lafleur, T.; Charles, C.; Boswell, R. W.

2011-08-15

It is shown analytically that the thrust from a simple plasma thruster (in the absence of a magnetic field) is given by the maximum upstream electron pressure, even if the plasma diverges downstream. Direct thrust measurements of a thruster are then performed using a pendulum thrust balance and a laser displacement sensor. A maximum thrust of about 2 mN is obtained at 700 W for a thruster length of 17.5 cm and a flow rate of 0.9 mg s{sup -1}, while a larger thrust of 4 mN is obtained at a similar power for a length of 9.5 cm andmore » a flow rate of 1.65 mg s{sup -1}. The measured thrusts are in good agreement with the maximum upstream electron pressure found from measurements of the plasma parameters and in fair agreement with a simple global approach used to model the thruster.« less

Design, Fabrication, and Performance of Foil Gas Thrust Bearings for Microturbomachinery Applications

NASA Technical Reports Server (NTRS)

Dykas, Brian; Bruckner, Robert; DellaCorte, Christopher; Edmonds, Brian; Prahl, Joseph

2008-01-01

A methodology for the design and construction of simple foil thrust bearings intended for parametric performance testing and low marginal costs is presented. Features drawn from a review of the open literature are discussed as they relate to bearing performance. The design of fixtures and tooling required to fabricate foil thrust bearings is presented, using conventional machining processes where possible. A prototype bearing with dimensions drawn from the literature is constructed, with all fabrication steps described. A load-deflection curve for the bearing is presented to illustrate structural stiffness characteristics. Start-top cycles are performed on the bearing at a temperature of 425 C to demonstrate early-life wear patterns. A test of bearing load capacity demonstrates useful performance when compared with data obtained from the open literature.

Amplification of radar and lidar signatures using quantum sensors

NASA Astrophysics Data System (ADS)

Lanzagorta, Marco

2013-05-01

One of the major scientific thrusts from recent years has been to try to harness quantum phenomena to dramat­ ically increase the performance of a wide variety of classical devices. These advances in quantum information science have had a considerable impact on the development of photonic-based quantum sensors. Even though quantum radar and quantum lidar remain theoretical proposals, preliminary results suggest that these sensors have the potential of becoming disruptive technologies able to revolutionize reconnaissance systems. In this paper we will discuss how quantum entanglement can be exploited to increase the radar and lidar signature of rectangular targets. In particular, we will show how the effective visibility of the target is increased if observed with an entangled multi-photon quantum sensor.

Scaling of Performance in Liquid Propellant Rocket Engine Combustors

NASA Technical Reports Server (NTRS)

Hulka, James

2008-01-01

The objectives are: a) Re-introduce to you the concept of scaling; b) Describe the scaling research conducted in the 1950s and early 1960s, and present some of their conclusions; c) Narrow the focus to scaling for performance of combustion devices for liquid propellant rocket engines; and d) Present some results of subscale to full-scale performance from historical programs. Scaling is "The ability to develop new combustion devices with predictable performance on the basis of test experience with old devices." Scaling can be used to develop combustion devices of any thrust size from any thrust size. Scaling is applied mostly to increase thrust. Objective is to use scaling as a development tool. - Move injector design from an "art" to a "science"

The design and operating characteristics of an advanced 30-kW ammonia arcjet engine

NASA Technical Reports Server (NTRS)

Deininger, William D.; Pivirotto, Thomas J.; Brophy, John R.

1987-01-01

Experimental investigations were conducted to evaluate the effects of a contoured nozzle and modified cathode shape on ammonia arcjet engine performance. The contoured nozzle performance data were compared to the performance data of an arcjet which had a 38-deg included-angle, conical nozzle. Thrust improvements of up to 10 percent were demonstrated which corresponded to 3 percent improvements in specific impulse and 10 percent improvements in thrust efficiency. Performance characterizations for the modified cathode tip were conducted with the contoured nozzle arcjet. A uniform 15 percent decrease in arc voltage was demonstrated over a mass flow range of 0.175 to 0.350 g/s. A 4 percent improvement in thrust efficiency was noted at 22.0 kW.

Performance Cycle Analysis of a Two-Spool, Separate-Exhaust Turbofan With Interstage Turbine Burner

NASA Technical Reports Server (NTRS)

Liew, K. H.; Urip, E.; Yang, S. L.; Mattingly, J. D.; Marek, C. J.

2005-01-01

This paper presents the performance cycle analysis of a dual-spool, separate-exhaust turbofan engine, with an Interstage Turbine Burner serving as a secondary combustor. The ITB, which is located at the transition duct between the high- and the low-pressure turbines, is a relatively new concept for increasing specific thrust and lowering pollutant emissions in modern jet engine propulsion. A detailed performance analysis of this engine has been conducted for steady-state engine performance prediction. A code is written and is capable of predicting engine performances (i.e., thrust and thrust specific fuel consumption) at varying flight conditions and throttle settings. Two design-point engines were studied to reveal trends in performance at both full and partial throttle operations. A mission analysis is also presented to assure the advantage of saving fuel by adding ITB.

A study of variable thrust, variable specific impulse trajectories for solar system exploration

NASA Astrophysics Data System (ADS)

Sakai, Tadashi

A study has been performed to determine the advantages and disadvantages of variable thrust and variable Isp (specific impulse) trajectories for solar system exploration. There have been several numerical research efforts for variable thrust, variable Isp, power-limited trajectory optimization problems. All of these results conclude that variable thrust, variable Isp (variable specific impulse, or VSI) engines are superior to constant thrust, constant Isp (constant specific impulse; or CSI) engines. However, most of these research efforts assume a mission from Earth to Mars, and some of them further assume that these planets are circular and coplanar. Hence they still lack the generality. This research has been conducted to answer the following questions: (1) Is a VSI engine always better than a CSI engine or a high thrust engine for any mission to any planet with any time of flight considering lower propellant mass as the sole criterion? (2) If a planetary swing-by is used for a VSI trajectory, is the fuel savings of a VSI swing-by trajectory better than that of a CSI swing-by or high thrust swing-by trajectory? To support this research, an unique, new computer-based interplanetary trajectory calculation program has been created. This program utilizes a calculus of variations algorithm to perform overall optimization of thrust, Isp, and thrust vector direction along a trajectory that minimizes fuel consumption for interplanetary travel. It is assumed that the propulsion system is power-limited, and thus the compromise between thrust and Isp is a variable to be optimized along the flight path. This program is capable of optimizing not only variable thrust trajectories but also constant thrust trajectories in 3-D space using a planetary ephemeris database. It is also capable of conducting planetary swing-bys. Using this program, various Earth-originating trajectories have been investigated and the optimized results have been compared to traditional CSI and high thrust trajectory solutions. Results show that VSI rocket engines reduce fuel requirements for any mission compared to CSI rocket engines. Fuel can be saved by applying swing-by maneuvers for VSI engines; but the effects of swing-bys due to VSI engines are smaller than that of CSI or high thrust engines.

Effects of bleed air extraction on thrust levels on the F404-GE-400 turbofan engine

NASA Technical Reports Server (NTRS)

Yuhas, Andrew J.; Ray, Ronald J.

1992-01-01

A ground test was performed to determine the effects of compressor bleed flow extraction on the performance of F404-GE-400 afterburning turbofan engines. The two engines were installed in the F/A-18 High Alpha Research Vehicle at the NASA Dryden Flight Research Facility. A specialized bleed ducting system was installed onto the aircraft to control and measure engine bleed airflow while the aircraft was tied down to a thrust measuring stand. The test was conducted on each engine and at various power settings. The bleed air extraction levels analyzed included flow rates above the manufacturer's maximum specification limit. The measured relationship between thrust and bleed flow extraction was shown to be essentially linear at all power settings with an increase in bleed flow causing a corresponding decrease in thrust. A comparison with the F404-GE-400 steady-state engine simulation showed the estimation to be within +/- 1 percent of measured thrust losses for large increases in bleed flow rate.

Manual and automatic flight control during severe turbulence penetration

NASA Technical Reports Server (NTRS)

Johnston, D. E.; Klein, R. H.; Hoh, R. H.

1976-01-01

An analytical and experimental investigation of possible contributing factors in jet aircraft turbulence upsets was conducted. Major contributing factors identified included autopilot and display deficiencies, the large aircraft inertia and associated long response time, and excessive pilot workload. An integrated flight and thrust energy management director system was synthesized. The system was incorporated in a moving-base simulation and evaluated using highly experienced airline pilots. The evaluation included comparison of pilot workload and flight performance during severe turbulence penetration utilizing four control/display concepts: manual control with conventional full panel display, conventional autopilot (A/P-A) with conventional full panel display, improved autopilot (A/P-B) with conventional full panel display plus thrust director display, and longitudinal flight director with conventional full panel display plus thrust director display. Simulation results show improved performance, reduced pilot workload, and a pilot preference for the autopilot system controlling to the flight director command and manual control of thrust following the trim thrust director.

Rapid design and optimization of low-thrust rendezvous/interception trajectory for asteroid deflection missions

NASA Astrophysics Data System (ADS)

Li, Shuang; Zhu, Yongsheng; Wang, Yukai

2014-02-01

Asteroid deflection techniques are essential in order to protect the Earth from catastrophic impacts by hazardous asteroids. Rapid design and optimization of low-thrust rendezvous/interception trajectories is considered as one of the key technologies to successfully deflect potentially hazardous asteroids. In this paper, we address a general framework for the rapid design and optimization of low-thrust rendezvous/interception trajectories for future asteroid deflection missions. The design and optimization process includes three closely associated steps. Firstly, shape-based approaches and genetic algorithm (GA) are adopted to perform preliminary design, which provides a reasonable initial guess for subsequent accurate optimization. Secondly, Radau pseudospectral method is utilized to transcribe the low-thrust trajectory optimization problem into a discrete nonlinear programming (NLP) problem. Finally, sequential quadratic programming (SQP) is used to efficiently solve the nonlinear programming problem and obtain the optimal low-thrust rendezvous/interception trajectories. The rapid design and optimization algorithms developed in this paper are validated by three simulation cases with different performance indexes and boundary constraints.

Aeroservoelastic Modeling and Validation of a Thrust-Vectoring F/A-18 Aircraft

NASA Technical Reports Server (NTRS)

Brenner, Martin J.

1996-01-01

An F/A-18 aircraft was modified to perform flight research at high angles of attack (AOA) using thrust vectoring and advanced control law concepts for agility and performance enhancement and to provide a testbed for the computational fluid dynamics community. Aeroservoelastic (ASE) characteristics had changed considerably from the baseline F/A-18 aircraft because of structural and flight control system amendments, so analyses and flight tests were performed to verify structural stability at high AOA. Detailed actuator models that consider the physical, electrical, and mechanical elements of actuation and its installation on the airframe were employed in the analysis to accurately model the coupled dynamics of the airframe, actuators, and control surfaces. This report describes the ASE modeling procedure, ground test validation, flight test clearance, and test data analysis for the reconfigured F/A-18 aircraft. Multivariable ASE stability margins are calculated from flight data and compared to analytical margins. Because this thrust-vectoring configuration uses exhaust vanes to vector the thrust, the modeling issues are nearly identical for modem multi-axis nozzle configurations. This report correlates analysis results with flight test data and makes observations concerning the application of the linear predictions to thrust-vectoring and high-AOA flight.

Design of a mixer for the thrust-vectoring system on the high-alpha research vehicle

NASA Technical Reports Server (NTRS)

Pahle, Joseph W.; Bundick, W. Thomas; Yeager, Jessie C.; Beissner, Fred L., Jr.

1996-01-01

One of the advanced control concepts being investigated on the High-Alpha Research Vehicle (HARV) is multi-axis thrust vectoring using an experimental thrust-vectoring (TV) system consisting of three hydraulically actuated vanes per engine. A mixer is used to translate the pitch-, roll-, and yaw-TV commands into the appropriate TV-vane commands for distribution to the vane actuators. A computer-aided optimization process was developed to perform the inversion of the thrust-vectoring effectiveness data for use by the mixer in performing this command translation. Using this process a new mixer was designed for the HARV and evaluated in simulation and flight. An important element of the Mixer is the priority logic, which determines priority among the pitch-, roll-, and yaw-TV commands.

78 FR 6202 - Airworthiness Directives; the Boeing Company

Federal Register 2010, 2011, 2012, 2013, 2014

2013-01-30

... Descent (If INITIATE Required). Passenger Oxygen Switch.... ON Thrust Levers CLOSE Speed Brakes FLIGHT... Engineer, Systems and Equipment Branch, ANM-130S, FAA, Seattle Aircraft Certification Office, 1601 Lind... regard. Request To Correct AFM Reference to Target Speed Delta Airlines (DAL) asked that the target speed...

Multiphysics Thrust Chamber Modeling for Nuclear Thermal Propulsion

NASA Technical Reports Server (NTRS)

Wang, Ten-See; Cheng, Gary; Chen, Yen-Sen

2006-01-01

The objective of this effort is to develop an efficient and accurate thermo-fluid computational methodology to predict environments for a solid-core, nuclear thermal engine thrust chamber. The computational methodology is based on an unstructured-grid, pressure-based computational fluid dynamics formulation. A two-pronged approach is employed in this effort: A detailed thermo-fluid analysis on a multi-channel flow element for mid-section corrosion investigation; and a global modeling of the thrust chamber to understand the effect of heat transfer on thrust performance. Preliminary results on both aspects are presented.

Studies of Operating Frequency Effects On Ejector-based Thrust Augmentation in a Pulse Detonation Engine

NASA Technical Reports Server (NTRS)

Landry, K.

2005-01-01

Studies were performed in order to characterize the thrust augmentation potential of an ejector in a Pulse Detonation Engine application. A 49-mm diameter tube of 0.914-m length was constructed with one open end and one closed end. Ethylene, oxygen, and nitrogen were introduced into the tube at the closed end through the implementation of a fast mixing injector. The tube was completely filled with a stoichiometric mixture containing a one to one molar ratio of nitrogen to oxygen. Ethylene was selected as the fuel due to its detonation sensitivity and the molar ratio of the oxidizer was chosen for heat transfer purposes. Detonations were initiated in the tube through the use of a spark ignition system. The PDE was operated in a multi-cycle mode at frequencies ranging from 20-Hz to 50-Hz. Baseline thrust measurements with no ejector present were performed while operating the engine at various frequencies and compared to theoretical estimates. The baseline values were observed to agree with the theoretical model at low operating frequencies and proved to be increasingly lower than the predicted values as the operating frequency was increased. The baseline thrust measurements were observed to agree within 15 percent of the model for all operating frequencies. A straight 152-mm diameter ejector was installed and thrust augmentation percentages were measured. The length of the ejector was varied while the overlap percentage (percent of the ejector length which overlapped the tube) was maintained at 25 percent for all tests. In addition, the effect of ejector inlet geometry was investigated by comparing results with a straight inlet to those of a 38-mm inlet diameter. The thrust augmentation of the straight inlet ejector proved to be independent of engine operating frequency, augmenting thrust by 40 percent for the 0.914-m length ejector. In contrast, the rounded lip ejector of the same length seemed to be highly dependent on the engine operating frequency. An optimum operating frequency observed with the rounded inlet occurred at an operating frequency of 30-Hz, resulting in thrust augmentation percentages greater than 100 percent. The effect that the engine operating frequency had on thrust augmentation levels attained with an ejector was characterized and optimum performance parameters were established. Insight into the frequency dependent nature of the ejector performance was pursued. Suggestions for future experiments which are needed to fully understand the means in which thrust augmentation is achieved in a PDE-ejector configuration were noted.

A lifting surface theory for thrust augmenting ejectors

NASA Technical Reports Server (NTRS)

Bevilaqua, P. M.

1977-01-01

The circulation theory of airfoil lift has been applied to calculate the performance of thrust augmenting ejectors. The ejector shroud is considered to be 'flying' in the secondary velocity field induced by the entrainment of the primary jet, so that the augmenting thrust is viewed as analogous to the lift on an airfoil. Vortex lattice methods are utilized to compute the thrust augmentation from the force on the flaps. The augmentation is shown to be a function of the length and shape of the flaps, as well as their position and orientation. Predictions of this new theory are compared with the results of classical methods of calculating the augmentation by integration of the stream thrust.

A Computational Study of a New Dual Throat Fluidic Thrust Vectoring Nozzle Concept

NASA Technical Reports Server (NTRS)

Deere, Karen A.; Berrier, Bobby L.; Flamm, Jeffrey D.; Johnson, Stuart K.

2005-01-01

A computational investigation of a two-dimensional nozzle was completed to assess the use of fluidic injection to manipulate flow separation and cause thrust vectoring of the primary jet thrust. The nozzle was designed with a recessed cavity to enhance the throat shifting method of fluidic thrust vectoring. Several design cycles with the structured-grid, computational fluid dynamics code PAB3D and with experiments in the NASA Langley Research Center Jet Exit Test Facility have been completed to guide the nozzle design and analyze performance. This paper presents computational results on potential design improvements for best experimental configuration tested to date. Nozzle design variables included cavity divergence angle, cavity convergence angle and upstream throat height. Pulsed fluidic injection was also investigated for its ability to decrease mass flow requirements. Internal nozzle performance (wind-off conditions) and thrust vector angles were computed for several configurations over a range of nozzle pressure ratios from 2 to 7, with the fluidic injection flow rate equal to 3 percent of the primary flow rate. Computational results indicate that increasing cavity divergence angle beyond 10 is detrimental to thrust vectoring efficiency, while increasing cavity convergence angle from 20 to 30 improves thrust vectoring efficiency at nozzle pressure ratios greater than 2, albeit at the expense of discharge coefficient. Pulsed injection was no more efficient than steady injection for the Dual Throat Nozzle concept.

Internal performance of a nonaxisymmetric nozzle with a rotating upper flap and a center-pivoted lower flap

NASA Technical Reports Server (NTRS)

Wing, David J.; Leavitt, Laurence D.; Re, Richard J.

1993-01-01

An investigation was conducted at wind-off conditions in the static-test facility of the Langley 16-Foot Transonic Tunnel to determine the internal performance characteristics of a single expansion-ramp nozzle with thrust-vectoring capability to 105 degrees. Thrust vectoring was accomplished by the downward rotation of an upper flap with adaptive capability for internal contouring and a corresponding rotation of a center-pivoted lower flap. The static internal performance of configurations with pitch thrust-vector angles of 0 degrees, 60 degrees, and 105 degrees each with two throat areas, was investigated. The nozzle pressure ratio was varied from 1.5 to approximately 8.0 (5.0 for the maximum throat area configurations). Results of this study indicated that the nozzle configuration of the present investigation, when vectored, provided excellent flow-turning capability with relatively high levels of internal performance. In all cases, the thrust vector angle was a function of the nozzle pressure ratio. This result is expected because the flow is bounded by a single expansion surface on both vectored- and unvectored-nozzle geometries.

An Experimental Investigation of Unsteady Thrust Augmentation Using a Speaker-Driven Jet

NASA Technical Reports Server (NTRS)

Paxson, Daniel E.; Wernet, Mark P.; John, Wentworth T.

2004-01-01

An experimental investigation is described in which a simple speaker-driven jet was used as a pulsed thrust source (driver) for an ejector configuration. The objectives of the investigation were twofold: first, to add to the experimental body of evidence showing that an unsteady thrust source, combined with a properly sized ejector generally yields higher thrust augmentation values than a similarly sized, steady driver of equivalent thrust. Second, to identify characteristics of the unsteady driver that may be useful for sizing ejectors, and predicting what thrust augmentation values may be achieved. The speaker-driven jet provided a convenient source for the investigation because it is entirely unsteady (having no mean component) and because relevant parameters such as frequency, time-averaged thrust, and diameter are easily variable. The experimental setup will be described, as will the various measurements made. These include both thrust and Digital Particle Imaging Velocimetry of the driver. It will be shown that thrust augmentation values as high as 1.8 were obtained, that the diameter of the best ejector scaled with the dimensions of the emitted vortex, and that the so-called Formation Number serves as a useful dimensionless number by which to characterize the jet and predict performance.

Characterizing Longitude-Dependent Orbital Debris Congestion in the Geosynchronous Orbit Regime

NASA Astrophysics Data System (ADS)

Anderson, Paul V.

The geosynchronous orbit (GEO) is a unique commodity of the satellite industry that is becoming increasingly contaminated with orbital debris, but is heavily populated with high-value assets from the civil, commercial, and defense sectors. The GEO arena is home to hundreds of communications, data transmission, and intelligence satellites collectively insured for an estimated 18.3 billion USD. As the lack of natural cleansing mechanisms at the GEO altitude renders the lifetimes of GEO debris essentially infinite, conjunction and risk assessment must be performed to safeguard operational assets from debris collisions. In this thesis, longitude-dependent debris congestion is characterized by predicting the number of near-miss events per day for every longitude slot at GEO, using custom debris propagation tools and a torus intersection metric. Near-miss events with the present-day debris population are assigned risk levels based on GEO-relative position and speed, and this risk information is used to prioritize the population for debris removal target selection. Long-term projections of debris growth under nominal launch traffic, mitigation practices, and fragmentation events are also discussed, and latitudinal synchronization of the GEO debris population is explained via node variations arising from luni-solar gravity. In addition to characterizing localized debris congestion in the GEO ring, this thesis further investigates the conjunction risk to operational satellites or debris removal systems applying low-thrust propulsion to raise orbit altitude at end-of-life to a super-synchronous disposal orbit. Conjunction risks as a function of thrust level, miss distance, longitude, and semi-major axis are evaluated, and a guidance method for evading conjuncting debris with continuous thrust by means of a thrust heading change via single-shooting is developed.

Preliminary Design of Low-Thrust Interplanetary Missions

NASA Technical Reports Server (NTRS)

Sims, Jon A.; Flanagan, Steve N.

1997-01-01

For interplanetary missions, highly efficient electric propulsion systems can be used to increase the mass delivered to the destination and/or reduce the trip time over typical chemical propulsion systems. This technology is being demonstrated on the Deep Space 1 mission - part of NASA's New Millennium Program validating technologies which can lower the cost and risk and enhance the performance of future missions. With the successful demonstration on Deep Space 1, future missions can consider electric propulsion as a viable propulsion option. Electric propulsion systems, while highly efficient, produce only a small amount of thrust. As a result, the engines operate during a significant fraction of the trajectory. This characteristic makes it much more difficult to find optimal trajectories. The methods for optimizing low-thrust trajectories are typically categorized as either indirect, or direct. Indirect methods are based on calculus of variations, resulting in a two-point boundary value problem that is solved by satisfying terminal constraints and targeting conditions. These methods are subject to extreme sensitivity to the initial guess of the variables - some of which are not physically intuitive. Adding a gravity assist to the trajectory compounds the sensitivity. Direct methods parameterize the problem and use nonlinear programming techniques to optimize an objective function by adjusting a set of variables. A variety of methods of this type have been examined with varying results. These methods are subject to the limitations of the nonlinear programming techniques. In this paper we present a direct method intended to be used primarily for preliminary design of low-thrust interplanetary trajectories, including those with multiple gravity assists. Preliminary design implies a willingness to accept limited accuracy to achieve an efficient algorithm that executes quickly.

Investigation of beamed-energy ERH thruster performance

NASA Technical Reports Server (NTRS)

Myrabo, Leik N.; Strayer, T. Darton; Bossard, John A.; Richard, Jacques C.; Gallimore, Alec D.

1986-01-01

The objective of this study was to determine the performance of an External Radiation Heated (ERH) thruster. In this thruster, high intensity laser energy is focused to ignite either a Laser Supported Combustion (LSC) wave or a Laser Supported Detonation (LSD) wave. Thrust is generated as the LSC or LSD wave propagates over the thruster's surface, or in the proposed thruster configuration, the vehicle afterbody. Thrust models for the LSC and LSD waves were developed and simulated on a computer. Performance parameters investigated include the effect of laser intensity, flight Mach number, and altitude on mean-thrust and coupling coefficient of the ERH thruster. Results from these models suggest that the ERH thruster using LSC/LSD wave ignition could provide propulsion performance considerably greater than any propulsion system currently available.

Method and system for monitoring and displaying engine performance parameters

NASA Technical Reports Server (NTRS)

Abbott, Terence S. (Inventor); Person, Lee H., Jr. (Inventor)

1988-01-01

The invention is believed a major improvement that will have a broad application in governmental and commercial aviation. It provides a dynamic method and system for monitoring and simultaneously displaying in easily scanned form the available, predicted, and actual thrust of a jet aircraft engine under actual operating conditions. The available and predicted thrusts are based on the performance of a functional model of the aircraft engine under the same operating conditions. Other critical performance parameters of the aircraft engine and functional model are generated and compared, the differences in value being simultaneously displayed in conjunction with the displayed thrust values. Thus, the displayed information permits the pilot to make power adjustments directly while keeping him aware of total performance at a glance of a single display panel.

Method and system for monitoring and displaying engine performance parameters

NASA Technical Reports Server (NTRS)

Abbott, Terence S. (Inventor); Person, Jr., Lee H. (Inventor)

1991-01-01

The invention is a method and system for monitoring and directly displaying the actual thrust produced by a jet aircraft engine under determined operating conditions and the available thrust and predicted (commanded) thrust of a functional model of an ideal engine under the same determined operating conditions. A first set of actual value output signals representative of a plurality of actual performance parameters of the engine under the determined operating conditions is generated and compared with a second set of predicted value output signals representative of the predicted value of corresponding performance parameters of a functional model of the engine under the determined operating conditions to produce a third set of difference value output signals within a range of normal, caution, or warning limit values. A thrust indicator displays when any one of the actual value output signals is in the warning range while shaping function means shape each of the respective difference output signals as each approaches the limit of the respective normal, caution, and warning range limits.

Analysis of gas turbine engines using water and oxygen injection to achieve high Mach numbers and high thrust

NASA Technical Reports Server (NTRS)

Henneberry, Hugh M.; Snyder, Christopher A.

1993-01-01

An analysis of gas turbine engines using water and oxygen injection to enhance performance by increasing Mach number capability and by increasing thrust is described. The liquids are injected, either separately or together, into the subsonic diffuser ahead of the engine compressor. A turbojet engine and a mixed-flow turbofan engine (MFTF) are examined, and in pursuit of maximum thrust, both engines are fitted with afterburners. The results indicate that water injection alone can extend the performance envelope of both engine types by one and one-half Mach numbers at which point water-air ratios reach 17 or 18 percent and liquid specific impulse is reduced to some 390 to 470 seconds, a level about equal to the impulse of a high energy rocket engine. The envelope can be further extended, but only with increasing sacrifices in liquid specific impulse. Oxygen-airflow ratios as high as 15 percent were investigated for increasing thrust. Using 15 percent oxygen in combination with water injection at high supersonic Mach numbers resulted in thrust augmentation as high as 76 percent without any significant decrease in liquid specific impulse. The stoichiometric afterburner exit temperature increased with increasing oxygen flow, reaching 4822 deg R in the turbojet engine at a Mach number of 3.5. At the transonic Mach number of 0.95 where no water injection is needed, an oxygen-air ratio of 15 percent increased thrust by some 55 percent in both engines, along with a decrease in liquid specific impulse of 62 percent. Afterburner temperature was approximately 4700 deg R at this high thrust condition. Water and/or oxygen injection are simple and straightforward strategies to improve engine performance and they will add little to engine weight. However, if large Mach number and thrust increases are required, liquid flows become significant, so that operation at these conditions will necessarily be of short duration.

Results From a Parametric Acoustic Liner Experiment Using P and W GEN1 HSR Mixer/Ejector Model

NASA Technical Reports Server (NTRS)

Boyd, Kathleen C.; Wolter, John D.

2004-01-01

This report documents the results of an acoustic liner test performed using a Gen 1 HSR mixer/ejector model installed on the Jet Exit Rig in the Nozzle Acoustic Test Rig in the Aeroacoustic Propulsion Laboratory or NASA Glenn Research Center. Acoustic liner effectiveness and single-component thrust performance results are discussed. Results from 26 different types of single-degree-of-freedom and bulk material liners are compared with each other and against a hardwall baseline. Design parameters involving all aspects of the facesheet, the backing cavity, and the type of bulk material were varied in order to study the effects of these design features on the acoustic impedance, acoustic effectiveness and on nozzle thrust performance. Overall, the bulk absorber liners are more effective at reducing the jet noise than the single-degree-of-freedom liners. Many of the design parameters had little effect on acoustic effectiveness, such as facesheeet hole diameter and honeycomb cell size. A relatively large variation in the impedance of the bulk absorber in a bulk liner is required to have a significant impact on the noise reduction. The thrust results exhibit a number of consistent trends, supporting the validity of this new addition to the facility. In general, the thrust results indicate that thrust performance benefits from increased facesheet thickness and decreased facesheet porosity.

Dynamics of high-bypass-engine thrust reversal using a variable-pitch fan

NASA Technical Reports Server (NTRS)

Schaefer, J. W.; Sagerser, D. R.; Stakolich, E. G.

1977-01-01

The test program demonstrated that successful and rapid forward-to reverse-thrust transients can be performed without any significant engine operational limitations for fan blade pitch changes through either feather pitch or flat pitch. For through-feather-pitch operation with a flight inlet, fan stall problems were encountered, and a fan blade overshoot technique was used to establish reverse thrust.

Performance Capability of Single-Cavity Vortex Gaseous Nuclear Rockets

NASA Technical Reports Server (NTRS)

Ragsdale, Robert G.

1963-01-01

An analysis was made to determine the maximum powerplant thrust-to-weight ratio possible with a single-cavity vortex gaseous reactor in which all the hydrogen propellant must diffuse through a fuel-rich region. An assumed radial temperature profile was used to represent conduction, convection, and radiation heat-transfer effects. The effect of hydrogen property changes due to dissociation and ionization was taken into account in a hydrodynamic computer program. It is shown that, even for extremely optimistic assumptions of reactor criticality and operating conditions, such a system is limited to reactor thrust-to-weight ratios of about 1.2 x 10(exp -3) for laminar flow. For turbulent flow, the maximum thrust-to-weight ratio is less than 10(exp -3). These low thrusts result from the fact that the hydrogen flow rate is limited by the diffusion process. The performance of a gas-core system with a specific impulse of 3000 seconds and a powerplant thrust-to-weight ratio of 10(exp -2) is shown to be equivalent to that of a 1000-second advanced solid-core system. It is therefore concluded that a single-cavity vortex gaseous reactor in which all the hydrogen must diffuse through the nuclear fuel is a low-thrust device and offers no improvement over a solid-core nuclear-rocket engine. To achieve higher thrust, additional hydrogen flow must be introduced in such a manner that it will by-pass the nuclear fuel. Obviously, such flow must be heated by thermal radiation. An illustrative model of a single-cavity vortex system employing supplementary flow of hydrogen through the core region is briefly examined. Such a system appears capable of thrust-to-weight ratios of approximately 1 to 10. For a high-impulse engine, this capability would be a considerable improvement over solid-core performance. Limits imposed by thermal radiation heat transfer to cavity walls are acknowledged but not evaluated. Alternate vortex concepts that employ many parallel vortices to achieve higher hydrogen flow rates offer the possibility of sufficiently high thrust-to-weight ratios, if they are not limited by short thermal-radiation path lengths.

Global Optimization of Low-Thrust Interplanetary Trajectories Subject to Operational Constraints

NASA Technical Reports Server (NTRS)

Englander, Jacob A.; Vavrina, Matthew A.; Hinckley, David

2016-01-01

Low-thrust interplanetary space missions are highly complex and there can be many locally optimal solutions. While several techniques exist to search for globally optimal solutions to low-thrust trajectory design problems, they are typically limited to unconstrained trajectories. The operational design community in turn has largely avoided using such techniques and has primarily focused on accurate constrained local optimization combined with grid searches and intuitive design processes at the expense of efficient exploration of the global design space. This work is an attempt to bridge the gap between the global optimization and operational design communities by presenting a mathematical framework for global optimization of low-thrust trajectories subject to complex constraints including the targeting of planetary landing sites, a solar range constraint to simplify the thermal design of the spacecraft, and a real-world multi-thruster electric propulsion system that must switch thrusters on and off as available power changes over the course of a mission.

Maximum sustained fin-kick thrust in underwater swimming.

PubMed

Yamaguchi, H; Shidara, F; Naraki, N; Mohri, M

1995-09-01

We examined the upper limit of a diver's fin-kick thrust force using a stationary-swimming ergometer. Heart rate, respiratory minute volume, oxygen uptake, and performance rate were measured in four male subjects who swam constantly for 8 min to maintain a horizontal position against an applied force at a depth of 0.7 m. The water temperature was controlled at 26 degrees +/- 1 degree C. The performance rate, which was the parameter of how well the subjects compensated for the applied load, showed an upper limit around 64 N of sustainable thrust force. This meant that the diver could generate the swimming thrust force within 64 N continuously for 8 min in a steady state. Heart rate, respiratory minute volume, and O2 uptake showed almost proportional increases to the applied load within 64 N and tended to plateau about 69 N.

Starting Vortex Identified as Key to Unsteady Ejector Performance

NASA Technical Reports Server (NTRS)

Paxson, Daniel E.

2004-01-01

Unsteady ejectors are currently under investigation for use in some pulse-detonation-engine-based propulsion systems. Experimental measurements made in the past, and recently at the NASA Glenn Research Center, have demonstrated that thrust augmentation can be enhanced considerably when the driver is unsteady. In ejector systems, thrust augmentation is defined as = T(sup Total)/T(sup j), where T(sup Total) is the total thrust of the combined ejector and driving jet and T(sup j) is the thrust due to the driving jet alone. There are three images in this figure, one for each of the named thrust sources. The images are color contours of measured instantaneous vorticity. Each image is an ensemble average of at least 150 phase-locked measurements. The flow is from right to left, and the shape and location of each driver is shown on the far right of each image. The emitted vortex is a clearly defined "doughnut" of highly vortical (spinning) flow. In these planar images, the vortex appears as two distorted circles, one above, and one below the axis of symmetry. Because they are spinning in the opposite direction, the two circles have vorticity of opposite sign and thus are different colors. There is also a rectangle shown in each image. Its width represents the ejector diameter that was found experimentally to yield the highest thrust augmentation. It is apparent that the optimal ejector diameter is that which just "captures" the vortex: that is, the diameter bounding the outermost edge of the vortex structure. The exact mechanism behind the enhanced performance is unclear; however, it is believed to be related to the powerful vortex emitted with each pulse of the unsteady driver. As such, particle imaging velocimetry (PIV) measurements were obtained for three unsteady drivers: a pulsejet, a resonance tube, and a speaker-driven jet. All the drivers were tested with ejectors, and all exhibited performance enhancement over similarly sized steady drivers. The characteristic starting vortices of each driver are shown in these images. The images are color contours of measured instantaneous vorticity. Each image is an ensemble average of at least 150 phase-locked measurements. The flow is from right to left. The shape and location of each driver is shown on the far right of each image. The rectangle shown in each image represents the ejector diameter that was found experimentally to yield the highest thrust augmentation. It is apparent that the optimal ejector diameter is that which just "captures" the vortex: that is, the diameter bounding the outermost edge of the vortex structure. Although not shown, it was observed that the emitted vortex spread as it traveled downstream. The spreading rate for the pulsejet is shown as the dashed lines in the top image. A tapered ejector was fabricated that matched this shape. When tested, the ejector demonstrated superior performance to all those previously tested at Glenn (which were essentially of straight, cylindrical form), achieving a remarkable thrust augmentation of 2. The measured thrust augmentation is shown as a function of ejector length. Also shown are the thrust augmentation values achieved with the straight, cylindrical ejectors of varying diameters. Here, thrust augmentation is plotted as a function of ejector length for several families of ejector diameters. It can be seen that large thrust augmentation values are indeed obtained and that they are sensitive to both ejector length and diameter, particularly the latter. Five curves are shown. Four correspond to straight ejector diameters of 2.2, 3.0, 4.0, and 6.0 in. The fifth curve corresponds to the tapered ejector contoured to bound the emitted vortex. For each curve, there are several data points corresponding to different lengths. The largest value of thrust augmentation is 2.0 for the tapered ejector and 1.81 for the straight ejectors. Regardless of their diameters, all the ejectors trend toward peak performance at a particular leng. That the cross-sectional dimensions of optimal ejectors scaled precisely with the vortex dimensions on three separate pulsed thrust sources demonstrates that the action of the vortex is responsible for the enhanced ejector performance. The result also suggests that, in the absence of a complete understanding of the entrainment and augmentation mechanisms, methods of characterizing starting vortices may be useful for correlating and predicting unsteady ejector performance.

Thrust Evaluation of an Arcjet Thruster Using Dimethyl Ether as a Propellant

NASA Astrophysics Data System (ADS)

Kakami, Akira; Beppu, Shinji; Maiguma, Muneyuki; Tachibana, Takeshi

This paper describes the performance of an arcjet thruster using dimethyl ether (DME) as a propellant. DME, an ether compound, has adequate characteristics for space propulsion systems; DME is storable in a liquid state without a high pressure or cryogenic device and requires no sophisticated temperature management. DME is gasified and liquefied simply by adjusting temperature, whereas hydrazine, a conventional propellant, requires an iridium-based particulate catalyst for its gasification. In this study, thrust of the designed kW-class DME arcjet thruster is measured with a torsional thrust stand. Thrust measurements show that thrust is increased with propellant mass flow rate, and that thrust using DME propellant is higher than when using nitrogen. The prototype DME arcjet thruster yields a specific impulse of 330 s, a thruster efficiency of 0.14, and a thrust of 0.19 N at 60-mg/s DME mass flow rate at 25-A discharge current. The corresponding discharge power and specific power are 2.3 kW and 39 MJ/kg.

Thrust Measurement of Dielectric Barrier Discharge (DBD) Plasma Actuators: New Anti-Thrust Hypothesis, Frequency Sweeps Methodology, Humidity and Enclosure Effects

NASA Technical Reports Server (NTRS)

Ashpis, David E.; Laun, Matthew C.

2014-01-01

We discuss thrust measurements of Dielectric Barrier Discharge (DBD) plasma actuators devices used for aerodynamic active flow control. After a review of our experience with conventional thrust measurement and significant non-repeatability of the results, we devised a suspended actuator test setup, and now present a methodology of thrust measurements with decreased uncertainty. The methodology consists of frequency scans at constant voltages. The procedure consists of increasing the frequency in a step-wise fashion from several Hz to the maximum frequency of several kHz, followed by frequency decrease back down to the start frequency of several Hz. This sequence is performed first at the highest voltage of interest, then repeated at lower voltages. The data in the descending frequency direction is more consistent and selected for reporting. Sample results show strong dependence of thrust on humidity which also affects the consistency and fluctuations of the measurements. We also observed negative values of thrust, or "anti-thrust", at low frequencies between 4 Hz and up to 64 Hz. The anti-thrust is proportional to the mean-squared voltage and is frequency independent. Departures from the parabolic anti-thrust curve are correlated with appearance of visible plasma discharges. We propose the anti-thrust hypothesis. It states that the measured thrust is a sum of plasma thrust and anti-thrust, and assumes that the anti-thrust exists at all frequencies and voltages. The anti-thrust depends on actuator geometry and materials and on the test installation. It enables the separation of the plasma thrust from the measured total thrust. This approach enables more meaningful comparisons between actuators at different installations and laboratories. The dependence on test installation was validated by surrounding the actuator with a grounded large-diameter metal sleeve. Strong dependence on humidity is also shown; the thrust significantly increased with decreasing humidity, e.g., 44 percent increase as relative humidity changed from 18 percent and dew point 33 degF to 50 percent and dew point of 57 degF.

Preliminary performance of a vertical-attitude takeoff and landing, supersonic cruise aircraft concept having thrust vectoring integrated into the flight control system

NASA Technical Reports Server (NTRS)

Robins, A. W.; Beissner, F. L., Jr.; Domack, C. S.; Swanson, E. E.

1985-01-01

A performance study was made of a vertical attitude takeoff and landing (VATOL), supersonic cruise aircraft concept having thrust vectoring integrated into the flight control system. Those characteristics considered were aerodynamics, weight, balance, and performance. Preliminary results indicate that high levels of supersonic aerodynamic performance can be achieved. Further, with the assumption of an advanced (1985 technology readiness) low bypass ratio turbofan engine and advanced structures, excellent mission performance capability is indicated.

A model for teaching and learning spinal thrust manipulation and its effect on participant confidence in technique performance.

PubMed

Wise, Christopher H; Schenk, Ronald J; Lattanzi, Jill Black

2016-07-01

Despite emerging evidence to support the use of high velocity thrust manipulation in the management of lumbar spinal conditions, utilization of thrust manipulation among clinicians remains relatively low. One reason for the underutilization of these procedures may be related to disparity in training in the performance of these techniques at the professional and post professional levels. To assess the effect of using a new model of active learning on participant confidence in the performance of spinal thrust manipulation and the implications for its use in the professional and post-professional training of physical therapists. A cohort of 15 DPT students in their final semester of entry-level professional training participated in an active training session emphasizing a sequential partial task practice (SPTP) strategy in which participants engaged in partial task practice over several repetitions with different partners. Participants' level of confidence in the performance of these techniques was determined through comparison of pre- and post-training session surveys and a post-session open-ended interview. The increase in scores across all items of the individual pre- and post-session surveys suggests that this model was effective in changing overall participant perception regarding the effectiveness and safety of these techniques and in increasing student confidence in their performance. Interviews revealed that participants greatly preferred the SPTP strategy, which enhanced their confidence in technique performance. Results indicate that this new model of psychomotor training may be effective at improving confidence in the performance of spinal thrust manipulation and, subsequently, may be useful for encouraging the future use of these techniques in the care of individuals with impairments of the spine. Inasmuch, this method of instruction may be useful for training of physical therapists at both the professional and post-professional levels.

Static Investigation of a Multiaxis Thrust-Vectoring Nozzle With Variable Internal Contouring Ability

NASA Technical Reports Server (NTRS)

Wing, David J.; Mills, Charles T. L.; Mason, Mary L.

1997-01-01

The thrust efficiency and vectoring performance of a convergent-divergent nozzle were investigated at static conditions in the model preparation area of the Langley 16-Foot Transonic Tunnel. The diamond-shaped nozzle was capable of varying the internal contour of each quadrant individually by using cam mechanisms and retractable drawers to produce pitch and yaw thrust vectoring. Pitch thrust vectoring was achieved by either retracting the lower drawers to incline the throat or varying the internal flow-path contours to incline the throat. Yaw thrust vectoring was achieved by reducing flow area left of the nozzle centerline and increasing flow area right of the nozzle centerline; a skewed throat deflected the flow in the lateral direction.

Airflow and thrust calibration of an F100 engine, S/N P680059, at selected flight conditions

NASA Technical Reports Server (NTRS)

Biesiadny, T. J.; Lee, D.; Rodriguez, J. R.

1978-01-01

An airflow and thrust calibration of an F100 engine, S/N P680059, was conducted to study airframe propulsion system integration losses in turbofan-powered high-performance aircraft. The tests were conducted with and without thrust augmentation for a variety of simulated flight conditions with emphasis on the transonic regime. The resulting corrected airflow data generalized into one curve with corrected fan speed while corrected gross thrust increased as simulated flight conditions increased. Overall agreement between measured data and computed results was 1 percent for corrected airflow and -1 1/2 percent for gross thrust. The results of an uncertainty analysis are presented for both parameters at each simulated flight condition.

Tank 12 data dump OME integrated thrust chamber test report, phase 1

NASA Technical Reports Server (NTRS)

Pauckert, R. P.; Tobin, R. D.

1974-01-01

The test program conducted to characterize the steady state stability, thermal, and performance characteristics of the integrated thrust chamber assembly, as well as limited tests to investigate transient characteristics are described.

Computational Investigation of Fluidic Counterflow Thrust Vectoring

NASA Technical Reports Server (NTRS)

Hunter, Craig A.; Deere, Karen A.

1999-01-01

A computational study of fluidic counterflow thrust vectoring has been conducted. Two-dimensional numerical simulations were run using the computational fluid dynamics code PAB3D with two-equation turbulence closure and linear Reynolds stress modeling. For validation, computational results were compared to experimental data obtained at the NASA Langley Jet Exit Test Facility. In general, computational results were in good agreement with experimental performance data, indicating that efficient thrust vectoring can be obtained with low secondary flow requirements (less than 1% of the primary flow). An examination of the computational flowfield has revealed new details about the generation of a countercurrent shear layer, its relation to secondary suction, and its role in thrust vectoring. In addition to providing new information about the physics of counterflow thrust vectoring, this work appears to be the first documented attempt to simulate the counterflow thrust vectoring problem using computational fluid dynamics.

Low-Thrust Many-Revolution Trajectory Optimization via Differential Dynamic Programming and a Sundman Transformation

NASA Technical Reports Server (NTRS)

Aziz, Jonathan D.; Parker, Jeffrey S.; Scheeres, Daniel J.; Englander, Jacob A.

2017-01-01

Low-thrust trajectories about planetary bodies characteristically span a high count of orbital revolutions. Directing the thrust vector over many revolutions presents a challenging optimization problem for any conventional strategy. This paper demonstrates the tractability of low-thrust trajectory optimization about planetary bodies by applying a Sundman transformation to change the independent variable of the spacecraft equations of motion to the eccentric anomaly and performing the optimization with differential dynamic programming. Fuel-optimal geocentric transfers are shown in excess of 1000 revolutions while subject to Earths J2 perturbation and lunar gravity.

Performance (Off-Design) Cycle Analysis for a Turbofan Engine With Interstage Turbine Burner

NASA Technical Reports Server (NTRS)

Liew, K. H.; Urip, E.; Yang, S. L.; Mattingly, J. D.; Marek, C. J.

2005-01-01

This report presents the performance of a steady-state, dual-spool, separate-exhaust turbofan engine, with an interstage turbine burner (ITB) serving as a secondary combustor. The ITB, which is located in the transition duct between the high- and the low-pressure turbines, is a relatively new concept for increasing specific thrust and lowering pollutant emissions in modern jet-engine propulsion. A detailed off-design performance analysis of ITB engines is written in Microsoft(Registered Trademark) Excel (Redmond, Washington) macrocode with Visual Basic Application to calculate engine performances over the entire operating envelope. Several design-point engine cases are pre-selected using a parametric cycle-analysis code developed previously in Microsoft(Registered Trademark) Excel, for off-design analysis. The off-design code calculates engine performances (i.e. thrust and thrust-specific-fuel-consumption) at various flight conditions and throttle settings.

Primary propulsion/large space system interactions

NASA Technical Reports Server (NTRS)

Dergance, R. H.

1980-01-01

Three generic types of structural concepts and nonstructural surface densities were selected and combined to represent potential LSS applications. The design characteristics of various classes of large space systems that are impacted by primary propulsion thrust required to effect orbit transfer were identified. The effects of propulsion system thrust-to-mass ratio, thrust transients, and performance on the mass, area, and orbit transfer characteristics of large space systems were determined.

Magnetohydrodynamic Augmentation of Pulse Detonation Rocket Engines (Preprint)

DTIC Science & Technology

2010-09-28

augmentation of the thrust . Ejectors typically transfer energy between streams through shear stress between separate flow streams, where a portion of the...the opportunity to extract energy and apply it to a separate stream where the net thrust can be increased. With MHD augmentation , such as in the Pulse...with the PDRIME for separate or additional thrust augmentation . Results show potential performance gains under many flight and operating conditions

Performance of a Rotary Wing Air Brake in Supersonic Flow

DTIC Science & Technology

1947-05-18

a pneumatic thrust measuring system (3) Strobotac and Audio Signal Generator for RPM measurements, (4) motion picture cameras and miscellaneous...produced by the head to keep the brake engaged. The " pneumatic thrust measuring system1* utilised the bellows described above to oppose the drag force...a brake , to prevent its rotation when desired. To release the brake it was necessary to inflate the internal thrust bellows, thereby disengaging

Static investigation of two STOL nozzle concepts with pitch thrust-vectoring capability

NASA Technical Reports Server (NTRS)

Mason, M. L.; Burley, J. R., II

1986-01-01

A static investigation of the internal performance of two short take-off and landing (STOL) nozzle concepts with pitch thrust-vectoring capability has been conducted. An axisymmetric nozzle concept and a nonaxisymmetric nozzle concept were tested at dry and afterburning power settings. The axisymmetric concept consisted of a circular approach duct with a convergent-divergent nozzle. Pitch thrust vectoring was accomplished by vectoring the approach duct without changing the nozzle geometry. The nonaxisymmetric concept consisted of a two dimensional convergent-divergent nozzle. Pitch thrust vectoring was implemented by blocking the nozzle exit and deflecting a door in the lower nozzle flap. The test nozzle pressure ratio was varied up to 10.0, depending on model geometry. Results indicate that both pitch vectoring concepts produced resultant pitch vector angles which were nearly equal to the geometric pitch deflection angles. The axisymmetric nozzle concept had only small thrust losses at the largest pitch deflection angle of 70 deg., but the two-dimensional convergent-divergent nozzle concept had large performance losses at both of the two pitch deflection angles tested, 60 deg. and 70 deg.

Magnet pole shape design for reduction of thrust ripple of slotless permanent magnet linear synchronous motor with arc-shaped magnets considering end-effect based on analytical method

NASA Astrophysics Data System (ADS)

Shin, Kyung-Hun; Park, Hyung-Il; Kim, Kwan-Ho; Jang, Seok-Myeong; Choi, Jang-Young

2017-05-01

The shape of the magnet is essential to the performance of a slotless permanent magnet linear synchronous machine (PMLSM) because it is directly related to desirable machine performance. This paper presents a reduction in the thrust ripple of a PMLSM through the use of arc-shaped magnets based on electromagnetic field theory. The magnetic field solutions were obtained by considering end effect using a magnetic vector potential and two-dimensional Cartesian coordinate system. The analytical solution of each subdomain (PM, air-gap, coil, and end region) is derived, and the field solution is obtained by applying the boundary and interface conditions between the subdomains. In particular, an analytical method was derived for the instantaneous thrust and thrust ripple reduction of a PMLSM with arc-shaped magnets. In order to demonstrate the validity of the analytical results, the back electromotive force results of a finite element analysis and experiment on the manufactured prototype model were compared. The optimal point for thrust ripple minimization is suggested.

Development Status of High-Thrust Density Electrostatic Engines

NASA Technical Reports Server (NTRS)

Patterson, Michael J.; Haag, Thomas W.; Foster, John E.; Young, Jason A.; Crofton, Mark W.

2017-01-01

Ion thruster technology offers the highest performance and efficiency of any mature electric propulsion thruster. It has by far the highest demonstrated total impulse of any technology option, demonstrated at input power levels appropriate for primary propulsion. It has also been successfully implemented for primary propulsion in both geocentric and heliocentric environments, with excellent ground/in-space correlation of both its performance and life. Based on these attributes there is compelling reasoning to continue the development of this technology: it is a leading candidate for high power applications; and it provides risk reduction for as-yet unproven alternatives. As such it is important that the operational limitations of ion thruster technology be critically examined and in particular for its application to primary propulsion its capabilities relative to thrust the density and thrust-to-power ratio be understood. This publication briefly addresses some of the considerations relative to achieving high thrust density and maximizing thrust-to-power ratio with ion thruster technology, and discusses the status of development work in this area being executed under a collaborative effort among NASA Glenn Research Center, the Aerospace Corporation, and the University of Michigan.

Engine performance analysis and optimization of a dual-mode scramjet with varied inlet conditions

NASA Astrophysics Data System (ADS)

Tian, Lu; Chen, Li-Hong; Chen, Qiang; Zhong, Feng-Quan; Chang, Xin-Yu

2016-02-01

A dual-mode scramjet can operate in a wide range of flight conditions. Higher thrust can be generated by adopting suitable combustion modes. Based on the net thrust, an analysis and preliminary optimal design of a kerosene-fueled parameterized dual-mode scramjet at a crucial flight Mach number of 6 were investigated by using a modified quasi-one-dimensional method and simulated annealing strategy. Engine structure and heat release distributions, affecting the engine thrust, were chosen as analytical parameters for varied inlet conditions (isolator entrance Mach number: 1.5-3.5). Results show that different optimal heat release distributions and structural conditions can be obtained at five different inlet conditions. The highest net thrust of the parameterized dual-mode engine can be achieved by a subsonic combustion mode at an isolator entrance Mach number of 2.5. Additionally, the effects of heat release and scramjet structure on net thrust have been discussed. The present results and the developed analytical method can provide guidance for the design and optimization of high-performance dual-mode scramjets.

Characterization of advanced electric propulsion systems

NASA Technical Reports Server (NTRS)

Ray, P. K.

1982-01-01

Characteristics of several advanced electric propulsion systems are evaluated and compared. The propulsion systems studied are mass driver, rail gun, MPD thruster, hydrogen free radical thruster and mercury electron bombardment ion engine. These are characterized by specific impulse, overall efficiency, input power, average thrust, power to average thrust ratio and average thrust to dry weight ratio. Several important physical characteristics such as dry system mass, accelerator length, bore size and current pulse requirement are also evaluated in appropriate cases. Only the ion engine can operate at a specific impulse beyond 2000 sec. Rail gun, MPD thruster and free radical thruster are currently characterized by low efficiencies. Mass drivers have the best performance characteristics in terms of overall efficiency, power to average thrust ratio and average thrust to dry weight ratio. But, they can only operate at low specific impulses due to large power requirements and are extremely long due to limitations of driving current. Mercury ion engines have the next best performance characteristics while operating at higher specific impulses. It is concluded that, overall, ion engines have somewhat better characteristics as compared to the other electric propulsion systems.

Thrust Stand for Electric Propulsion Performance Evaluation

NASA Technical Reports Server (NTRS)

Markusic, T. E.; Jones, J. E.; Cox, M. D.

2004-01-01

An electric propulsion thrust stand capable of supporting thrusters with total mass of up to 125 kg and 1 mN to 1 N thrust levels has been developed and tested. The mechanical design features a conventional hanging pendulum arm attached to a balance mechanism that transforms horizontal motion into amplified vertical motion, with accommodation for variable displacement sensitivity. Unlike conventional hanging pendulum thrust stands, the deflection is independent of the length of the pendulum arm, and no reference structure is required at the end of the pendulum. Displacement is measured using a non-contact, optical linear gap displacement transducer. Mechanical oscillations are attenuated using a passive, eddy current damper. An on-board microprocessor-based level control system, which includes a two axis accelerometer and two linear-displacement stepper motors, continuously maintains the level of the balance mechanism - counteracting mechanical %era drift during thruster testing. A thermal control system, which includes heat exchange panels, thermocouples, and a programmable recirculating water chiller, continuously adjusts to varying thermal loads to maintain the balance mechanism temperature, to counteract thermal drifts. An in-situ calibration rig allows for steady state calibration both prior to and during thruster testing. Thrust measurements were carried out on a well-characterized 1 kW Hall thruster; the thrust stand was shown to produce repeatable results consistent with previously published performance data.

Hydrodynamic thrust generation and power consumption investigations for piezoelectric fins with different aspect ratios

NASA Astrophysics Data System (ADS)

Shahab, S.; Tan, D.; Erturk, A.

2015-12-01

Bio-inspired hydrodynamic thrust generation using piezoelectric transduction has recently been explored using Macro-Fiber Composite (MFC) actuators. The MFC technology strikes a balance between the actuation force and structural deformation levels for effective swimming performance, and additionally offers geometric scalability, silent operation, and ease of fabrication. Recently we have shown that mean thrust levels comparable to biological fish of similar size can be achieved using MFC fins. The present work investigates the effect of length-to-width (L/b) aspect ratio on the hydrodynamic thrust generation performance of MFC cantilever fins by accounting for the power consumption level. It is known that the hydrodynamic inertia and drag coefficients are controlled by the aspect ratio especially for L/b< 5. The three MFC bimorph fins explored in this work have the aspect ratios of 2.1, 3.9, and 5.4. A nonlinear electrohydroelastic model is employed to extract the inertia and drag coefficients from the vibration response to harmonic actuation for the first bending mode. Experiments are then conducted for various actuation voltage levels to quantify the mean thrust resultant and power consumption levels for different aspect ratios. Variation of the thrust coefficient of the MFC bimorph fins with changing aspect ratio is also semi-empirically modeled and presented.

Aeroassisted orbit transfer vehicle trajectory analysis

NASA Technical Reports Server (NTRS)

Braun, Robert D.; Suit, William T.

1988-01-01

The emphasis in this study was on the use of multiple pass trajectories for aerobraking. However, for comparison, single pass trajectories, trajectories using ballutes, and trajectories corrupted by atmospheric anomolies were run. A two-pass trajectory was chosen to determine the relation between sensitivity to errors and payload to orbit. Trajectories that used only aerodynamic forces for maneuvering could put more weight into the target orbits but were very sensitive to variations from the planned trajectors. Using some thrust control resulted in less payload to orbit, but greatly reduced the sensitivity to variations from nominal trajectories. When compared to the non-thrusting trajectories investigated, the judicious use of thrusting resulted in multiple pass trajectories that gave 97 percent of the payload to orbit with almost none of the sensitivity to variations from the nominal.

Oxygen/Alcohol Dual Thrust RCS Engines

NASA Technical Reports Server (NTRS)

Angstadt, Tara; Hurlbert, Eric

1999-01-01

A non-toxic dual thrust RCS engine offers significant operational, safety, and performance advantages to the space shuttle and the next generation RLVs. In this concept, a single engine produces two thrust levels of 25 and 870 lbf. The low thrust level is provided by the spark torch igniter, which, with the addition of 2 extra valves, can also be made to function as a vernier. A dual thrust RCS engine allows 38 verniers to be packaged more efficiently on a vehicle. These 38 vemiers improve translation and reduce cross coupling, thereby providing more pure roll, pitch, and yaw maneuvers of the vehicle. Compared to the 6 vemiers currently on the shuttle, the 38 dual thrust engines would be 25 to 40% more efficient for the same maneuvers and attitude control. The vernier thrust level also reduces plume impingement and contamination concerns. Redundancy is also improved, thereby improving mission success reliability. Oxygen and ethanol are benign propellants which do not create explosive reaction products or contamination, as compared to hypergolic propellants. These characteristics make dual-thrust engines simpler to implement on a non-toxic reaction control system. Tests at WSTF in August 1999 demonstrated a dual-thrust concept that is successful with oxygen and ethanol. Over a variety of inlet pressures and mixture ratios at 22:1 area ratio, the engine produced between 230 and 297 sec Isp, and thrust levels from 8 lbf. to 50 lbf. This paper describes the benefits of dual-thrust engines and the recent results from tests at WSTF.

Static investigation of a two-dimensional convergent-divergent exhaust nozzle with multiaxis thrust-vectoring capability

NASA Technical Reports Server (NTRS)

Taylor, John G.

1990-01-01

An investigation was conducted in the Static Test Facility of the NASA Langley 16-Foot Transonic Tunnel to determine the internal performance of two-dimensional convergent-divergent nozzles designed to have simultaneous pitch and yaw thrust vectoring capability. This concept utilized divergent flap rotation of thrust vectoring in the pitch plane and deflection of flat yaw flaps hinged at the end of the sidewalls for yaw thrust vectoring. The hinge location of the yaw flaps was varied at four positions from the nozzle exit plane to the throat plane. The yaw flaps were designed to contain the flow laterally independent of power setting. In order to eliminate any physical interference between the yaw flap deflected into the exhaust stream and the divergent flaps, the downstream corners of both upper and lower divergent flaps were cut off to allow for up to 30 deg of yaw flap deflection. The impact of varying the nozzle pitch vector angle, throat area, yaw flap hinge location, yaw flap length, and yaw flap deflection angle on nozzle internal performance characteristics, was studied. High-pressure air was used to simulate jet exhaust at nozzle pressure ratios up to 7.0. Static results indicate that configurations with the yaw flap hinge located upstream of the exit plane provide relatively high levels of thrust vectoring efficiency without causing large losses in resultant thrust ratio. Therefore, these configurations represent a viable concept for providing simultaneous pitch and yaw thrust vectoring.

Experimental Investigation of Unsteady Thrust Augmentation Using a Speaker-Driven Jet

NASA Technical Reports Server (NTRS)

Paxson, Daniel E.; Wernet, Mark P.; John, Wentworth T.

2007-01-01

An experimental investigation is described in which a simple speaker-driven jet was used as a pulsed thrust source (driver) for an ejector configuration. The objectives of the investigation were twofold. The first was to expand the experimental body of evidence showing that an unsteady thrust source, combined with a properly sized ejector generally yields higher thrust augmentation values than a similarly sized, steady driver of equivalent thrust. The second objective was to identify characteristics of the unsteady driver that may be useful for sizing ejectors, and for predicting the thrust augmentation levels that may be achieved. The speaker-driven jet provided a convenient source for the investigation because it is entirely unsteady (i.e., it has no mean velocity component) and because relevant parameters such as frequency, time-averaged thrust, and diameter are easily variable. The experimental setup will be described, as will the two main measurements techniques employed. These are thrust and digital particle imaging velocimetry of the driver. It will be shown that thrust augmentation values as high as 1.8 were obtained, that the diameter of the best ejector scaled with the dimensions of the emitted vortex, and that the so-called formation time serves as a useful dimensionless parameter by which to characterize the jet and predict performance.

Effects of Cavity on the Performance of Dual Throat Nozzle During the Thrust-Vectoring Starting Transient Process.

PubMed

Gu, Rui; Xu, Jinglei

2014-01-01

The dual throat nozzle (DTN) technique is capable to achieve higher thrust-vectoring efficiencies than other fluidic techniques, without compromising thrust efficiency significantly during vectoring operation. The excellent performance of the DTN is mainly due to the concaved cavity. In this paper, two DTNs of different scales have been investigated by unsteady numerical simulations to compare the parameter variations and study the effects of cavity during the vector starting process. The results remind us that during the vector starting process, dynamic loads may be generated, which is a potentially challenging problem for the aircraft trim and control.

Performance of a Low-Power Cylindrical Hall Thruster

NASA Technical Reports Server (NTRS)

Polzin, Kurt A.; Markusic, Thomas E.; Stanojev, Boris J.; Dehoyos, Amado; Raitses, Yevgeny; Smirnov, Artem; Fisch, Nathaniel J.

2007-01-01

Recent mission studies have shown that a Hall thruster which operates at relatively constant thrust efficiency (45-55%) over a broad power range (300W - 3kW) is enabling for deep space science missions when compared with slate-of-the-art ion thrusters. While conventional (annular) Hall thrusters can operate at high thrust efficiency at kW power levels, it is difficult to construct one that operates over a broad power envelope down to 0 (100 W) while maintaining relatively high efficiency. In this note we report the measured performance (I(sub sp), thrust and efficiency) of a cylindrical Hall thruster operating at 0 (100 W) input power.

Advanced engine study for mixed-mode orbit-transfer vehicles

NASA Technical Reports Server (NTRS)

Mellish, J. A.

1978-01-01

Engine design, performance, weight and envelope data were established for three mixed-mode orbit-transfer vehicle engine candidates. Engine concepts evaluated are the tripropellant, dual-expander and plug cluster. Oxygen, RP-1 and hydrogen are the propellants considered for use in these engines. Theoretical performance and propellant properties were established for bipropellant and tripropellant mixes of these propellants. RP-1, hydrogen and oxygen were evaluated as coolants and the maximum attainable chamber pressures were determined for each engine concept within the constraints of the propellant properties and the low cycle thermal fatigue (300 cycles) requirement. The baseline engine design and component operating characteristics are determined at a thrust level of 88,964N (20,000 lbs) and a thrust split of 0.5. The parametric data is generated over ranges of thrust and thrust split of 66.7 to 400kN (15 to 90 klb) and 0.4 to 0.8, respectively.

Full scale hover test of a 25 foot tilt rotor

NASA Technical Reports Server (NTRS)

Helf, S.; Broman, E.; Gatchel, S.; Charles, B.

1973-01-01

The tilt rotor underwent a hover performance test on the Aero Propulsion Laboratory whirl stand at Wright-Patterson Air Force Base. The maximum thrust over density ratio measured at the design tip speed of 740 feet per second was 10,016 pounds. This occurred when the power over density ratio was 1721 horsepower. At the hover overspeed rpm, the thrust and power, over density ratio, were 11,008 pounds and 1866 horsepower. During the test, the maximum measured thrust coefficient was 0.177, and the rotor figure of merit exceeded 0.81. Measured lifting efficiency was 8.35 pounds per horsepower at the thrust a 13,000-pound aircraft would require for hover at sea level on a standard day. No effect of compressibility on performance is discernible in the test results (the range of tip Mach numbers tested was 0.55 to 0.71).

A Double-Sided Linear Primary Permanent Magnet Vernier Machine

PubMed Central

2015-01-01

The purpose of this paper is to present a new double-sided linear primary permanent magnet (PM) vernier (DSLPPMV) machine, which can offer high thrust force, low detent force, and improved power factor. Both PMs and windings of the proposed machine are on the short translator, while the long stator is designed as a double-sided simple iron core with salient teeth so that it is very robust to transmit high thrust force. The key of this new machine is the introduction of double stator and the elimination of translator yoke, so that the inductance and the volume of the machine can be reduced. Hence, the proposed machine offers improved power factor and thrust force density. The electromagnetic performances of the proposed machine are analyzed including flux, no-load EMF, thrust force density, and inductance. Based on using the finite element analysis, the characteristics and performances of the proposed machine are assessed. PMID:25874250

A double-sided linear primary permanent magnet vernier machine.

PubMed

Du, Yi; Zou, Chunhua; Liu, Xianxing

2015-01-01

The purpose of this paper is to present a new double-sided linear primary permanent magnet (PM) vernier (DSLPPMV) machine, which can offer high thrust force, low detent force, and improved power factor. Both PMs and windings of the proposed machine are on the short translator, while the long stator is designed as a double-sided simple iron core with salient teeth so that it is very robust to transmit high thrust force. The key of this new machine is the introduction of double stator and the elimination of translator yoke, so that the inductance and the volume of the machine can be reduced. Hence, the proposed machine offers improved power factor and thrust force density. The electromagnetic performances of the proposed machine are analyzed including flux, no-load EMF, thrust force density, and inductance. Based on using the finite element analysis, the characteristics and performances of the proposed machine are assessed.

Froghopper-inspired direction-changing concept for miniature jumping robots.

PubMed

Jung, Gwang-Pil; Cho, Kyu-Jin

2016-09-14

To improve the maneuverability and agility of jumping robots, several researchers have studied steerable jumping mechanisms. This steering ability enables robots to reach a particular target by controlling their jumping direction. To this end, we propose a novel direction-changing concept for miniature jumping robots. The proposed concept allows robots to be steerable while exerting minimal effects on jumping performance. The key design principles were adopted from the froghopper's power-producing hind legs and the moment cancellation accomplished by synchronized leg operation. These principles were applied via a pair of symmetrically positioned legs and conventional gears, which were modeled on the froghopper's anatomy. Each leg has its own thrusting energy, which improves jumping performance by allowing the mechanism to thrust itself with both power-producing legs. Conventional gears were utilized to simultaneously operate the legs and cancel out the moments that they induce, which minimizes body spin. A prototype to verify the concept was built and tested by varying the initial jumping posture. Three jumping postures (synchronous, asynchronous, and single-legged) were tested to investigate how synchronization and moment cancelling affect jumping performance. The results show that synchronous jumping allows the mechanism to change direction from -40° to 40°, with an improved take-off speed. The proposed concept can only be steered in a limited range of directions, but it has potential for use in miniature jumping robots that can change jumping direction with a minimal drop in jumping performance.

Static investigation of two fluidic thrust-vectoring concepts on a two-dimensional convergent-divergent nozzle

NASA Technical Reports Server (NTRS)

Wing, David J.

1994-01-01

A static investigation was conducted in the static test facility of the Langley 16-Foot Transonic Tunnel of two thrust-vectoring concepts which utilize fluidic mechanisms for deflecting the jet of a two-dimensional convergent-divergent nozzle. One concept involved using the Coanda effect to turn a sheet of injected secondary air along a curved sidewall flap and, through entrainment, draw the primary jet in the same direction to produce yaw thrust vectoring. The other concept involved deflecting the primary jet to produce pitch thrust vectoring by injecting secondary air through a transverse slot in the divergent flap, creating an oblique shock in the divergent channel. Utilizing the Coanda effect to produce yaw thrust vectoring was largely unsuccessful. Small vector angles were produced at low primary nozzle pressure ratios, probably because the momentum of the primary jet was low. Significant pitch thrust vector angles were produced by injecting secondary flow through a slot in the divergent flap. Thrust vector angle decreased with increasing nozzle pressure ratio but moderate levels were maintained at the highest nozzle pressure ratio tested. Thrust performance generally increased at low nozzle pressure ratios and decreased near the design pressure ratio with the addition of secondary flow.

Multi-Objective Hybrid Optimal Control for Multiple-Flyby Interplanetary Mission Design using Chemical Propulsion

NASA Technical Reports Server (NTRS)

Englander, Jacob A.; Vavrina, Matthew A.

2015-01-01

Preliminary design of high-thrust interplanetary missions is a highly complex process. The mission designer must choose discrete parameters such as the number of flybys and the bodies at which those flybys are performed. For some missions, such as surveys of small bodies, the mission designer also contributes to target selection. In addition, real-valued decision variables, such as launch epoch, flight times, maneuver and flyby epochs, and flyby altitudes must be chosen. There are often many thousands of possible trajectories to be evaluated. The customer who commissions a trajectory design is not usually interested in a point solution, but rather the exploration of the trade space of trajectories between several different objective functions. This can be a very expensive process in terms of the number of human analyst hours required. An automated approach is therefore very desirable. This work presents such an approach by posing the impulsive mission design problem as a multi-objective hybrid optimal control problem. The method is demonstrated on several real-world problems. Two assumptions are frequently made to simplify the modeling of an interplanetary high-thrust trajectory during the preliminary design phase. The first assumption is that because the available thrust is high, any maneuvers performed by the spacecraft can be modeled as discrete changes in velocity. This assumption removes the need to integrate the equations of motion governing the motion of a spacecraft under thrust and allows the change in velocity to be modeled as an impulse and the expenditure of propellant to be modeled using the time-independent solution to Tsiolkovsky's rocket equation [1]. The second assumption is that the spacecraft moves primarily under the influence of the central body, i.e. the sun, and all other perturbing forces may be neglected in preliminary design. The path of the spacecraft may then be modeled as a series of conic sections. When a spacecraft performs a close approach to a planet, the central body switches from the sun to that planet and the trajectory is modeled as a hyperbola with respect to the planet. This is known as the method of patched conics. The impulsive and patched-conic assumptions significantly simplify the preliminary design problem.

Analysis of an all-metallic resilient-pad gas-lubricated thrust bearing

NASA Technical Reports Server (NTRS)

Anderson, W. J.

1974-01-01

A resilient-pad gas thrust bearing that does not contain any elastomers in the bearing assembly is described and analyzed. The bearing consists of sector-shaped pads mounted asymmetrically on resilient foil beams. The effects of bearing design parameters on performance are shown. Performance of a resilient-pad bearing is compared with that of a pivoted-pad bearing.

A Monte Carlo Analysis of the Thrust Imbalance for the RSRMV Booster During Both the Ignition Transient and Steady State Operation

NASA Technical Reports Server (NTRS)

Foster, Winfred A., Jr.; Crowder, Winston; Steadman, Todd E.

2014-01-01

This paper presents the results of statistical analyses performed to predict the thrust imbalance between two solid rocket motor boosters to be used on the Space Launch System (SLS) vehicle. Two legacy internal ballistics codes developed for the Space Shuttle program were coupled with a Monte Carlo analysis code to determine a thrust imbalance envelope for the SLS vehicle based on the performance of 1000 motor pairs. Thirty three variables which could impact the performance of the motors during the ignition transient and thirty eight variables which could impact the performance of the motors during steady state operation of the motor were identified and treated as statistical variables for the analyses. The effects of motor to motor variation as well as variations between motors of a single pair were included in the analyses. The statistical variations of the variables were defined based on data provided by NASA's Marshall Space Flight Center for the upgraded five segment booster and from the Space Shuttle booster when appropriate. The results obtained for the statistical envelope are compared with the design specification thrust imbalance limits for the SLS launch vehicle

A Monte Carlo Analysis of the Thrust Imbalance for the Space Launch System Booster During Both the Ignition Transient and Steady State Operation

NASA Technical Reports Server (NTRS)

Foster, Winfred A., Jr.; Crowder, Winston; Steadman, Todd E.

2014-01-01

This paper presents the results of statistical analyses performed to predict the thrust imbalance between two solid rocket motor boosters to be used on the Space Launch System (SLS) vehicle. Two legacy internal ballistics codes developed for the Space Shuttle program were coupled with a Monte Carlo analysis code to determine a thrust imbalance envelope for the SLS vehicle based on the performance of 1000 motor pairs. Thirty three variables which could impact the performance of the motors during the ignition transient and thirty eight variables which could impact the performance of the motors during steady state operation of the motor were identified and treated as statistical variables for the analyses. The effects of motor to motor variation as well as variations between motors of a single pair were included in the analyses. The statistical variations of the variables were defined based on data provided by NASA's Marshall Space Flight Center for the upgraded five segment booster and from the Space Shuttle booster when appropriate. The results obtained for the statistical envelope are compared with the design specification thrust imbalance limits for the SLS launch vehicle.

Predicted Performance of a Thrust-Enhanced SR-71 Aircraft with an External Payload

NASA Technical Reports Server (NTRS)

Conners, Timothy R.

1997-01-01

NASA Dryden Flight Research Center has completed a preliminary performance analysis of the SR-71 aircraft for use as a launch platform for high-speed research vehicles and for carrying captive experimental packages to high altitude and Mach number conditions. Externally mounted research platforms can significantly increase drag, limiting test time and, in extreme cases, prohibiting penetration through the high-drag, transonic flight regime. To provide supplemental SR-71 acceleration, methods have been developed that could increase the thrust of the J58 turbojet engines. These methods include temperature and speed increases and augmentor nitrous oxide injection. The thrust-enhanced engines would allow the SR-71 aircraft to carry higher drag research platforms than it could without enhancement. This paper presents predicted SR-71 performance with and without enhanced engines. A modified climb-dive technique is shown to reduce fuel consumption when flying through the transonic flight regime with a large external payload. Estimates are included of the maximum platform drag profiles with which the aircraft could still complete a high-speed research mission. In this case, enhancement was found to increase the SR-71 payload drag capability by 25 percent. The thrust enhancement techniques and performance prediction methodology are described.

Experimental investigation of the effects of variable expanding channel on the performance of a low-power cusped field thruster

NASA Astrophysics Data System (ADS)

Liu, Hui; Zeng, Ming; Jiang, Wenjia; Yang, Chiyu; Ning, Zhongxi; Yu, Daren

2018-04-01

Due to a special magnetic field structure, the multi-cusped field thruster shows advantages of low wall erosion, low noise and high thrust density over a wide range of thrust. In this paper, expanding discharge channels are employed to make up for deficiencies on the range of thrust and plume divergence, which often emerges in conventional straight cylindrical channels. Three thruster geometries are fabricated with different expanding-angle channels, and a group of experiments are carried out to find out their influence on the performance and discharge characteristics of the thruster. A retarding potential analyzer and a Faraday probe are employed to analyze the structures of the plume in these three models. The results show that when the thrusters operate at low mass flow rate, the gradually-expanding channels exhibit lower propellant utilization and lower overall performance by amounts not exceeding 44.8% in ionization rate and 19.5% in anode efficiency, respectively. But the weakening of magnetic field intensity near the exit of expanding channels leads to an extended thrust throttling ability, a smaller plume divergence angle, and a relatively larger stable operating space without mode converting and the consequent performance degradation.

Propulsive performance of biologically inspired flapping foils at high Reynolds numbers.

PubMed

Techet, Alexandra H

2008-01-01

Propulsion and maneuvering underwater by flapping foil motion, optimized through years of evolution, is ubiquitous in nature, yet marine propulsors inspired by examples of highly maneuverable marine life or aquatic birds are not widely implemented in engineering. Performance data from flapping foils, moving in a rolling and pitching motion, are presented at high Reynolds numbers, Re=Uc/nu, or O(10(4)), where U is the relative inflow velocity, c is the chord length of the foil, and nu is the kinematic viscosity of the fluid, from water tunnel experiments using a foil actuator module designed after an aquatic penguin or turtle fin. The average thrust coefficients and efficiency measurements are recorded over a range of kinematic flapping amplitudes and frequencies. Results reveal a maximum thrust coefficient of 2.09, and for low values of angle of attack the thrust generally increases with Strouhal number, without much penalty to efficiency. Strouhal number is defined as St=2h(0)f/U, where f is the frequency of flapping, and 2h(0) is the peak-to-peak amplitude of flapping. The thrust and efficiency contour plots also present a useful performance trend where, at low angles of attack, high thrust and efficiency can be gained at sufficiently high Strouhal numbers. Understanding the motion of aquatic penguins and turtle wings and emulating these motions mechanically can yield insight into the hydrodynamics of how these animals swim and also improve performance of biologically inspired propulsive devices.

NASA Double Asteroid Redirection Test (DART) Trajectory Validation and Robutness

NASA Technical Reports Server (NTRS)

Sarli, Bruno V.; Ozimek, Martin T.; Atchison, Justin A.; Englander, Jacob A.; Barbee, Brent W.

2017-01-01

The Double Asteroid Redirection Test (DART) mission will be the first to test the concept of a kinetic impactor. Several studies have been made on asteroid redirection and impact mitigation, however, to this date no mission tested the proposed concepts. An impact study on a representative body allows the measurement of the effects on the target's orbit and physical structure. With this goal, DART's objective is to verify the effectiveness of the kinetic impact concept for planetary defense. The spacecraft uses solar electric propulsion to escape Earth, fly by (138971) 2001 CB21 for impact rehearsal, and impact Didymos-B, the secondary body of the binary (65803) Didymos system. This work focuses on the heliocentric transfer design part of the mission with the validation of the baseline trajectory, performance comparison to other mission objectives, and assessment of the baseline robustness to missed thrust events. Results show a good performance of the selected trajectory for different mission objectives: latest possible escape date, maximum kinetic energy on impact, shortest possible time of flight, and use of an Earth swing-by. The baseline trajectory was shown to be robust to a missed thrust with 1% of fuel margin being enough to recover the mission for failures of more than 14 days.

Structural dynamics and aerodynamics measurements of biologically inspired flexible flapping wings.

PubMed

Wu, P; Stanford, B K; Sällström, E; Ukeiley, L; Ifju, P G

2011-03-01

Flapping wing flight as seen in hummingbirds and insects poses an interesting unsteady aerodynamic problem: coupling of wing kinematics, structural dynamics and aerodynamics. There have been numerous studies on the kinematics and aerodynamics in both experimental and computational cases with both natural and artificial wings. These studies tend to ignore wing flexibility; however, observation in nature affirms that passive wing deformation is predominant and may be crucial to the aerodynamic performance. This paper presents a multidisciplinary experimental endeavor in correlating a flapping micro air vehicle wing's aeroelasticity and thrust production, by quantifying and comparing overall thrust, structural deformation and airflow of six pairs of hummingbird-shaped membrane wings of different properties. The results show that for a specific spatial distribution of flexibility, there is an effective frequency range in thrust production. The wing deformation at the thrust-productive frequencies indicates the importance of flexibility: both bending and twisting motion can interact with aerodynamic loads to enhance wing performance under certain conditions, such as the deformation phase and amplitude. By measuring structural deformations under the same aerodynamic conditions, beneficial effects of passive wing deformation can be observed from the visualized airflow and averaged thrust. The measurements and their presentation enable observation and understanding of the required structural properties for a thrust effective flapping wing. The intended passive responses of the different wings follow a particular pattern in correlation to their aerodynamic performance. Consequently, both the experimental technique and data analysis method can lead to further studies to determine the design principles for micro air vehicle flapping wings.

Study of Jet-Propulsion System Comprising Blower, Burner, and Nozzle

NASA Technical Reports Server (NTRS)

Hall, Eldon W

1944-01-01

A study was made of the performance of a jet-propulsion system composed of an engine-driven blower, a combustion chamber, and a discharge nozzle. A simplified analysis is made of this system for the purpose of showing in concise form the effect of the important design variables and operating conditions on jet thrust, thrust horsepower, and fuel consumption. Curves are presented that permit a rapid evaluation of the performance of this system for a range of operating conditions. The performance for an illustrative case of a power plant of the type under consideration id discussed in detail. It is shown that for a given airplane velocity the jet thrust horsepower depends mainly on the blower power and the amount of fuel burned in the jet; the higher the thrust horsepower is for a given blower power, the higher the fuel consumption per thrust horsepower. Within limits the amount of air pumped has only a secondary effect on the thrust horsepower and efficiency. A lower limit on air flow for a given fuel flow occurs where the combustion-chamber temperature becomes excessive on the basis of the strength of the structure. As the air-flow rate is increased, an upper limit is reached where, for a given blower power, fuel-flow rate, and combustion-chamber size, further increase in air flow causes a decrease in power and efficiency. This decrease in power is caused by excessive velocity through the combustion chamber, attended by an excessive pressure drop caused by momentum changes occurring during combustion.

Navigation and Guidance for Low-Thrust Trajectories, LOTNAV

NASA Astrophysics Data System (ADS)

Cano, J. L.; Bello, M.; Rodriguez-Canabal, J.

A number of interplanetary low-thrust missions have already been flown by many space agencies. Examples of already flown missions based on the use of electric propulsion are Deep Space 1, Hayabusa and SMART-1. Many others are already in the assessment phase or in the development phase itself. In such perspective, it is required by the space agencies the procurement and utilisation of assessment tools for fast prototyping in the areas of mission design and navigation. The Low-Thrust Interplanetary Navigation Tool, which is the subject of this paper, allows the mission analyst performing such type of quick assessment studies for the early phases in the development of low-thrust missions. A number of test cases on low-thrust missions are also presented along with the utilities composing the LOTNAV tool.

Comparative jet wake structure and swimming performance of salps.

PubMed

Sutherland, Kelly R; Madin, Laurence P

2010-09-01

Salps are barrel-shaped marine invertebrates that swim by jet propulsion. Morphological variations among species and life-cycle stages are accompanied by differences in swimming mode. The goal of this investigation was to compare propulsive jet wakes and swimming performance variables among morphologically distinct salp species (Pegea confoederata, Weelia (Salpa) cylindrica, Cyclosalpa sp.) and relate swimming patterns to ecological function. Using a combination of in situ dye visualization and particle image velocimetry (PIV) measurements, we describe properties of the jet wake and swimming performance variables including thrust, drag and propulsive efficiency. Locomotion by all species investigated was achieved via vortex ring propulsion. The slow-swimming P. confoederata produced the highest weight-specific thrust (T=53 N kg(-1)) and swam with the highest whole-cycle propulsive efficiency (eta(wc)=55%). The fast-swimming W. cylindrica had the most streamlined body shape but produced an intermediate weight-specific thrust (T=30 N kg(-1)) and swam with an intermediate whole-cycle propulsive efficiency (eta(wc)=52%). Weak swimming performance variables in the slow-swimming C. affinis, including the lowest weight-specific thrust (T=25 N kg(-1)) and lowest whole-cycle propulsive efficiency (eta(wc)=47%), may be compensated by low energetic requirements. Swimming performance variables are considered in the context of ecological roles and evolutionary relationships.

Static performance investigation of a skewed-throat multiaxis thrust-vectoring nozzle concept

NASA Technical Reports Server (NTRS)

Wing, David J.

1994-01-01

The static performance of a jet exhaust nozzle which achieves multiaxis thrust vectoring by physically skewing the geometric throat has been characterized in the static test facility of the 16-Foot Transonic Tunnel at NASA Langley Research Center. The nozzle has an asymmetric internal geometry defined by four surfaces: a convergent-divergent upper surface with its ridge perpendicular to the nozzle centerline, a convergent-divergent lower surface with its ridge skewed relative to the nozzle centerline, an outwardly deflected sidewall, and a straight sidewall. The primary goal of the concept is to provide efficient yaw thrust vectoring by forcing the sonic plane (nozzle throat) to form at a yaw angle defined by the skewed ridge of the lower surface contour. A secondary goal is to provide multiaxis thrust vectoring by combining the skewed-throat yaw-vectoring concept with upper and lower pitch flap deflections. The geometric parameters varied in this investigation included lower surface ridge skew angle, nozzle expansion ratio (divergence angle), aspect ratio, pitch flap deflection angle, and sidewall deflection angle. Nozzle pressure ratio was varied from 2 to a high of 11.5 for some configurations. The results of the investigation indicate that efficient, substantial multiaxis thrust vectoring was achieved by the skewed-throat nozzle concept. However, certain control surface deflections destabilized the internal flow field, which resulted in substantial shifts in the position and orientation of the sonic plane and had an adverse effect on thrust-vectoring and weight flow characteristics. By increasing the expansion ratio, the location of the sonic plane was stabilized. The asymmetric design resulted in interdependent pitch and yaw thrust vectoring as well as nonzero thrust-vector angles with undeflected control surfaces. By skewing the ridges of both the upper and lower surface contours, the interdependency between pitch and yaw thrust vectoring may be eliminated and the location of the sonic plane may be further stabilized.

Internal performance of a 10 deg conical plug nozzle with a multispoke primary and translating external shroud

NASA Technical Reports Server (NTRS)

Bresnahan, D. L.

1972-01-01

An experimental investigation was conducted in a nozzle static test facility to determine the performance characteristics of a cold-flow, 21.59-centimeter-diameter plug nozzle with a multispoke primary. Two multispoke primary nozzles, a 12-spoke and a 24-spoke, were tested and compared with an annular plug nozzle. The supersonic cruise configurations for both spoke primaries performed about the same, with a gross thrust coefficient of 0.974, a decrease of approximately 1.5 percent from the reference nozzle. The takeoff configuration for the 12-spoke primary had a gross thrust coefficient of 0.957, a decrease of 1.5 percent from the reference nozzle, and the 24-spoke primary had a gross thrust coefficient of 0.95.

Astrium Preparation of Future Cryogenic Thrust Chamber Development

NASA Astrophysics Data System (ADS)

Nicolay, Rolf

2002-01-01

The scenarios for future cryogenic propulsion in Europe are mainly governed by cost issues on the one side and performance issues on the other. Certain relationships of the different issues exist to either the application for ELVs or RLVs respectively. Taking into account the limited budgets of the Europeans Agency Market, flexible development strategies are and have to be defined to fulfill both applications requirements. Investigations aiming at identifying the optimum development strategy serving both applications have been performed. Based on the experience of the different cryogenic thrust chamber developments already performed, Astrium worked out a flexible development strategy for future cryogenic thrust chambers in order to: This paper is going to report about this development strategy and the associated derived needs for technological investigations and development work.

Internal Performance of Several Divergent-Shroud Ejector Nozzles with High Divergence Angles

NASA Technical Reports Server (NTRS)

Trout, Arthur M.; Papell, S. Stephen; Povolny, John H.

1957-01-01

Nine divergent-shroud ejector configurations were investigated to determine the effect of shroud divergence angle on ejector internal performance. Unheated dry air was used for both the primary and secondary flows. The decrease in the design-point thrust coefficient with increasing flow divergence angle (angle measured from primary exit to shroud exit) followed very closely a simple relation involving the cosine of the angle. This indicates that design-point thrust performance for divergent-shroud ejectors can be predicted with reasonable accuracy within the range investigated. The decrease in design-point thrust coefficient due to increasing the flow divergence engle from 120deg to 30deg (half-singles) was approximately 6 percent. Ejector air-handling characteristics and the primary-nozzle flow coefficient were not significantly affected by change in shroud divergence angle.

Modes of thrust generation in flying animals

NASA Astrophysics Data System (ADS)

Luo, Haoxiang; Song, Jialei; Tobalske, Bret; Luo Team; Tobalske Team

2016-11-01

For flying animals in forward flight, thrust is usually much smaller as compared with weight support and has not been given the same amount of attention. Several modes of thrust generation are discussed in this presentation. For insects performing slow flight that is characterized by low advance ratios (i.e., the ratio between flight speed and wing speed), thrust is usually generated by a "backward flick" mode, in which the wings moves upward and backward at a faster speed than the flight speed. Paddling mode is another mode used by some insects like fruit flies who row their wings backward during upstroke like paddles (Ristroph et al., PRL, 2011). Birds wings have high advance ratios and produce thrust during downstroke by directing aerodynamic lift forward. At intermediate advance ratios around one (e.g., hummingbirds and bats), the animal wings generate thrust during both downstroke and upstroke, and thrust generation during upstroke may come at cost of negative weight support. These conclusions are supported by previous experiment studies of insects, birds, and bats, as well as our recent computational modeling of hummingbirds. Supported by the NSF.

Nuclear Thermal Rocket Simulation in NPSS

NASA Technical Reports Server (NTRS)

Belair, Michael L.; Sarmiento, Charles J.; Lavelle, Thomas M.

2013-01-01

Four nuclear thermal rocket (NTR) models have been created in the Numerical Propulsion System Simulation (NPSS) framework. The models are divided into two categories. One set is based upon the ZrC-graphite composite fuel element and tie tube-style reactor developed during the Nuclear Engine for Rocket Vehicle Application (NERVA) project in the late 1960s and early 1970s. The other reactor set is based upon a W-UO2 ceramic-metallic (CERMET) fuel element. Within each category, a small and a large thrust engine are modeled. The small engine models utilize RL-10 turbomachinery performance maps and have a thrust of approximately 33.4 kN (7,500 lbf ). The large engine models utilize scaled RL-60 turbomachinery performance maps and have a thrust of approximately 111.2 kN (25,000 lbf ). Power deposition profiles for each reactor were obtained from a detailed Monte Carlo N-Particle (MCNP5) model of the reactor cores. Performance factors such as thermodynamic state points, thrust, specific impulse, reactor power level, and maximum fuel temperature are analyzed for each engine design.

Nuclear Thermal Rocket Simulation in NPSS

NASA Technical Reports Server (NTRS)

Belair, Michael L.; Sarmiento, Charles J.; Lavelle, Thomas L.

2013-01-01

Four nuclear thermal rocket (NTR) models have been created in the Numerical Propulsion System Simulation (NPSS) framework. The models are divided into two categories. One set is based upon the ZrC-graphite composite fuel element and tie tube-style reactor developed during the Nuclear Engine for Rocket Vehicle Application (NERVA) project in the late 1960s and early 1970s. The other reactor set is based upon a W-UO2 ceramic- metallic (CERMET) fuel element. Within each category, a small and a large thrust engine are modeled. The small engine models utilize RL-10 turbomachinery performance maps and have a thrust of approximately 33.4 kN (7,500 lbf ). The large engine models utilize scaled RL-60 turbomachinery performance maps and have a thrust of approximately 111.2 kN (25,000 lbf ). Power deposition profiles for each reactor were obtained from a detailed Monte Carlo N-Particle (MCNP5) model of the reactor cores. Performance factors such as thermodynamic state points, thrust, specific impulse, reactor power level, and maximum fuel temperature are analyzed for each engine design.

Thrust shock vector control of an axisymmetric conical supersonic nozzle via secondary transverse gas injection

NASA Astrophysics Data System (ADS)

Zmijanovic, V.; Lago, V.; Sellam, M.; Chpoun, A.

2014-01-01

Transverse secondary gas injection into the supersonic flow of an axisymmetric convergent-divergent nozzle is investigated to describe the effects of the fluidic thrust vectoring within the framework of a small satellite launcher. Cold-flow dry-air experiments are performed in a supersonic wind tunnel using two identical supersonic conical nozzles with the different transverse injection port positions. The complex three-dimensional flow field generated by the supersonic cross-flows in these test nozzles was examined. Valuable experimental data were confronted and compared with the results obtained from the numerical simulations. Different nozzle models are numerically simulated under experimental conditions and then further investigated to determine which parameters significantly affect thrust vectoring. Effects which characterize the nozzle and thrust vectoring performances are established. The results indicate that with moderate secondary to primary mass flow rate ratios, ranging around 5 %, it is possible to achieve pertinent vector side forces. It is also revealed that injector positioning and geometry have a strong effect on the shock vector control system and nozzle performances.

Variable Speed CMG Control of a Dual-Spin Stabilized Unconventional VTOL Air Vehicle

NASA Technical Reports Server (NTRS)

Lim, Kyong B.; Moerder, Daniel D.; Shin, J-Y.

2004-01-01

This paper describes an approach based on using both bias momentum and multiple control moment gyros for controlling the attitude of statically unstable thrust-levitated vehicles in hover or slow translation. The stabilization approach described in this paper uses these internal angular momentum transfer devices for stability, augmented by thrust vectoring for trim and other outer loop control functions, including CMG stabilization/ desaturation under persistent external disturbances. Simulation results show the feasibility of (1) improved vehicle performance beyond bias momentum assisted vector thrusting control, and (2) using control moment gyros to significantly reduce the external torque required from the vector thrusting machinery.

Computation of optimal Mars trajectories via combined chemical/electrical propulsion, Part 3: Compromise solutions

NASA Astrophysics Data System (ADS)

Miele, A.; Wang, T.; Williams, P. N.

2005-12-01

The success of the solar-electric ion engine powering the DS1 spacecraft has paved the way toward the use of low-thrust electrical engines in future planetary/interplanetary missions. Vis-à-vis a chemical engine, an electrical engine has a higher specific impulse, implying a possible decrease in propellant mass; however, the low-thrust aspect discourages the use of an electrical engine in the near-planet phases of a trip, since this might result in an increase in flight time. Therefore, a fundamental design problem is to find the best combination of chemical propulsion and electrical propulsion for a given mission, for example, a mission from Earth to Mars. With this in mind, this paper is the third of a series dealing with the optimization of Earth Mars missions via the use of hybrid engines, namely the combination of high-thrust chemical engines for planetary flight and low-thrust electrical engines for interplanetary flight. We look at the deep-space interplanetary portion of the trajectory under rather idealized conditions. The two major performance indexes, the propellant mass and the flight time, are in conflict with one another for the following reason: any attempt at reducing the former causes an increase in the latter and vice versa. Therefore, it is natural to consider a compromise performance index involving the scaled values of the propellant mass and flight time weighted respectively by the compromise factor C and its complement 1-C. We use the compromise factor as the parameter of the one-parameter family of compromise trajectories. Analyses carried out with the sequential gradient-restoration algorithm for optimal control problems lead to results which can be highlighted as follows. Thrust profile. Generally speaking, the thrust profile of the compromise trajectory includes three subarcs: the first subarc is characterized by maximum thrust in conjunction with positive (upward) thrust direction; the second subarc is characterized by zero thrust (coasting flight); the third subarc is characterized by maximum thrust in conjunction with negative (downward) thrust direction. Effect of the compromise factor. As the compromise factor increases, the propellant mass decreases and the flight time increases; correspondingly, the following changes in the thrust profile take place: (a) the time lengths of the first and third subarcs (powered phases) decrease slightly, meaning that thrust application occurs for shorter duration; also, the average value of the thrust direction in the first and third subarcs decreases, implying higher efficiency of thrust application wrt the spacecraft energy level; as a result, the total propellant mass decreases; (b) the time length of the second subarc (coasting) increases considerably, resulting in total time increase. Minimum time trajectory. If C=0, the resulting minimum time trajectory has the following characteristics: (a) the time length of the coasting subarc reduces to zero and the three-subarc trajectory degenerates into a two-subarc trajectory; (b) maximum thrust is applied at all times and the thrust direction switches from upward to downward at midcourse. Minimum propellant mass trajectory. If C=1, the resulting minimum propellant mass trajectory has the following characteristics: (a) the thrust magnitude has a bang-zero-bang profile; (b) for the powered subarcs, the thrust direction is tangent to the flight path at all times.

Upper cervical and upper thoracic thrust manipulation versus nonthrust mobilization in patients with mechanical neck pain: a multicenter randomized clinical trial.

PubMed

Dunning, James R; Cleland, Joshua A; Waldrop, Mark A; Arnot, Cathy F; Young, Ian A; Turner, Michael; Sigurdsson, Gisli

2012-01-01

Randomized clinical trial. To compare the short-term effects of upper cervical and upper thoracic high-velocity low-amplitude (HVLA) thrust manipulation to nonthrust mobilization in patients with neck pain. Although upper cervical and upper thoracic HVLA thrust manipulation and nonthrust mobilization are common interventions for the management of neck pain, no studies have directly compared the effects of both upper cervical and upper thoracic HVLA thrust manipulation to nonthrust mobilization in patients with neck pain. Patients completed the Neck Disability Index, the numeric pain rating scale, the flexion-rotation test for measurement of C1-2 passive rotation range of motion, and the craniocervical flexion test for measurement of deep cervical flexor motor performance. Following the baseline evaluation, patients were randomized to receive either HVLA thrust manipulation or nonthrust mobilization to the upper cervical (C1-2) and upper thoracic (T1-2) spines. Patients were reexamined 48-hours after the initial examination and again completed the outcome measures. The effects of treatment on disability, pain, C1-2 passive rotation range of motion, and motor performance of the deep cervical flexors were examined with a 2-by-2 mixed-model analysis of variance (ANOVA). One hundred seven patients satisfied the eligibility criteria, agreed to participate, and were randomized into the HVLA thrust manipulation (n = 56) and nonthrust mobilization (n = 51) groups. The 2-by-2 ANOVA demonstrated that patients with mechanical neck pain who received the combination of upper cervical and upper thoracic HVLA thrust manipulation experienced significantly (P<.001) greater reductions in disability (50.5%) and pain (58.5%) than those of the nonthrust mobilization group (12.8% and 12.6%, respectively) following treatment. In addition, the HVLA thrust manipulation group had significantly (P<.001) greater improvement in both passive C1-2 rotation range of motion and motor performance of the deep cervical flexor muscles as compared to the group that received nonthrust mobilization. The number needed to treat to avoid an unsuccessful outcome was 1.8 and 2.3 at 48-hour follow-up, using the global rating of change and Neck Disability Index cut scores, respectively. The combination of upper cervical and upper thoracic HVLA thrust manipulation is appreciably more effective in the short term than nonthrust mobilization in patients with mechanical neck pain. Therapy, level 1b.

An approximate theoretical method for modeling the static thrust performance of non-axisymmetric two-dimensional convergent-divergent nozzles. M.S. Thesis - George Washington Univ.

NASA Technical Reports Server (NTRS)

Hunter, Craig A.

1995-01-01

An analytical/numerical method has been developed to predict the static thrust performance of non-axisymmetric, two-dimensional convergent-divergent exhaust nozzles. Thermodynamic nozzle performance effects due to over- and underexpansion are modeled using one-dimensional compressible flow theory. Boundary layer development and skin friction losses are calculated using an approximate integral momentum method based on the classic karman-Polhausen solution. Angularity effects are included with these two models in a computational Nozzle Performance Analysis Code, NPAC. In four different case studies, results from NPAC are compared to experimental data obtained from subscale nozzle testing to demonstrate the capabilities and limitations of the NPAC method. In several cases, the NPAC prediction matched experimental gross thrust efficiency data to within 0.1 percent at a design NPR, and to within 0.5 percent at off-design conditions.

Flight Measurements of the Effect of a Controllable Thrust Reverser on the Flight Characteristics of a Single-Engine Jet Airplane

NASA Technical Reports Server (NTRS)

Anderson, Seth B.; Cooper, George E.; Faye, Alan E., Jr.

1959-01-01

A flight investigation was undertaken to determine the effect of a fully controllable thrust reverser on the flight characteristics of a single-engine jet airplane. Tests were made using a cylindrical target-type reverser actuated by a hydraulic cylinder through a "beep-type" cockpit control mounted at the base of the throttle. The thrust reverser was evaluated as an in-flight decelerating device, as a flight path control and airspeed control in landing approach, and as a braking device during the ground roll. Full deflection of the reverser for one reverser configuration resulted in a reverse thrust ratio of as much as 85 percent, which at maximum engine power corresponded to a reversed thrust of 5100 pounds. Use of the reverser in landing approach made possible a wide selection of approach angles, a large reduction in approach speed at steep approach angles, improved control of flight path angle, and more accuracy in hitting a given touchdown point. The use of the reverser as a speed brake at lower airspeeds was compromised by a longitudinal trim change. At the lower airspeeds and higher engine powers there was insufficient elevator power to overcome the nose-down trim change at full reverser deflection.

Numerical and experimental investigations of micro air bearings for micro systems

NASA Astrophysics Data System (ADS)

Zhang, Qide; Shan, X. C.

2006-04-01

The paper investigated performance of air bearing system in a micro device. A parametric study is carried out. The dynamic performance of a very short journal bearing (L/D < 0.1) and thrust bearing is studied. The parameters that affect the performance of the air bearing are discussed. The optimum values of the important parameters are explored, and the stability of the thrust bearing is discussed. The prototype and test result are presented.

Effect of basement structure and salt tectonics on deformation styles along strike: An example from the Kuqa fold-thrust belt, West China

NASA Astrophysics Data System (ADS)

Neng, Yuan; Xie, Huiwen; Yin, Hongwei; Li, Yong; Wang, Wei

2018-04-01

The Kuqa fold-thrust belt (KFTB) has a complex thrust-system geometry and comprises basement-involved thrusts, décollement thrusts, triangle zones, strike-slip faults, transpressional faults, and pop-up structures. These structures, combined with the effects of Paleogene salt tectonics and Paleozoic basement uplift form a complex structural zone trending E-W. Interpretation and comprehensive analysis of recent high-quality seismic data, field observations, boreholes, and gravity data covering the KFTB has been performed to understand the characteristics and mechanisms of the deformation styles along strike. Regional sections, fold-thrust system maps of the surface and the sub-salt layer, salt and basement structure distribution maps have been created, and a comprehensive analysis of thrust systems performed. The results indicate that the thrust-fold system in Paleogene salt range can be divided into five segments from east to west: the Kela-3, Keshen, Dabei, Bozi, and Awate segments. In the easternmost and westernmost parts of the Paleogene salt range, strike-slip faulting and basement-involved thrusting are the dominant deformation styles, as basement uplift and the limits of the Cenozoic evaporite deposit are the main controls on deformation. Salt-core detachment fold-thrust systems coincide with areas of salt tectonics, and pop-up, imbricate, and duplex structures are associated with the main thrust faults in the sub-salt layer. Distribution maps of thrust systems, basement structures, and salt tectonics show that Paleozoic basement uplift controlled the Paleozoic foreland basin morphology and the distribution of Cenozoic salt in the KFTB, and thus had a strong influence on the segmented structural deformation and evolution of the fold-thrust belt. Three types of transfer zone are identified, based on the characteristics of the salt layer and basement uplift, and the effects of these zones on the fault systems are evaluated. Basement uplift and the boundary of the salt deposit generated strike-slip faults in the sub-salt layer and supra-salt layers at the basin boundary (Model A). When changes in the basement occurred within the salt basin, strike-slip faults controlled the deformation styles in the sub-salt layer and shear-zone dominated in the supra-salt layer (Model B). A homogeneous basement and discontinues salt layer formed different accommodation zones in the sub- and supra-salt layers (Model C). In the sub-salt layer the thrusts form imbricate structures on the basal décollement, whereas the supra-salt layer shows overlapping, discontinuous faults and folds with kinds of salt tectonics, and has greater structural variation than the sub-salt layer.

CAMELOT: Computational-Analytical Multi-fidElity Low-thrust Optimisation Toolbox

NASA Astrophysics Data System (ADS)

Di Carlo, Marilena; Romero Martin, Juan Manuel; Vasile, Massimiliano

2018-03-01

Computational-Analytical Multi-fidElity Low-thrust Optimisation Toolbox (CAMELOT) is a toolbox for the fast preliminary design and optimisation of low-thrust trajectories. It solves highly complex combinatorial problems to plan multi-target missions characterised by long spirals including different perturbations. To do so, CAMELOT implements a novel multi-fidelity approach combining analytical surrogate modelling and accurate computational estimations of the mission cost. Decisions are then made using two optimisation engines included in the toolbox, a single-objective global optimiser, and a combinatorial optimisation algorithm. CAMELOT has been applied to a variety of case studies: from the design of interplanetary trajectories to the optimal de-orbiting of space debris and from the deployment of constellations to on-orbit servicing. In this paper, the main elements of CAMELOT are described and two examples, solved using the toolbox, are presented.

Fuel-optimal, low-thrust transfers between libration point orbits

NASA Astrophysics Data System (ADS)

Stuart, Jeffrey R.

Mission design requires the efficient management of spacecraft fuel to reduce mission cost, increase payload mass, and extend mission life. High efficiency, low-thrust propulsion devices potentially offer significant propellant reductions. Periodic orbits that exist in a multi-body regime and low-thrust transfers between these orbits can be applied in many potential mission scenarios, including scientific observation and communications missions as well as cargo transport. In light of the recent discovery of water ice in lunar craters, libration point orbits that support human missions within the Earth-Moon region are of particular interest. This investigation considers orbit transfer trajectories generated by a variable specific impulse, low-thrust engine with a primer-vector-based, fuel-optimizing transfer strategy. A multiple shooting procedure with analytical gradients yields rapid solutions and serves as the basis for an investigation into the trade space between flight time and consumption of fuel mass. Path and performance constraints can be included at node points along any thrust arc. Integration of invariant manifolds into the design strategy may also yield improved performance and greater fuel savings. The resultant transfers offer insight into the performance of the variable specific impulse engine and suggest novel implementations of conventional impulsive thrusters. Transfers incorporating invariant manifolds demonstrate the fuel savings and expand the mission design capabilities that are gained by exploiting system symmetry. A number of design applications are generated.

Thrust Augmented Nozzle for a Hybrid Rocket with a Helical Fuel Port

NASA Astrophysics Data System (ADS)

Marshall, Joel H.

A thrust augmented nozzle for hybrid rocket systems is investigated. The design lever-ages 3-D additive manufacturing to embed a helical fuel port into the thrust chamber of a hybrid rocket burning gaseous oxygen and ABS plastic as propellants. The helical port significantly increases how quickly the fuel burns, resulting in a fuel-rich exhaust exiting the nozzle. When a secondary gaseous oxygen flow is injected into the nozzle downstream of the throat, all of the remaining unburned fuel in the plume spontaneously ignites. This secondary reaction produces additional high pressure gases that are captured by the nozzle and significantly increases the motor's performance. Secondary injection and combustion allows a high expansion ratio (area of the nozzle exit divided by area of the throat) to be effective at low altitudes where there would normally be significantly flow separation and possibly an embedded shock wave due. The result is a 15 percent increase in produced thrust level with no loss in engine efficiency due to secondary injection. Core flow efficiency was increased significantly. Control tests performed using cylindrical fuel ports with secondary injection, and helical fuel ports without secondary injection did not exhibit this performance increase. Clearly, both the fuel-rich plume and secondary injection are essential features allowing the hybrid thrust augmentation to occur. Techniques for better design optimization are discussed.

Recommended Practices in Thrust Measurements

DTIC Science & Technology

2013-10-01

Turin.5,38 This stand consists of two BeCu plates which hang from flexible BeCu mounts on a rigid block of Zerodur c, a material with a very low coe...2013 Figure 4. Example of a state-of-the-art hanging pendulum thrust stand. 38 Two spherical mirrors mounted on the plates form an optical cavity for...the Zerodur frame. Temperature control and careful choice of materials were used to minimize and correct for thermal drift. 2. Thrust Stand Performance

Thrust Augmentation Study of Cross-Flow Fan for Vertical Take-Off and Landing Aircraft

DTIC Science & Technology

2012-09-01

configuration by varying the gap between the CFFs. Computational fluid simulations of the dual CFF configuration was performed using ANSYS CFX to find the...Computational fluid simulations of the dual CFF configuration was performed using ANSYS CFX to find the thrust generated as well as the optimal operating point...RECOMMENDATIONS ...............................................................................43 APPENDIX A. ANSYS CFX SETTINGS FOR DUAL CFF (8,000

Analysis of an all-metallic resilient pad gas-lubricated thrust bearing

NASA Technical Reports Server (NTRS)

Anderson, W. J.

1974-01-01

A new type of resilient pad gas thrust bearing that does not contain any elastomers in the bearing assembly is described and analyzed. The bearing consists of sector-shaped pads mounted asymmetrically on resilient foil beams. The effects of bearing design parameters on performance are shown. Performance of a resilient pad bearing is compared with that of a pivoted pad bearing.

Experimental Study on the Propulsion Performance of the M-shape flapping wing’s bending angle

NASA Astrophysics Data System (ADS)

Chen, Jingxian; Nie, Xiaofang; Zhou, Ximing

2017-10-01

To study the the effect of flapping wing with different bending angles α on the thrust, in this paper, 9 M-shape flapping wing models with different bending angles, ranging for 0° to 22°, were designed. The rotating arm experiment was adopted to conduct the thrust test on the flapping wing models with different bending angels under the wind speed of 15m/s. The result shows that the span-wise flapping wing’s curvature could rectify the airflow, the proper curvature could prevent the span-wise airflow at the surface the flapping wing and leads the airflow towards backward, the amount of air pushed backwards by the flapping wing is larger, therefore the value of thrust is increased; As well as the rectification of M-shape flapping wing increases the thrust value, the flapping wing’s form drag also increased due to the bending angle. According to the results of the experiment, when the bending angle is less than 12°, the increment of the thrust is larger than the decrease of the form drag, so the thrust value increases gradually. However, when the bending angle is larger than 12°, the increment of the thrust is less than the decrease of the form drag, so the thrust value decreases. The thrust value is the largest when the bending angle is 12°.

PPT Thrust Stand

NASA Technical Reports Server (NTRS)

Haag, Thomas W.

1995-01-01

A torsional-type thrust stand has been designed and built to test Pulsed Plasma Thrusters (PPT's) in both single shot and repetitive operating modes. Using this stand, momentum per pulse was determined strictly as a function of thrust stand deflection, spring constant, and natural frequency. No empirical corrections were required. The accuracy of the method was verified using a swinging impact pendulum. Momentum transfer data between the thrust stand and the pendulum were consistent to within 1%. Following initial calibrations, the stand was used to test a Lincoln Experimental Satellite (LES-8/9) thruster. The LES-8/9 system had a mass of approximately 7.5 kg, with a nominal thrust to weight ratio of 1.3 x 10(exp -5). A total of 34 single shot thruster pulses were individually measured. The average impulse bit per pulse was 266 microN-s, which was slightly less than the value of 300 microN-s published in previous reports on this device. Repetitive pulse measurements were performed similar to ordinary steady-state thrust measurements. The thruster was operated for 30 minutes at a repetition rate of 132 pulses per minute and yielded an average thrust of 573 microN. Using average thrust, the average impulse bit per pulse was estimated to be 260 microN-s, which was in agreement with the single shot data. Zero drift during the repetitive pulse test was found to be approximately 1% of the measured thrust.

A new method for optimization of low-thrust gravity-assist sequences

NASA Astrophysics Data System (ADS)

Maiwald, V.

2017-09-01

Recently missions like Hayabusa and Dawn have shown the relevance and benefits of low-thrust spacecraft concerning the exploration of our solar system. In general, the efficiency of low-thrust propulsion is one means of improving mission payload mass. At the same time, gravity-assist maneuvers can serve as mission enablers, as they have the capability to provide "free energy." A combination of both, gravity-assist and low-thrust propulsion, has the potential to generally improve mission performance, i.e. planning and optimization of gravity-assist sequences for low-thrust missions is a desirable asset. Currently no established methods exist to include the gravity-assist partners as optimization variable for low-thrust missions. The present paper explains how gravity-assists are planned and optimized, including the gravity-assist partners, for high-thrust missions and discusses the possibility to transfer the established method, based on the Tisserand Criterion, to low-thrust missions. It is shown how the Tisserand Criterion needs to be adapted using a correction term for the low-thrust situation. It is explained why this necessary correction term excludes an a priori evaluation of sequences and therefore their planning and an alternate approach is proposed. Preliminary results of this method, by application of a Differential Evolution optimization algorithm, are presented and discussed, showing that the method is valid but can be improved. Two constraints on the search space are briefly presented for that aim.

Hybrid propulsion technology program. Volume 1: Conceptional design package

NASA Technical Reports Server (NTRS)

Jensen, Gordon E.; Holzman, Allen L.; Leisch, Steven O.; Keilbach, Joseph; Parsley, Randy; Humphrey, John

1989-01-01

A concept design study was performed to configure two sizes of hybrid boosters; one which duplicates the advanced shuttle rocket motor vacuum thrust time curve and a smaller, quarter thrust level booster. Two sizes of hybrid boosters were configured for either pump-fed or pressure-fed oxygen feed systems. Performance analyses show improved payload capability relative to a solid propellant booster. Size optimization and fuel safety considerations resulted in a 4.57 m (180 inch) diameter large booster with an inert hydrocarbon fuel. The preferred diameter for the quarter thrust level booster is 2.53 m (96 inches). As part of the design study critical technology issues were identified and a technology acquisition and demonstration plan was formulated.

Hybrid propulsion technology program. Volume 2: Technology definition package

NASA Technical Reports Server (NTRS)

Jensen, Gordon E.; Holzman, Allen L.; Leisch, Steven O.; Keilbach, Joseph; Parsley, Randy; Humphrey, John

1989-01-01

A concept design study was performed to configure two sizes of hybrid boosters; one which duplicates the advanced shuttle rocket motor vacuum thrust time curve and a smaller, quarter thrust level booster. Two sizes of hybrid boosters were configured for either pump-fed or pressure-fed oxygen feed systems. Performance analyses show improved payload capability relative to a solid propellant booster. Size optimization and fuel safety considerations resulted in a 4.57 m (180 inch) diameter large booster with an inert hydrocarbon fuel. The preferred diameter for the quarter thrust level booster is 2.53 m (96 inches). The demonstration plan would culminate with test firings of a 3.05 m (120 inch) diameter hybrid booster.

Geometric effects in applied-field MPD thrusters

NASA Technical Reports Server (NTRS)

Myers, R. M.; Mantenieks, M.; Sovey, J.

1990-01-01

Three applied-field magnetoplasmadynamic (MPD) thruster geometries were tested with argon propellant to establish the influence of electrode geometry on thruster performance. The thrust increased approximately linearly with anode radius, while the discharge and electrode fall voltages increased quadratically with anode radius. All these parameters increased linearly with applied-field strength. Thrust efficiency, on the other hand, was not significantly influenced by changes in geometry over the operating range studied, though both thrust and thermal efficiencies increased monotonically with applied field strength. The best performance, 1820 sec I (sub sp) at 20 percent efficiency, was obtained with the largest radius anode at the highest discharge current (1500 amps) and applied field strength (0.4 Tesla).

Geometric effects in applied-field MPD thrusters

NASA Technical Reports Server (NTRS)

Myers, R. M.; Mantenieks, M.; Sovey, James S.

1990-01-01

Three applied-field magnetoplasmadynamic (MPD) thruster geometries were tested with argon propellant to establish the influence of electrode geometry on thruster performance. The thrust increased approximately linearly with anode radius, while the discharge and electrode fall voltages increased quadratically with anode radius. All these parameters increased linearly with applied-field strength. Thrust efficiency, on the other hand, was not significantly influenced by changes in geometry over the operating range studied, though both thrust and thermal efficiencies increased monotonically with applied field strength. The best performance, 1820 sec I(sub sp) at 20 percent efficiency, was obtained with the largest radius anode at the highest discharge current (1500 amps) and applied field strength (0.4 Tesla).

100-lbf LO2/CH4 RCS Thruster Testing and Validation

NASA Technical Reports Server (NTRS)

Barnes, Frank; Cannella, Matthew; Gomez, Carlos; Hand, Jeffrey; Rosenberg, David

2009-01-01

100 pound thrust liquid Oxygen-Methane thruster sized for RCS (Reaction Control System) applications. Innovative Design Characteristics include: a) Simple compact design with minimal part count; b) Gaseous or Liquid propellant operation; c) Affordable and Reusable; d) Greater flexibility than existing systems; e) Part of NASA'S study of "Green Propellants." Hot-fire testing validated performance and functionality of thruster. Thruster's dependence on mixture ratio has been evaluated. Data has been used to calculate performance parameters such as thrust and Isp. Data has been compared with previous test results to verify reliability and repeatability. Thruster was found to have an Isp of 131 s and 82 lbf thrust at a mixture ratio of 1.62.

Evaluation of Phantom-Based Education System for Acupuncture Manipulation

PubMed Central

Lee, In-Seon; Lee, Ye-Seul; Park, Hi-Joon; Lee, Hyejung; Chae, Younbyoung

2015-01-01

Background Although acupuncture manipulation has been regarded as one of the important factors in clinical outcome, it has been difficult to train novice students to become skillful experts due to a lack of adequate educational program and tools. Objectives In the present study, we investigated whether newly developed phantom acupoint tools would be useful to practice-naïve acupuncture students for practicing the three different types of acupuncture manipulation to enhance their skills. Methods We recruited 12 novice students and had them practice acupuncture manipulations on the phantom acupoint (5% agarose gel). We used the Acusensor 2 and compared their acupuncture manipulation techniques, for which the target criteria were depth and time factors, at acupoint LI11 in the human body before and after 10 training sessions. The outcomes were depth of needle insertion, depth error from target criterion, time of rotating, lifting, and thrusting, time error from target criteria and the time ratio. Results After 10 training sessions, the students showed significantly improved outcomes in depth of needle, depth error (rotation, reducing lifting/thrusting), thumb-forward time error, thumb-backward time error (rotation), and lifting time (reinforcing lifting/thrusting). Conclusions The phantom acupoint tool could be useful in a phantom-based education program for acupuncture-manipulation training for students. For advanced education programs for acupuncture manipulation, we will need to collect additional information, such as patient responses, acupoint-specific anatomical characteristics, delicate tissue-like modeling, haptic and visual feedback, and data from an acupuncture practice simulator. PMID:25689598

Flight Management System Execution of Idle-Thrust Descents in Operations

NASA Technical Reports Server (NTRS)

Stell, Laurel L.

2011-01-01

To enable arriving aircraft to fly optimized descents computed by the flight management system (FMS) in congested airspace, ground automation must accurately predict descent trajectories. To support development of the trajectory predictor and its error models, commercial flights executed idle-thrust descents, and the recorded data includes the target speed profile and FMS intent trajectories. The FMS computes the intended descent path assuming idle thrust after top of descent (TOD), and any intervention by the controllers that alters the FMS execution of the descent is recorded so that such flights are discarded from the analysis. The horizontal flight path, cruise and meter fix altitudes, and actual TOD location are extracted from the radar data. Using more than 60 descents in Boeing 777 aircraft, the actual speeds are compared to the intended descent speed profile. In addition, three aspects of the accuracy of the FMS intent trajectory are analyzed: the meter fix crossing time, the TOD location, and the altitude at the meter fix. The actual TOD location is within 5 nmi of the intent location for over 95% of the descents. Roughly 90% of the time, the airspeed is within 0.01 of the target Mach number and within 10 KCAS of the target descent CAS, but the meter fix crossing time is only within 50 sec of the time computed by the FMS. Overall, the aircraft seem to be executing the descents as intended by the designers of the onboard automation.

Evaluation of phantom-based education system for acupuncture manipulation.

PubMed

Lee, In-Seon; Lee, Ye-Seul; Park, Hi-Joon; Lee, Hyejung; Chae, Younbyoung

2015-01-01

Although acupuncture manipulation has been regarded as one of the important factors in clinical outcome, it has been difficult to train novice students to become skillful experts due to a lack of adequate educational program and tools. In the present study, we investigated whether newly developed phantom acupoint tools would be useful to practice-naïve acupuncture students for practicing the three different types of acupuncture manipulation to enhance their skills. We recruited 12 novice students and had them practice acupuncture manipulations on the phantom acupoint (5% agarose gel). We used the Acusensor 2 and compared their acupuncture manipulation techniques, for which the target criteria were depth and time factors, at acupoint LI11 in the human body before and after 10 training sessions. The outcomes were depth of needle insertion, depth error from target criterion, time of rotating, lifting, and thrusting, time error from target criteria and the time ratio. After 10 training sessions, the students showed significantly improved outcomes in depth of needle, depth error (rotation, reducing lifting/thrusting), thumb-forward time error, thumb-backward time error (rotation), and lifting time (reinforcing lifting/thrusting). The phantom acupoint tool could be useful in a phantom-based education program for acupuncture-manipulation training for students. For advanced education programs for acupuncture manipulation, we will need to collect additional information, such as patient responses, acupoint-specific anatomical characteristics, delicate tissue-like modeling, haptic and visual feedback, and data from an acupuncture practice simulator.

Tubular copper thrust chamber design study

NASA Technical Reports Server (NTRS)

Masters, A. I.; Galler, D. E.

1992-01-01

The use of copper tubular thrust chambers is particularly important in high performance expander cycle space engines. Tubular chambers have more surface area than flat wall chambers, and this extra surface area provides enhanced heat transfer for additional energy to power the cycle. This paper was divided into two sections: (1) a thermal analysis and sensitivity study; and (2) a preliminary design of a selected thrust chamber configuration. The thermal analysis consisted of a statistical optimization to determine the optimum tube geometry, tube booking, thrust chamber geometry, and cooling routing to achieve the maximum upper limit chamber pressure for a 25,000 pound thrust engine. The preliminary design effort produced a layout drawing of a tubular thrust chamber that is three inches shorter than the Advanced Expander Test Bed (AETB) milled channel chamber but is predicted to provide a five percent increase in heat transfer. Testing this chamber in the AETB would confirm the inherent advantages of tubular chamber construction and heat transfer.

Characterization of DBD Plasma Actuators Performance without External Flow . Part I; Thrust-Voltage Quadratic Relationship in Logarithmic Space for Sinusoidal Excitation

NASA Technical Reports Server (NTRS)

Ashpis, David E.; Laun, Matthew C.

2016-01-01

We present results of thrust measurements of Dielectric Barrier Discharge (DBD) plasma actuators. We have used a test setup, measurement, and data processing methodology that we developed in prior work. The tests were conducted with High Density Polyethylene (HDPE) actuators of three thicknesses. The applied voltage driving the actuators was a pure sinusoidal waveform. The test setup was suspended actuators with a partial liquid interface. The tests were conducted at low ambient humidity. The thrust was measured with an analytical balance and the results were corrected for anti-thrust to isolate the plasma generated thrust. Applying this approach resulted in smooth and repeatable data. It also enabled curve fitting that yielded quadratic relations between the plasma thrust and voltage in log-log space at constant frequencies. The results contrast power law relationships developed in literature that appear to be a rough approximation over a limited voltage range.

A static investigation of the thrust vectoring system of the F/A-18 high-alpha research vehicle

NASA Technical Reports Server (NTRS)

Mason, Mary L.; Capone, Francis J.; Asbury, Scott C.

1992-01-01

A static (wind-off) test was conducted in the static test facility of the Langley 16-foot Transonic Tunnel to evaluate the vectoring capability and isolated nozzle performance of the proposed thrust vectoring system of the F/A-18 high alpha research vehicle (HARV). The thrust vectoring system consisted of three asymmetrically spaced vanes installed externally on a single test nozzle. Two nozzle configurations were tested: A maximum afterburner-power nozzle and a military-power nozzle. Vane size and vane actuation geometry were investigated, and an extensive matrix of vane deflection angles was tested. The nozzle pressure ratios ranged from two to six. The results indicate that the three vane system can successfully generate multiaxis (pitch and yaw) thrust vectoring. However, large resultant vector angles incurred large thrust losses. Resultant vector angles were always lower than the vane deflection angles. The maximum thrust vectoring angles achieved for the military-power nozzle were larger than the angles achieved for the maximum afterburner-power nozzle.

An Oil-Free Thrust Foil Bearing Facility Design, Calibration, and Operation

NASA Technical Reports Server (NTRS)

Bauman, Steve

2005-01-01

New testing capabilities are needed in order to foster thrust foil air bearing technology development and aid its transition into future Oil-Free gas turbines. This paper describes a new test apparatus capable of testing thrust foil air bearings up to 100 mm in diameter at speeds to 80,000 rpm and temperatures to 650 C (1200 F). Measured parameters include bearing torque, load capacity, and bearing temperatures. This data will be used for design performance evaluations and for validation of foil bearing models. Preliminary test results demonstrate that the rig is capable of testing thrust foil air bearings under a wide range of conditions which are anticipated in future Oil-Free gas turbines. Torque as a function of speed and temperature corroborates results expected from rudimentary performance models. A number of bearings were intentionally failed with no resultant damage whatsoever to the test rig. Several test conditions (specific speeds and loads) revealed undesirable axial shaft vibrations which have been attributed to the magnetic bearing control system and are under study. Based upon these preliminary results, this test rig will be a valuable tool for thrust foil bearing research, parametric studies and technology development.

Annular Internal-External-Expansion Rocket Nozzles for Large Booster Applications

NASA Technical Reports Server (NTRS)

Connors, James F.; Cubbison, Robert W.; Mitchell, Glenn A.

1961-01-01

For large-thrust booster applications, annular rocket nozzles employing both internal and external expansion are investigated. In these nozzles, free-stream air flows through the center as well as around the outside of the exiting jet. Flaps for deflecting the rocket exhaust are incorporated on the external-expansion surface for thrust-vector control. In order to define nozzle off-design performance, thrust vectoring effectiveness, and external stream effects, an experimental investigation was conducted on two annular nozzles with area ratios of 15 and 25 at Mach 0, 2, and 3 in the Lewis 10- by 10-foot wind tunnel. Air, pressurized to 600 pounds per square inch absolute, was used to simulate the exhaust flow. For a nozzle-pressure-ratio range of 40 to 1000, the ratio of actual to ideal thrust was essentially constant at 0.98 for both nozzles. Compared with conventional convergent-divergent configurations on hypothetical boost missions, the performance gains of the annular nozzle could yield significant orbital payload increases (possibly 8 to 17 percent). A single flap on the external-expansion surface of the area-ratio-25 annular nozzle produced a side force equal to 4 percent of the axial force with no measurable loss in axial thrust.

A Note on Rocket Performance Comparison Through Impulse and Thrust Coefficients

NASA Astrophysics Data System (ADS)

Taylor, N. V.

Comparison of rocket motor systems is important when generating data to be used in making design decisions. In order to present meaningful comparisons, non-dimensional numbers related to performance are beneficial, as they remove effects of scale. Traditionally thrust coefficients and C* have been used to quantify the aerodynamic and chemical performance of a system respectively. However, it is argued here that in fact the thrust coefficient does not fully account for aerodynamic performance, as the impact of non-uniform flow at the throat is not accounted for. This discharge coefficient is usually allocated to the chemical efficiency through a correction to C*. However, this causes a coupling between chemical and aerodynamic efficiencies which may lead to poor design decisions. Through the use of a specific impulse coefficient, this risk is avoided, and furthermore comparison of unconventional nozzles becomes more straightforward. It is admitted, however, that this has no actual impact on real motor performance, being more in the way of a tidier `accounting' system.

Theoretical peak performance and optical constraints for the deflection of an S-type asteroid with a continuous wave laser

NASA Astrophysics Data System (ADS)

Thiry, Nicolas; Vasile, Massimiliano

2017-03-01

This paper presents a theoretical model to evaluate the thrust generated by a continuous wave (CW) laser, operating at moderate intensity (<100 GW/m2), ablating an S-type asteroid made of Forsterite. The key metric to assess the performance of the laser system is the thrust coupling coefficient which is given by the ratio between thrust and associated optical power. Three different models are developed in the paper: a one dimensional steady state model, a full 3D steady state model and a one dimensional model accounting for transient effects resulting from the tumbling motion of the asteroid. The results obtained with these models are used to derive key requirements and constraints on the laser system that allow approaching the ideal performance in a realistic case.

Orbit transfer vehicle engine study, phase A extension. Volume 2A: Study results

NASA Technical Reports Server (NTRS)

1980-01-01

Engine trade studies and systems analyses leading to a baseline engine selection for advanced expander cycle engine are discussed with emphasis on: (1) performance optimization of advanced expander cycle engines in the 10 to 20K pound thrust range; (2) selection of a recommended advanced expander engine configuration based on maximized performance and minimized mission risk, and definition of the components for this configuration; (3) characterization of the low thrust adaptation requirements and performance for the staged combustion engine; (4) generation of a suggested safety and reliability approach for OTV engines independent of engine cycle; (5) definition of program risk relationships between expander and staged combustion cycle engines; and (6) development of schedules and costs for the DDT&E, production, and operation phases of the 10K pound thrust expander engine program.

Performance Evaluation of the T6 Ion Engine

NASA Technical Reports Server (NTRS)

Snyder, John Steven; Goebel, Dan M.; Hofer, Richard R.; Polk, James E.; Wallace, Neil C.; Simpson, Huw

2010-01-01

The T6 ion engine is a 22-cm diameter, 4.5-kW Kaufman-type ion thruster produced by QinetiQ, Ltd., and is baselined for the European Space Agency BepiColombo mission to Mercury and is being qualified under ESA sponsorship for the extended range AlphaBus communications satellite platform. The heritage of the T6 includes the T5 ion thruster now successfully operating on the ESA GOCE spacecraft. As a part of the T6 development program, an engineering model thruster was subjected to a suite of performance tests and plume diagnostics at the Jet Propulsion Laboratory. The engine was mounted on a thrust stand and operated over its nominal throttle range of 2.5 to 4.5 kW. In addition to the typical electrical and flow measurements, an E x B mass analyzer, scanning Faraday probe, thrust vector probe, and several near-field probes were utilized. Thrust, beam divergence, double ion content, and thrust vector movement were all measured at four separate throttle points. The engine performance agreed well with published data on this thruster. At full power the T6 produced 143 mN of thrust at a specific impulse of 4120 seconds and an efficiency of 64%; optimization of the neutralizer for lower flow rates increased the specific impulse to 4300 seconds and the efficiency to nearly 66%. Measured beam divergence was less than, and double ion content was greater than, the ring-cusp-design NSTAR thruster that has flown on NASA missions. The measured thrust vector offset depended slightly on throttle level and was found to increase with time as the thruster approached thermal equilibrium.

Experimental and Computational Investigation of Multiple Injection Ports in a Convergent-Divergent Nozzle for Fluidic Thrust Vectoring

NASA Technical Reports Server (NTRS)

Waithe, Kenrick A.; Deere, Karen A.

2003-01-01

A computational and experimental study was conducted to investigate the effects of multiple injection ports in a two-dimensional, convergent-divergent nozzle, for fluidic thrust vectoring. The concept of multiple injection ports was conceived to enhance the thrust vectoring capability of a convergent-divergent nozzle over that of a single injection port without increasing the secondary mass flow rate requirements. The experimental study was conducted at static conditions in the Jet Exit Test Facility of the 16-Foot Transonic Tunnel Complex at NASA Langley Research Center. Internal nozzle performance was obtained at nozzle pressure ratios up to 10 with secondary nozzle pressure ratios up to 1 for five configurations. The computational study was conducted using the Reynolds Averaged Navier-Stokes computational fluid dynamics code PAB3D with two-equation turbulence closure and linear Reynolds stress modeling. Internal nozzle performance was predicted for nozzle pressure ratios up to 10 with a secondary nozzle pressure ratio of 0.7 for two configurations. Results from the experimental study indicate a benefit to multiple injection ports in a convergent-divergent nozzle. In general, increasing the number of injection ports from one to two increased the pitch thrust vectoring capability without any thrust performance penalties at nozzle pressure ratios less than 4 with high secondary pressure ratios. Results from the computational study are in excellent agreement with experimental results and validates PAB3D as a tool for predicting internal nozzle performance of a two dimensional, convergent-divergent nozzle with multiple injection ports.

A model for teaching and learning spinal thrust manipulation and its effect on participant confidence in technique performance

PubMed Central

Wise, Christopher H.; Schenk, Ronald J.; Lattanzi, Jill Black

2016-01-01

Background Despite emerging evidence to support the use of high velocity thrust manipulation in the management of lumbar spinal conditions, utilization of thrust manipulation among clinicians remains relatively low. One reason for the underutilization of these procedures may be related to disparity in training in the performance of these techniques at the professional and post professional levels. Purpose To assess the effect of using a new model of active learning on participant confidence in the performance of spinal thrust manipulation and the implications for its use in the professional and post-professional training of physical therapists. Methods A cohort of 15 DPT students in their final semester of entry-level professional training participated in an active training session emphasizing a sequential partial task practice (SPTP) strategy in which participants engaged in partial task practice over several repetitions with different partners. Participants’ level of confidence in the performance of these techniques was determined through comparison of pre- and post-training session surveys and a post-session open-ended interview. Results The increase in scores across all items of the individual pre- and post-session surveys suggests that this model was effective in changing overall participant perception regarding the effectiveness and safety of these techniques and in increasing student confidence in their performance. Interviews revealed that participants greatly preferred the SPTP strategy, which enhanced their confidence in technique performance. Conclusion Results indicate that this new model of psychomotor training may be effective at improving confidence in the performance of spinal thrust manipulation and, subsequently, may be useful for encouraging the future use of these techniques in the care of individuals with impairments of the spine. Inasmuch, this method of instruction may be useful for training of physical therapists at both the professional and post-professional levels. PMID:27559284

Inertia and Couple-Stress Effects in a Curvilinear Thrust Hydrostatic Bearing

NASA Astrophysics Data System (ADS)

Walicka, A.; Jurczak, P.; Falicki, J.

2017-08-01

The flow of a couple-stress lubricant in a clearance of a curvilinear thrust hydrostatic bearing with impermeable walls is considered. The flow in the bearing clearance is considered with inertia forces. The equations of motion are solved by an averaged inertia method. As a result, the formulae for pressure distributions without and with inertia effects were obtained. Radial thrust bearings and spherical bearings are discussed as numerical examples. It is shown that inertia effects influence the bearing performance considerably.

Low thrust propulsion literature survey

NASA Technical Reports Server (NTRS)

Monroe, Darrel

1989-01-01

A literature search was performed to investigate the area of low thrust propulsion. In an effort to evaluate this technology, a number of articles, obtained through the use of the NASA-RECON database, were collected and categorized. The study indicates that although much was done, particularly in the 1960's and 1970's, more can be done in the area of practical navigation and guidance. It is suggested that the older studies be reinvestigated to see what potential there exists for future low thrust applications.

Investigation of advanced thrust vectoring exhaust systems for high speed propulsive lift

NASA Technical Reports Server (NTRS)

Hutchison, R. A.; Petit, J. E.; Capone, F. J.; Whittaker, R. W.

1980-01-01

The paper presents the results of a wind tunnel investigation conducted at the NASA-Langley research center to determine thrust vectoring/induced lift characteristics of advanced exhaust nozzle concepts installed on a supersonic tactical airplane model. Specific test objectives include: (1) basic aerodynamics of a wing body configuration, (2) investigation of induced lift effects, (3) evaluation of static and forward speed performance, and (4) the effectiveness of a canard surface to trim thrust vectoring/induced lift forces and moments.

Anomalous Thrust Production from an RF Test Device Measured on a Low-Thrust Torsion Pendulum

NASA Technical Reports Server (NTRS)

Brady, David; White, Harold G.; March, Paul; Lawrence, James T.; Davies, Frank J.

2014-01-01

This paper describes the eight-day August 2013 test campaign designed to investigate and demonstrate viability of using classical magnetoplasmadynamics to obtain a propulsive momentum transfer via the quantum vacuum virtual plasma. This paper will not address the physics of the quantum vacuum plasma thruster, but instead will describe the test integration, test operations, and the results obtained from the test campaign. Approximately 30-50 micro-Newtons of thrust were recorded from an electric propulsion test article consisting primarily of a radio frequency (RF) resonant cavity excited at approximately 935 megahertz. Testing was performed on a low-thrust torsion pendulum that is capable of detecting force at a single-digit micronewton level, within a stainless steel vacuum chamber with the door closed but at ambient atmospheric pressure. Several different test configurations were used, including two different test articles as well as a reversal of the test article orientation. In addition, the test article was replaced by an RF load to verify that the force was not being generated by effects not associated with the test article. The two test articles were designed by Cannae LLC of Doylestown, Pennsylvania. The torsion pendulum was designed, built, and operated by Eagleworks Laboratories at the NASA Johnson Space Center of Houston, Texas. Approximately six days of test integration were required, followed by two days of test operations, during which, technical issues were discovered and resolved. Integration of the two test articles and their supporting equipment was performed in an iterative fashion between the test bench and the vacuum chamber. In other words, the test article was tested on the bench, then moved to the chamber, then moved back as needed to resolve issues. Manual frequency control was required throughout the test. Thrust was observed on both test articles, even though one of the test articles was designed with the expectation that it would not produce thrust. Specifically, one test article contained internal physical modifications that were designed to produce thrust, while the other did not (with the latter being referred to as the "null" test article). Test data gathered includes torsion pendulum displacement measurements which are used to calculate generated force, still imagery in the visible spectrum to document the physical configuration, still imagery in the infrared spectrum to characterize the thermal environment, and video imagery. Post-test data includes static and animated graphics produced during RF resonant cavity characterization using the COMSOL Multiphysics® software application. Excerpts from all of the above are included and discussed in this paper. Lessons learned from test integration and operations include identification of the need to replace manual control of the resonant cavity target frequency with an automated frequency control capability. Future test plans include the development of an automatic frequency control circuit. Test results indicate that the RF resonant cavity thruster design, which is unique as an electric propulsion device, is producing a force that is not attributable to any classical electromagnetic phenomenon and therefore is potentially demonstrating an interaction with the quantum vacuum virtual plasma. Future test plans include independent verification and validation at other test facilities.

Low-Thrust Many-Revolution Trajectory Optimization via Differential Dynamic Programming and a Sundman Transformation

NASA Astrophysics Data System (ADS)

Aziz, Jonathan D.; Parker, Jeffrey S.; Scheeres, Daniel J.; Englander, Jacob A.

2018-01-01

Low-thrust trajectories about planetary bodies characteristically span a high count of orbital revolutions. Directing the thrust vector over many revolutions presents a challenging optimization problem for any conventional strategy. This paper demonstrates the tractability of low-thrust trajectory optimization about planetary bodies by applying a Sundman transformation to change the independent variable of the spacecraft equations of motion to an orbit angle and performing the optimization with differential dynamic programming. Fuel-optimal geocentric transfers are computed with the transfer duration extended up to 2000 revolutions. The flexibility of the approach to higher fidelity dynamics is shown with Earth's J 2 perturbation and lunar gravity included for a 500 revolution transfer.

Low-Thrust Many-Revolution Trajectory Optimization via Differential Dynamic Programming and a Sundman Transformation

NASA Astrophysics Data System (ADS)

Aziz, Jonathan D.; Parker, Jeffrey S.; Scheeres, Daniel J.; Englander, Jacob A.

2018-06-01

Low-thrust trajectories about planetary bodies characteristically span a high count of orbital revolutions. Directing the thrust vector over many revolutions presents a challenging optimization problem for any conventional strategy. This paper demonstrates the tractability of low-thrust trajectory optimization about planetary bodies by applying a Sundman transformation to change the independent variable of the spacecraft equations of motion to an orbit angle and performing the optimization with differential dynamic programming. Fuel-optimal geocentric transfers are computed with the transfer duration extended up to 2000 revolutions. The flexibility of the approach to higher fidelity dynamics is shown with Earth's J 2 perturbation and lunar gravity included for a 500 revolution transfer.

Multiphysics Analysis of a Solid-Core Nuclear Thermal Engine Thrust Chamber

NASA Technical Reports Server (NTRS)

Wang, Ten-See; Canabal, Francisco; Cheng, Gary; Chen, Yen-Sen

2006-01-01

The objective of this effort is to develop an efficient and accurate thermo-fluid computational methodology to predict environments for a hypothetical solid-core, nuclear thermal engine thrust chamber. The computational methodology is based on an unstructured-grid, pressure-based computational fluid dynamics methodology. Formulations for heat transfer in solids and porous media were implemented and anchored. A two-pronged approach was employed in this effort: A detailed thermo-fluid analysis on a multi-channel flow element for mid-section corrosion investigation; and a global modeling of the thrust chamber to understand the effect of hydrogen dissociation and recombination on heat transfer and thrust performance. The formulations and preliminary results on both aspects are presented.

Comparison of straight and 15 degree vectored nozzles using a six component thrust stand

NASA Technical Reports Server (NTRS)

Carpenter, Thomas W.; Flake, Scott

1991-01-01

This project compared the forces and moments produced by straight and 15 degree vectored nozzles. Using the six component thrust stand in the engines laboratory at California Polytechnic State University, several trials were performed. This data was then reduced using first a computer program and then later an electronic spreadsheet. This reduced data was graphed and compared. As a result of these comparisons some unexpected forces were discovered. Several more tests were run including a zero thrust test and a statistical comparison were done to discover the source of these discrepancies. As a direct result several nozzle changes were made and significant revisions to the thrust stand are being made.

Effects of Altitude on Turbojet Engine Performance

NASA Technical Reports Server (NTRS)

Fleming, William A

1951-01-01

Component and over-all performance characteristics of several turbojet engines investigated in the altitude facilities of the NACA Lewis Laboratory during the last several years are summarized to indicate the effects of altitude on turbojet engine performance. Data presented show that failure of turbojet engine performance to generalize for all altitudes can be traced to reductions in compressor efficiency, corrected air flow, and combustion efficiency at altitude. In addition, it is shown that although engines of different design may have equal thrusts at sea level, the thrusts at altitude may vary widely because of differences in compressor performance characteristics from one engine to another.

A review of technologies applicable to low-speed flight of high-performance aircraft investigated in the Langley 14- x 22-foot subsonic tunnel

NASA Technical Reports Server (NTRS)

Paulson, John W., Jr.; Quinto, P. Frank; Banks, Daniel W.; Kemmerly, Guy T.; Gatlin, Gregory M.

1988-01-01

An extensive research program has been underway at the NASA Langley Research Center to define and develop the technologies required for low-speed flight of high-performance aircraft. This 10-year program has placed emphasis on both short takeoff and landing (STOL) and short takeoff and vertical landing (STOVL) operations rather than on regular up and away flight. A series of NASA in-house as well as joint projects have studied various technologies including high lift, vectored thrust, thrust-induced lift, reversed thrust, an alternate method of providing trim and control, and ground effects. These technologies have been investigated on a number of configurations ranging from industry designs for advanced fighter aircraft to generic wing-canard research models. Test conditions have ranged from hover (or static) through transition to wing-borne flight at angles of attack from -5 to 40 deg at representative thrust coefficients.

On-board orbit determination for low thrust LEO-MEO transfer by Consider Kalman Filtering and multi-constellation GNSS

NASA Astrophysics Data System (ADS)

Menzione, Francesco; Renga, Alfredo; Grassi, Michele

2017-09-01

In the framework of the novel navigation scenario offered by the next generation satellite low thrust autonomous LEO-to-MEO orbit transfer, this study proposes and tests a GNSS based navigation system aimed at providing on-board precise and robust orbit determination strategy to override rising criticalities. The analysis introduces the challenging design issues to simultaneously deal with the variable orbit regime, the electric thrust control and the high orbit GNSS visibility conditions. The Consider Kalman Filtering approach is here proposed as the filtering scheme to process the GNSS raw data provided by a multi-antenna/multi-constellation receiver in presence of uncertain parameters affecting measurements, actuation and spacecraft physical properties. Filter robustness and achievable navigation accuracy are verified using a high fidelity simulation of the low-thrust rising scenario and performance are compared with the one of a standard Extended Kalman Filtering approach to highlight the advantages of the proposed solution. Performance assessment of the developed navigation solution is accomplished for different transfer phases.

The Chevron Foil Thrust Bearing: Improved Performance Through Passive Thermal Management and Effective Lubricant Mixing

NASA Technical Reports Server (NTRS)

Bruckner, Robert

2013-01-01

An improved foil thrust bearing is described that eliminates or reduces the need for forced cooling of the bearing foils while at the same time improves the load capacity of the bearing, enhances damping, provides overload tolerance, and eliminates the high speed load capacity drop-off that plagues the current state of the art. The performance improvement demonstrated by the chevron foil thrust bearing stems from a novel trailing edge shape that splays the hot lubricant in the thin film radially, thus preventing hot lubricant carry-over into the ensuing bearing sector. Additionally, the chevron shaped trailing edge induces vortical mixing of the hot lubricant with the gas that is naturally resident within the inter-pad region of a foil thrust bearing. The elimination of hot gas carry-over in combination with the enhanced mixing has enabled a completely passive thermally managed foil bearing design. Laboratory testing at NASA has confirmed the original analysis and reduced this concept to practice.

Analysis of oil lubricated, fluid film, thrust bearings with allowance for temperature dependent viscosity

NASA Technical Reports Server (NTRS)

Pan, C. H. T.; Malanoski, S. B.

1972-01-01

A preliminary design study was performed to seek a fluid-film thrust bearing design intended to be part of a high-speed, hybrid (rolling element/fluid film) bearing configuration. The base line used is a design previously tested. To improve the accuracy of theoretical predictions of load capacity, flow rate, and friction power loss, an analytical procedure was developed to include curvature effects inherent in thrust bearings and to allow for the temperature rise in the fluid due to viscous heating. Also, a narrow-groove approximation in the treatment of the temperature field was formulated to apply the procedure to the Whipple thrust bearing. A comparative trade-off study was carried out assuming isothermal films; its results showed the shrouded-step design to be superior to the Whipple design for the intended application. An extensive parametric study was performed, employing isoviscous calculations, to determine the optimized design, which was subsequently recalculated allowing for temperature effects.

Trajectory optimization for lunar rover performing vertical takeoff vertical landing maneuvers in the presence of terrain

NASA Astrophysics Data System (ADS)

Ma, Lin; Wang, Kexin; Xu, Zuhua; Shao, Zhijiang; Song, Zhengyu; Biegler, Lorenz T.

2018-05-01

This study presents a trajectory optimization framework for lunar rover performing vertical takeoff vertical landing (VTVL) maneuvers in the presence of terrain using variable-thrust propulsion. First, a VTVL trajectory optimization problem with three-dimensional kinematics and dynamics model, boundary conditions, and path constraints is formulated. Then, a finite-element approach transcribes the formulated trajectory optimization problem into a nonlinear programming (NLP) problem solved by a highly efficient NLP solver. A homotopy-based backtracking strategy is applied to enhance the convergence in solving the formulated VTVL trajectory optimization problem. The optimal thrust solution typically has a "bang-bang" profile considering that bounds are imposed on the magnitude of engine thrust. An adaptive mesh refinement strategy based on a constant Hamiltonian profile is designed to address the difficulty in locating the breakpoints in the thrust profile. Four scenarios are simulated. Simulation results indicate that the proposed trajectory optimization framework has sufficient adaptability to handle VTVL missions efficiently.

H2 arcjet performance mapping program

NASA Astrophysics Data System (ADS)

1992-01-01

Work performed during the period of Mar. 1991 to Jan. 1992 is reviewed. High power H2 arcjets are being considered for electric powered orbit transfer vehicles (EOTV). Mission analyses indicate that the overall arcjet thrust efficiency is very important since increasing the efficiency increases the thrust, and thereby reduces the total trip time for the same power. For example, increasing the thrust efficiency at the same specific impulse from 30 to 40 percent will reduce the trip time by 25 percent. For a 200 day mission, this equates to 50 days, which results in lower ground costs and less time during which the payload is dormant. Arcjet performance levels of 1200 seconds specific impulse (lsp) at 35 to 40 percent efficiency with lifetimes over 1000 hours are needed to support EOTV missions. Because of the potential very high efficiency levels, the objective of this program was to evaluate the ability of a scaled Giannini-style thruster to achieve the performance levels while operating at a reduced nominal power of 10 kW. To meet this objective, a review of past literature was conducted; scaling relationships were developed and applied to establish critical dimensions; a development thruster was designed with the aid of the plasma analysis model KARNAC and finite element thermal modeling; test hardware was fabricated; and a series of performance tests were conducted in RRC's Cell 11 vacuum chamber with its null-balance thrust stand.

Characterization of Low-Frequency Combustion Stability of the Fastrac Engine

NASA Technical Reports Server (NTRS)

Rocker, Marvin; Jones, Preston (Technical Monitor)

2002-01-01

A series of tests were conducted to measure the combustion performance of the Fastrac engine thrust chamber. During mainstage, the thrust chamber exhibited no large-amplitude chamber pressure oscillations that could be identified as low-frequency combustion instability or 'chug'. However, during start-up and shutdown, the thrust chamber very briefly exhibited large-amplitude chamber pressure oscillations that were identified as chug. These instabilities during start-up and shutdown were regarded as benign due to their brevity. Linear models of the thrust chamber and the propellant feed systems were formulated for both the thrust chamber component tests and the flight engine tests. These linear models determined the frequency and decay rate of chamber pressure oscillations given the design and operating conditions of the thrust chamber and feed system. The frequency of chamber pressure oscillations determined from the model closely matched the frequency of low-amplitude, low-frequency chamber pressure oscillations exhibited in some of the later thrust chamber mainstage tests. The decay rate of the chamber pressure oscillations determined from the models indicated that these low-frequency oscillations were stable. Likewise, the decay rate, determined from the model of the flight engine tests indicated that the low-frequency chamber pressure oscillations would be stable.

Studies of $${\\rm Nb}_{3}{\\rm Sn}$$ Strands Based on the Restacked-Rod Process for High Field Accelerator Magnets

DOE PAGES

Barzi, E.; Bossert, M.; Gallo, G.; ...

2011-12-21

A major thrust in Fermilab's accelerator magnet R&D program is the development of Nb 3Sn wires which meet target requirements for high field magnets, such as high critical current density, low effective filament size, and the capability to withstand the cabling process. The performance of a number of strands with 150/169 restack design produced by Oxford Superconducting Technology was studied for round and deformed wires. To optimize the maximum plastic strain, finite element modeling was also used as an aid in the design. Results of mechanical, transport and metallographic analyses are presented for round and deformed wires.

Docking simulation analysis of range data requirements for the orbital maneuvering vehicle

NASA Technical Reports Server (NTRS)

Micheal, J. D.; Vinz, F. L.

1985-01-01

The results of an initial study are reported assess the controllability of the Orbital Maneuvering Vehicle (OMV) for terminal closure and docking are reported. The vehicle characteristics used in this study are those of the Marshall Space Flight Center (MSFC) baseline OMV which were published with the request for proposals for preliminary design of this vehicle. This simulation was conducted at MSFC using the Target Motion Simulator. The study focused on the OMV manual mode capability to accommodate both stabilized and tumbling target engagements with varying complements of range and range rate data displayed to the OMV operator. Four trained test subjects performed over 400 simulated orbital dockings during this study. A firm requirement for radar during the terminal closure and dock phase of the OMV mission was not established by these simulations. Fifteen pound thrusters recommended in the MSFC baseline design were found to be advantageous for initial rate matching maneuvers with unstabilized targets; however, lower thrust levels were desirable for making the final docking maneuvers.

A torsion balance for impulse and thrust measurements of micro-Newton thrusters

NASA Astrophysics Data System (ADS)

Yang, Yuan-Xia; Tu, Liang-Cheng; Yang, Shan-Qing; Luo, Jun

2012-01-01

This paper reports the performance of a torsion-type thrust stand suitable for studies of micro-Newton thrusters, which is developed for ground testing the micro-Newton thruster in Chinese Test of the Equivalence Principle with Optical readout space mission. By virtue of specially suspending design and precise assembly of torsion balance configuration, the thrust stand with load capacity up to several kilograms is able to measure the impulse bit up to 1350 μNs with a resolution of 0.47 μNs, and the average thrust up to 264 μN with a resolution of 0.09 μN in both open and close loop operation. A pulsed plasma thruster, the preliminary prototype developed for Chinese TEPO space mission, is tested by the thrust stand, and the results reveal that the average impulse bit per pulse is measured to be 58.4 μNs with a repeatability of about 5%.

The calculated effect of trailing-edge flaps on the take-off of flying boats

NASA Technical Reports Server (NTRS)

Parkinson, J E; Bell, J W

1934-01-01

The results of take-off calculations are given for an application of simple trailing-edge flaps to two hypothetical flying boats, one having medium wing and power loading and consequently considerable excess of thrust over total resistance during the take-off run, the other having high wing and power loading and a very low excess thrust. For these seaplanes the effect of downward flap settings was: (1) to increase the total resistance below the stalling speed, (2) to decrease the get-away speed, (3) to improve the take-off performance of the seaplane having considerable excess thrust, and (4) to hinder the take-off of the seaplane having low excess thrust. It is indicated that flaps would allow a decrease in the high angles of wing setting necessary with most seaplanes, provided that the excess thrust is not too low.

Unsteady Ejector Performance: an Experimental Investigation Using a Pulsejet Driver

NASA Technical Reports Server (NTRS)

Paxson, Daniel E.; Wilson, Jack; Dougherty, Kevin T.

2002-01-01

An experimental investigation is described in which thrust augmentation and mass entrainment were measured for a variety of simple cylindrical ejectors driven by a gasoline-fueled pulsejet. The ejectors were of varying length, diameter, and inlet radius. Measurements were also taken to determine the effect on performance of the distance between pulsejet exit and ejector inlet. Limited tests were also conducted to determine the effect of driver cross-sectional shape. Optimal values were found for all three ejector parameters with respect to thrust augmentation. This was not the case with mass entrainment, which increased monotonically with ejector diameter. Thus, it was found that thrust augmentation is not necessarily directly related to mass entrainment, as is often supposed for ejectors. Peak thrust augmentation values of 1.8 were obtained. Peak mass entrainment values of 30 times the driver mass flow were also observed. Details of the experimental setup and results are presented. Preliminary analysis of the results indicates that the enhanced performance obtained with an unsteady jet (primary source) over comparably sized ejectors driven with steady jets is due primarily to the structure of the starting vortex-type flow associated with the former.

Performance Characteristics of a DME Propellant Arcjet Thruster

NASA Astrophysics Data System (ADS)

Kakami, Akira; Beeppu, Shinji; Maiguma, Muneyuki; Tachibana, Takeshi

This paper describes the influence of cathode configuration on performance of an arcjet thruster using dimethyl ether (DME) propellant. DME, an ether compound, has suitable characteristics for a space propulsion system; DME is storable in a liquid state without being kept under a high pressure, and requires no sophisticated temperature management such as a cryogenic device. DME can be gasified and liquefied simply by adjusting temperature whereas hydrazine, a conventional propellant, requires an iridium-based particulate catalyst for its gasification. In this study, thrust of a 1-kW class DME arcjet thruster is measured at a discharge current of 13 A, DME mass flow rates ranging 15 to 60 mg/s under three cathode configurations: flat-tip rods of 2 and 4 mm in diam. and 4-mm-diam. rod having a cavity of 2 mm in diameter. Thrust measurements show that thrust is increased with propellant mass flow rate. Among the tested cathodes, the flat-tip rod of 4 mm in diam. with 55 mg/s DME flow rate yielded the highest performance: specific impulse of 330 s, thrust of 0.18 N, discharge power of 1400 W and specific power of 25 MJ/kg.

Parametric study of a simultaneous pitch/yaw thrust vectoring single expansion ramp nozzle

NASA Technical Reports Server (NTRS)

Schirmer, Alberto W.; Capone, Francis J.

1989-01-01

In the course of the last eleven years, the concept of thrust vectoring has emerged as a promising method of enhancing aircraft control capabilities in post-stall flight incursions during combat. In order to study the application of simultaneous pitch and yaw vectoring to single expansion ramp nozzles, a static test was conducted in the NASA-Langley 16 foot transonic tunnel. This investigation was based on internal performance data provided by force, mass flow and internal pressure measurements at nozzle pressure ratios up to 8. The internal performance characteristics of the nozzle were studied for several combinations of six different parameters: yaw vectoring angle, pitch vectoring angle, upper ramp cutout, sidewall hinge location, hinge inclination angle and sidewall containment. Results indicated a 2-to- 3-percent decrease in resultant thrust ratio with vectoring in either pitch or yaw. Losses were mostly associated with the turning of supersonic flow. Resultant thrust ratios were also decreased by sideways expansion of the jet. The effects of cutback corners in the upper ramp and lower flap on performance were small. Maximum resultant yaw vector angles, about half of the flap angle, were achieved for the configuration with the most forward hinge location.

Shallow seismic imaging of folds above the Puente Hills blind-thrust fault, Los Angeles, California

USGS Publications Warehouse

Pratt, T.L.; Shaw, J.H.; Dolan, J.F.; Christofferson, S.A.; Williams, R.A.; Odum, J.K.; Plesch, A.

2002-01-01

High-resolution seismic reflection profiles image discrete folds in the shallow subsurface (<600 m) above two segments of the Puente Hills blind-thrust fault system, Los Angeles basin, California. The profiles demonstrate late Quaternary activity at the fault tip, precisely locate the axial surfaces of folds within the upper 100 m, and constrain the geometry and kinematics of recent folding. The Santa Fe Springs segment of the Puente Hills fault zone shows an upward-narrowing kink band with an active anticlinal axial surface, consistent with fault-bend folding above an active thrust ramp. The Coyote Hills segment shows an active synclinal axial surface that coincides with the base of a 9-m-high scarp, consistent with tip-line folding or the presence of a backthrust. The seismic profiles pinpoint targets for future geologic work to constrain slip rates and ages of past events on this important fault system.

Low-thrust mission risk analysis, with application to a 1980 rendezvous with the comet Encke

NASA Technical Reports Server (NTRS)

Yen, C. L.; Smith, D. B.

1973-01-01

A computerized failure process simulation procedure is used to evaluate the risk in a solar electric space mission. The procedure uses currently available thrust-subsystem reliability data and performs approximate simulations of the thrust sybsystem burn operation, the system failure processes, and the retargeting operations. The method is applied to assess the risks in carrying out a 1980 rendezvous mission to the comet Encke. Analysis of the results and evaluation of the effects of various risk factors on the mission show that system component failure rates are the limiting factors in attaining a high mission relability. It is also shown that a well-designed trajectory and system operation mode can be used effectively to partially compensate for unreliable thruster performance.

Thrust performance of a variable-geometry, divergent exhaust nozzle on a turbojet engine at altitude

NASA Technical Reports Server (NTRS)

Straight, D. M.; Collom, R. R.

1983-01-01

A variable geometry, low aspect ratio, nonaxisymmetric, two dimensional, convergent-divergent exhaust nozzle was tested at simulated altitude on a turbojet engine to obtain baseline axial, dry thrust performance over wide ranges of operating nozzle pressure ratios, throat areas, and internal expansion area ratios. The thrust data showed good agreement with theory and scale model test results after the data were corrected for seal leakage and coolant losses. Wall static pressure profile data were also obtained and compared with one dimensional theory and scale model data. The pressure data indicate greater three dimensional flow effects in the full scale tests than with models. The leakage and coolant penalties were substantial, and the method to determine them is included.

Experimental performance of 75-millimeter-bore arched outer-race ball bearings to 2.1 million DN. [thrust loads

NASA Technical Reports Server (NTRS)

Coe, H. H.

1976-01-01

An experimental investigation was performed to determine the operating characteristics of arched outer-race bearings and to compare the data with those for a similar, but conventional, deep-groove ball bearing. The bearings were tested over a range of shaft speeds up to 28,000 rpm at a thrust load of 2200 newtons (500 lb). One bearing was operated at 26,000 rpm with a range of thrust loads. The amounts of arching were 0.13, 0.25, and 0.51 millimeter (0.005, 0.010, and 0.020 in.). All bearings operated satisfactorily; the arched bearing outer-race temperatures and torques were consistently higher than those for the conventional bearing.

Comparison of theoretical and experimental thrust performance of a 1030:1 area ratio rocket nozzle at a chamber pressure of 2413 kN/m2 (350 psia)

NASA Technical Reports Server (NTRS)

Smith, Tamara A.; Pavli, Albert J.; Kacynski, Kenneth J.

1987-01-01

The joint Army. Navy, NASA. Air Force (JANNAF) rocket engine peformnace prediction procedure is based on the use of various reference computer programs. One of the reference programs for nozzle analysis is the Two-Dimensional Kinetics (TDK) Program. The purpose of this report is to calibrate the JANNAF procedure incorporated into the December l984 version of the TDK program for the high-area-ratio rocket engine regime. The calibration was accomplished by modeling the performance of a 1030:1 rocket nozzle tested at NASA Lewis Research Center. A detailed description of the experimental test conditions and TDK input parameters is given. The results show that the computer code predicts delivered vacuum specific impulse to within 0.12 to 1.9 percent of the experimental data. Vacuum thrust coefficient predictions were within + or - 1.3 percent of experimental results. Predictions of wall static pressure were within approximately + or - 5 percent of the measured values. An experimental value for inviscid thrust was obtained for the nozzle extension between area ratios of 427.5 and 1030 by using an integration of the measured wall static pressures. Subtracting the measured thrust gain produced by the nozzle between area ratios of 427.5 and 1030 from the inviscid thrust gain yielded experimental drag decrements of 10.85 and 27.00 N (2.44 and 6.07 lb) for mixture ratios of 3.04 and 4.29, respectively. These values correspond to 0.45 and 1.11 percent of the total vacuum thrust. At a mixture ratio of 4.29, the TDK predicted drag decrement was 16.59 N (3.73 lb), or 0.71 percent of the predicted total vacuum thrust.

A Crewed Mission to Apophis Using a Hybrid Bimodal Nuclear Thermal Electric Propulsion (BNTEP) System

NASA Technical Reports Server (NTRS)

Mccurdy, David R.; Borowski, Stanley K.; Burke, Laura M.; Packard, Thomas W.

2014-01-01

A BNTEP system is a dual propellant, hybrid propulsion concept that utilizes Bimodal Nuclear Thermal Rocket (BNTR) propulsion during high thrust operations, providing 10's of kilo-Newtons of thrust per engine at a high specific impulse (Isp) of 900 s, and an Electric Propulsion (EP) system during low thrust operations at even higher Isp of around 3000 s. Electrical power for the EP system is provided by the BNTR engines in combination with a Brayton Power Conversion (BPC) closed loop system, which can provide electrical power on the order of 100's of kWe. High thrust BNTR operation uses liquid hydrogen (LH2) as reactor coolant propellant expelled out a nozzle, while low thrust EP uses high pressure xenon expelled by an electric grid. By utilizing an optimized combination of low and high thrust propulsion, significant mass savings over a conventional NTR vehicle can be realized. Low thrust mission events, such as midcourse corrections (MCC), tank settling burns, some reaction control system (RCS) burns, and even a small portion at the end of the departure burn can be performed with EP. Crewed and robotic deep space missions to a near Earth asteroid (NEA) are best suited for this hybrid propulsion approach. For these mission scenarios, the Earth return V is typically small enough that EP alone is sufficient. A crewed mission to the NEA Apophis in the year 2028 with an expendable BNTEP transfer vehicle is presented. Assembly operations, launch element masses, and other key characteristics of the vehicle are described. A comparison with a conventional NTR vehicle performing the same mission is also provided. Finally, reusability of the BNTEP transfer vehicle is explored.

SST Technology Follow-on Program - Phase I, Performance Evaluation of an SST Noise Suppressor Nozzle System. Volume 1. Suppressed Mode.

DTIC Science & Technology

ACOUSTIC INSULATION, *TURBOJET EXHAUST NOZZLES, *JET ENGINE NOISE, REDUCTION, JET TRANSPORT AIRCRAFT, THRUST AUGMENTATION , SUPERSONIC NOZZLES, DUCT...INLETS, CONVERGENT DIVERGENT NOZZLES, SUBSONIC FLOW, SUPERSONIC FLOW, SUPPRESSORS, TURBOJET INLETS, BAFFLES, JET PUMPS, THRUST , DRAG, TEMPERATURE

Parametric scramjet analysis

NASA Astrophysics Data System (ADS)

Choi, Jongseong

The performance of a hypersonic flight vehicle will depend on existing materials and fuels; this work presents the performance of the ideal scramjet engine for three different combustion chamber materials and three different candidate fuels. Engine performance is explored by parametric cycle analysis for the ideal scramjet as a function of material maximum service temperature and the lower heating value of jet engine fuels. The thermodynamic analysis is based on the Brayton cycle as similarly employed in describing the performance of the ramjet, turbojet, and fanjet ideal engines. The objective of this work is to explore material operating temperatures and fuel possibilities for the combustion chamber of a scramjet propulsion system to show how they relate to scramjet performance and the seven scramjet engine parameters: specific thrust, fuel-to-air ratio, thrust-specific fuel consumption, thermal efficiency, propulsive efficiency, overall efficiency, and thrust flux. The information presented in this work has not been done by others in the scientific literature. This work yields simple algebraic equations for scramjet performance which are similar to that of the ideal ramjet, ideal turbojet and ideal turbofan engines.

Thrust Stand for Electric Propulsion Performance Evaluation

NASA Technical Reports Server (NTRS)

Polzin, Kurt A.; Markusic, Thomas E.; Stanojev, Boris J.; Dehoyos, Amado; Spaun, Benjamin

2006-01-01

An electric propulsion thrust stand capable of supporting testing of thrusters having a total mass of up to 125 kg and producing thrust levels between 100 microN to 1 N has been developed and tested. The design features a conventional hanging pendulum arm attached to a balance mechanism that converts horizontal deflections produced by the operating thruster into amplified vertical motion of a secondary arm. The level of amplification is changed through adjustment of the location of one of the pivot points linking the system. Response of the system depends on the relative magnitudes of the restoring moments applied by the displaced thruster mass and the twisting torsional pivots connecting the members of the balance mechanism. Displacement is measured using a non-contact, optical linear gap displacement transducer and balance oscillatory motion is attenuated using a passive, eddy-current damper. The thrust stand employs an automated leveling and thermal control system. Pools of liquid gallium are used to deliver power to the thruster without using solid wire connections, which can exert undesirable time-varying forces on the balance. These systems serve to eliminate sources of zero-drift that can occur as the stand thermally or mechanically shifts during the course of an experiment. An in-situ calibration rig allows for steady-state calibration before, during and after thruster operation. Thrust measurements were carried out on a cylindrical Hall thruster that produces mN-level thrust. The measurements were very repeatable, producing results that compare favorably with previously published performance data, but with considerably smaller uncertainty.

Preliminary Results of Performance Measurements on a Cylindrical Hall-Effect Thruster with Magnetic Field Generated by Permanent Magnets

NASA Technical Reports Server (NTRS)

Polzin, K. A.; Raitses, Y.; Merino, E.; Fisch, N. J.

2008-01-01

The performance of a low-power cylindrical Hall thruster, which more readily lends itself to miniaturization and low-power operation than a conventional (annular) Hall thruster, was measured using a planar plasma probe and a thrust stand. The field in the cylindrical thruster was produced using permanent magnets, promising a power reduction over previous cylindrical thruster iterations that employed electromagnets to generate the required magnetic field topology. Two sets of ring-shaped permanent magnets are used, and two different field configurations can be produced by reorienting the poles of one magnet relative to the other. A plasma probe measuring ion flux in the plume is used to estimate the current utilization for the two magnetic configurations. The measurements indicate that electron transport is impeded much more effectively in one configuration, implying a higher thrust efficiency. Preliminary thruster performance measurements on this configuration were obtained over a power range of 100-250 W. The thrust levels over this power range were 3.5-6.5 mN, with anode efficiencies and specific impulses spanning 14-19% and 875- 1425 s, respectively. The magnetic field in the thruster was lower for the thrust measurements than the plasma probe measurements due to heating and weakening of the permanent magnets, reducing the maximum field strength from 2 kG to roughly 750-800 G. The discharge current levels observed during thrust stand testing were anomalously high compared to those levels measured in previous experiments with this thruster.

Static stress change from the 8 October, 2005 M = 7.6 Kashmir earthquake

USGS Publications Warehouse

Parsons, T.; Yeats, R.S.; Yagi, Y.; Hussain, A.

2006-01-01

We calculated static stress changes from the devastating M = 7.6 earthquake that shook Kashmir on 8 October, 2005. We mapped Coulomb stress change on target fault planes oriented by assuming a regional compressional stress regime with greatest principal stress directed orthogonally to the mainshock strike. We tested calculation sensitivity by varying assumed stress orientations, target-fault friction, and depth. Our results showed no impact on the active Salt Range thrust southwest of the rupture. Active faults north of the Main Boundary thrust near Peshawar fall in a calculated stress-decreased zone, as does the Raikot fault zone to the northeast. We calculated increased stress near the rupture where most aftershocks occurred. The greatest increase to seismic hazard is in the Indus-Kohistan seismic zone near the Indus River northwest of the rupture termination, and southeast of the rupture termination near the Kashmir basin.

Design and test of a high power electromechanical actuator for thrust vector control

NASA Technical Reports Server (NTRS)

Cowan, J. R.; Myers, W. N.

1992-01-01

NASA-Marshall is involved in the development of electromechanical actuators (EMA) for thrust-vector control (TVC) system testing and implementation in spacecraft control/gimballing systems, with a view to the replacement of hydraulic hardware. TVC system control is furnished by solid state controllers and power supplies; a pair of resolvers supply position feedback to the controller for precise positioning. Performance comparisons between EMA and hydraulic TVC systems are performed.

Low-thrust chemical propulsion system pump technology

NASA Technical Reports Server (NTRS)

Meadville, J. W.

1980-01-01

A study was conducted within the thrust range 450 to 9000 N (100 to 2000 pounds). Performance analyses were made on centrifugal, pitot, Barske, drag, Tesla, gear, piston, lobe, and vane pumps with liquid hydrogen, liquid methane, and liquid oxygen as propellants. Gaseous methane and hydrogen driven axial impulse turbines, vane expanders, piston expanders, and electric motors were studied as drivers. Data are presented on performance, sizes, weights, and estimated service lives and costs.

Design and test of a high power electromechanical actuator for thrust vector control

NASA Astrophysics Data System (ADS)

Cowan, J. R.; Myers, W. N.

1992-07-01

NASA-Marshall is involved in the development of electromechanical actuators (EMA) for thrust-vector control (TVC) system testing and implementation in spacecraft control/gimballing systems, with a view to the replacement of hydraulic hardware. TVC system control is furnished by solid state controllers and power supplies; a pair of resolvers supply position feedback to the controller for precise positioning. Performance comparisons between EMA and hydraulic TVC systems are performed.

Extended Performance 8-cm Mercury Ion Thruster

NASA Technical Reports Server (NTRS)

Mantenieks, M. A.

1981-01-01

A slightly modified 8-cm Hg ion thruster demonstrated significant increase in performance. Thrust was increased by almost a factor of five over that of the baseline thruster. Thruster operation with various three grid ion optics configurations; thruster performance as a function of accelerator grid open area, cathode baffle, and cathode orifice size; and a life test of 614 hours at a beam current of 250 mA (17.5 mN thrust) are discussed. Highest thruster efficiency was obtained with the smallest open area accelerator grid. The benefits in efficiency from the low neutral loss grids were mitigated, however, by the limitation such grids place on attainable ion beam current densities. The thruster components suffered negligible weight losses during a life test, which indicated that operation of the 8-cm thruster at extended levels of thrust and power is possible with no significant loss of lifetime.

Study to adapt solar electric propulsion to the Pioneer F and G spacecraft

NASA Technical Reports Server (NTRS)

1972-01-01

The addition of an electric thrust subsystem to the spin-stabilized Pioneer F and G spacecraft to improve performance capability for certain missions is discussed. The evaluation was performed for the Atlas and Titan launch vehicles with Centaur and TE-364-4 stages and for electric thrust stages of 8- and 5-kw with three 30- and five 15-cm thrusters respectively. The combination of a spinning spacecraft with electric propulsion is a concept only recently evaluated and the penalty from spinning over three-axis stabilized is not as significant as might initally be thought. There are major gains in weight, cost, and reliability, the disadvantages being lower data rate during the thrust phase and less efficient pointing. A variety of missions were evaluated from a solar approach mission into 0.14 AU to a flyby mission of Neptune at approximately 30 AU. Performance improvements were present for all missions evaluated.

Scale model test results of several STOVL ventral nozzle concepts

NASA Technical Reports Server (NTRS)

Meyer, B. E.; Re, R. J.; Yetter, J. A.

1991-01-01

Short take-off and vertical landing (STOVL) ventral nozzle concepts are investigated by means of a static cold flow scale model at a NASA facility. The internal aerodynamic performance characteristics of the cruise, transition, and vertical lift modes are considered for four ventral nozzle types. The nozzle configurations examined include those with: butterfly-type inner doors and vectoring exit vanes; circumferential inner doors and thrust vectoring vanes; a three-port segmented version with circumferential inner doors; and a two-port segmented version with cylindrical nozzle exit shells. During the testing, internal and external pressure is measured, and the thrust and flow coefficients and resultant vector angles are obtained. The inner door used for ventral nozzle flow control is found to affect performance negatively during the initial phase of transition. The best thrust performance is demonstrated by the two-port segmented ventral nozzle due to the elimination of the inner door.

Innovative Double Bypass Engine for Increased Performance

NASA Astrophysics Data System (ADS)

Manoharan, Sanjivan

Engines continue to grow in size to meet the current thrust requirements of the civil aerospace industry. Large engines pose significant transportation problems and require them to be split in order to be shipped. Thus, large amounts of time have been spent in researching methods to increase thrust capabilities while maintaining a reasonable engine size. Unfortunately, much of this research has been focused on increasing the performance and efficiencies of individual components while limited research has been done on innovative engine configurations. This thesis focuses on an innovative engine configuration, the High Double Bypass Engine, aimed at increasing fuel efficiency and thrust while maintaining a competitive fan diameter and engine length. The 1-D analysis was done in Excel and then compared to the results from Numerical Propulsion Simulation System (NPSS) software and were found to be within 4% error. Flow performance characteristics were also determined and validated against their criteria.

Simulator evaluation of the effects of reduced spoiler and thrust authority on a decoupled longitudinal control system during landings in wind shear

NASA Technical Reports Server (NTRS)

Miller, G. K., Jr.

1981-01-01

The effect of reduced control authority, both in symmetric spoiler travel and thrust level, on the effectiveness of a decoupled longitudinal control system was examined during the approach and landing of the NASA terminal configured vehicle (TCV) aft flight deck simulator in the presence of wind shear. The evaluation was conducted in a fixed-base simulator that represented the TCV aft cockpit. There were no statistically significant effects of reduced spoiler and thrust authority on pilot performance during approach and landing. Increased wind severity degraded approach and landing performance by an amount that was often significant. However, every attempted landing was completed safely regardless of the wind severity. There were statistically significant differences in performance between subjects, but the differences were generally restricted to the control wheel and control-column activity during the approach.

Nuclear-Rocket Propulsion

NASA Technical Reports Server (NTRS)

Rom, Frank E.

1968-01-01

The three basic types of nuclear power-plants (solid, liquid, and gas core) are compared on the bases of performance potential and the status of current technology. The solid-core systems are expected to have impulses in the range of 850 seconds, any thrust level (as long as it is greater than 10,000 pounds (44,480 newtons)), and thrust-to-engine-weight ratios of 2 to 20 pounds per pound (19.7 to 197 newtons per kilogram). There is negligible or no fuel loss from the solid-core system. The solid-core system, of course, has had the most work done on it. Large-scale tests have been performed on a breadboard engine that has produced specific impulses greater than 700 seconds at thrust levels of about 50,000 pounds (222,000 newtons). The liquid-core reactor would be interesting in the specific impulse range of 1200 to 1500 seconds. Again, any thrust level can be obtained depending on how big or small the reactor is made. The thrust-to-engine weight ratio for these systems would be in the range of 1 to 10. The discouraging feature of the liquid-core system is the high fuel-loss ratio anticipated. Values of 0.01 to 0.1 pound (0.00454 to 0.0454 kilograms) or uranium loss per pound (0.454 kilograms) of hydrogen are expected, if impulses in the range of 1200 to 1500 seconds are desired. The gas-core reactor shows specific impulses in the range of 1500 to 2500 seconds. The thrust levels should be at least as high as the weight so that the thrust-to-weight ratio does not go below 1. Because the engine weight is not expected to be under 100,000 pounds (444,800 newtons), thrust levels higher than 100,000 pounds (448,000 newtons) are of interest. The thrust-to-engine weights, in that case, would run from 1 to 20 pounds per pound (9.8 to 19.7 kilograms). Gas-core reactors tend to be very large, and can have high thrust-to-weight ratios. As in the case of the liquid-core system, the fuel loss that will be attendant with gas cores as envisioned today will be rather high. The loss rates will be 0.01 to 0.1 pound of uranium (0.00454 to 0.0454 kilograms) for each pound (0.454 kilograms) of hydrogen.

Non-invasive 3D geometry extraction and robotic modeling of a Sea lion foreflipper

NASA Astrophysics Data System (ADS)

Patel, R. K.; Leftwich, M. C.; Friedman, C.

2016-02-01

California Sea Lions are very agile swimmers and unlike many marine animals, they use their fore flipper rather than their hind flipper undulations to generate high thrust values. To date there exist limited amount of qualitative studies for sea lions swimming that show the flippers are used for thrust, stability, and control during swimming motions. Quantitative studies mainly measured drag used for cost of transport, and analyzed banked turn performance. Recently, the kinematics of a California sea lion flipper during the thrust phase was extracted using video tracking in two dimensions. This work extends the tracking ability to three dimensions using a non-invasive Direct Linear Transformation (DLT) technique employed on non-research sea lions at the Smithsonian National Zoological Park. The flippers are therefore marker-less and tracking is carried out manually in post processing after capturing complete dorsal-ventral flipper motions. Two cameras are used (3840 × 2160 pixels resolution) and calibrated in space using a calibration target inserted into the sea lion habitat. They are synchronized in time using a simple light flash. The fluid flow and forces generated by a sea lion clap is also being explored. Recently, a sea lion flipper from a deceased subject was externally scanned in high detail for fluid dynamics research. The flipper's geometry is being used in this work to design and build an articulate flipper model that is approximately 60% of the full size span. The model is actuated by servo motors and is designed to mimic a sea lion flipper clap motion based on the previously extracted kinematics from above. The model incorporates three axles, simulating the movements of the sea lion's elbow, wrist, and knuckles. The flipper tip speed is designed to match typical Reynolds numbers for the full-scale flipper for an acceleration from rest maneuver. The model will be tested in a water flume to obtain the forces during the thrust production phase of the flipper motion.

Air Force electrochemical power research and technology program for space applications

NASA Technical Reports Server (NTRS)

Allen, Douglas

1987-01-01

An overview is presented of the existing Air Force electrochemical power, battery, and fuel cell programs for space application. Present thrusts are described along with anticipated technology availability dates. Critical problems to be solved before system applications occur are highlighted. Areas of needed performance improvement of batteries and fuel cells presently used are outlined including target dates for key demonstrations of advanced technology. Anticipated performance and current schedules for present technology programs are reviewed. Programs that support conventional military satellite power systems and special high power applications are reviewed. Battery types include bipolar lead-acid, nickel-cadmium, silver-zinc, nickel-hydrogen, sodium-sulfur, and some candidate advanced couples. Fuel cells for pulsed and transportation power applications are discussed as are some candidate advanced regenerative concepts.

Orbit transfer vehicle engine study. Volume 2: Technical report

NASA Technical Reports Server (NTRS)

1980-01-01

The orbit transfer vehicle (OTV) engine study provided parametric performance, engine programmatic, and cost data on the complete propulsive spectrum that is available for a variety of high energy, space maneuvering missions. Candidate OTV engines from the near term RL 10 (and its derivatives) to advanced high performance expander and staged combustion cycle engines were examined. The RL 10/RL 10 derivative performance, cost and schedule data were updated and provisions defined which would be necessary to accommodate extended low thrust operation. Parametric performance, weight, envelope, and cost data were generated for advanced expander and staged combustion OTV engine concepts. A prepoint design study was conducted to optimize thrust chamber geometry and cooling, engine cycle variations, and controls for an advanced expander engine. Operation at low thrust was defined for the advanced expander engine and the feasibility and design impact of kitting was investigated. An analysis of crew safety and mission reliability was conducted for both the staged combustion and advanced expander OTV engine candidates.

77 FR 67309 - Special Conditions: Embraer S.A., Model EMB-550 Airplanes; Flight Envelope Protection...

Federal Register 2010, 2011, 2012, 2013, 2014

2012-11-09

... Envelope Protection: Performance Credit for Automatic Takeoff Thrust Control System (ATTCS) During Go... Automatic Takeoff Thrust Control System (ATTCS) during go-around. The applicable airworthiness regulations... FAA-2012-1199 using any of the following methods: Federal eRegulations Portal: Go to http://www...

Human Centered Modeling and Simulation

Science.gov Websites

Contacts Researchers Thrust Area 2: Human Centered Modeling and Simulation Thrust Area Leader: Dr. Matthew performance of human occupants and operators are paramount in the achievement of ground vehicle design objectives, but these occupants are also the most variable components of the human-machine system. Modeling

Static thrust-vectoring performance of nonaxisymmetric convergent-divergent nozzles with post-exit yaw vanes. M.S. Thesis - George Washington Univ., Aug. 1988

NASA Technical Reports Server (NTRS)

Foley, Robert J.; Pendergraft, Odis C., Jr.

1991-01-01

A static (wind-off) test was conducted in the Static Test Facility of the 16-ft transonic tunnel to determine the performance and turning effectiveness of post-exit yaw vanes installed on two-dimensional convergent-divergent nozzles. One nozzle design that was previously tested was used as a baseline, simulating dry power and afterburning power nozzles at both 0 and 20 degree pitch vectoring conditions. Vanes were installed on these four nozzle configurations to study the effects of vane deflection angle, longitudinal and lateral location, size, and camber. All vanes were hinged at the nozzle sidewall exit, and in addition, some were also hinged at the vane quarter chord (double-hinged). The vane concepts tested generally produced yaw thrust vectoring angles much less than the geometric vane angles, for (up to 8 percent) resultant thrust losses. When the nozzles were pitch vectored, yawing effectiveness decreased as the vanes were moved downstream. Thrust penalties and yawing effectiveness both decreased rapidly as the vanes were moved outboard (laterally). Vane length and height changes increased yawing effectiveness and thrust ratio losses, while using vane camber, and double-hinged vanes increased resultant yaw angles by 50 to 100 percent.

Numerical analysis of the static performance of an annular aerostatic gas thrust bearing applied in the cryogenic turbo-expander of the EAST subsystem

NASA Astrophysics Data System (ADS)

Lai, Tianwei; Fu, Bao; Chen, Shuangtao; Zhang, Qiyong; Hou, Yu

2017-02-01

The EAST superconducting tokamak, an advanced steady-state plasma physics experimental device, has been built at the Institute of Plasma Physics, Chinese Academy of Sciences. All the toroidal field magnets and poloidal field magnets, made of NbTi/Cu cable-in-conduit conductor, are cooled with forced flow supercritical helium at 3.8 K. The cryogenic system of EAST consists of a 2 kW/4 K helium refrigerator and a helium distribution system for the cooling of coils, structures, thermal shields, bus-lines, etc. The high-speed turbo-expander is an important refrigerating component of the EAST cryogenic system. In the turbo-expander, the axial supporting technology is critical for the smooth operation of the rotor bearing system. In this paper, hydrostatic thrust bearings are designed based on the axial load of the turbo-expander. Thereafter, a computational fluid dynamics-based numerical model of the aerostatic thrust bearing is set up to evaluate the bearing performance. Tilting effect on the pressure distribution and bearing load is analyzed for the thrust bearing. Bearing load and stiffness are compared with different static supply pressures. The net force from the thrust bearings can be calculated for different combinations of bearing clearance and supply pressure.

Performance of Several Conical Convergent-Divergent Rocket-Type Exhaust Nozzles

NASA Technical Reports Server (NTRS)

Campbell, C. E.; Farley, J. M.

1960-01-01

An investigation was conducted to obtain nozzle performance data with relatively large-scale models at pressure ratios as high as 120. Conical convergent-divergent nozzles with divergence angles alpha of 15, 25, and 29 deg. were each tested at area ratios of approximately 10, 25, and 40. Heated air (1200 F) was supplied at the nozzle inlet at pressures up to 145 pounds per square inch absolute and was exhausted into quiescent air at pressures as low as 1.2 pounds per square inch absolute. Thrust ratios for all nozzle configurations are presented over the range of pressure ratios attainable and were extrapolated when possible to design pressure ratio and beyond. Design thrust ratios decreased with increasing nozzle divergence angle according to the trend predicted by the (1 + cos alpha)/2 parameter. Decreasing the nozzle divergence angle resulted in sizable increases in thrust ratio for a given surface-area ratio (nozzle weight), particularly at low nozzle pressure ratios. Correlations of the nozzle static pressure at separation and of the average static pressure downstream of separation with various nozzle parameters permitted the calculation of thrust in the separated-flow region from unseparated static-pressure distributions. Thrust ratios calculated by this method agreed with measured values within about 1 percent.

Performance of gas-lubricated nonconforming pivoted-pad journal bearings and a flexibly mounted spiral-groove thrust bearing

NASA Technical Reports Server (NTRS)

Ream, L. W.

1973-01-01

A test program was conducted to determine the performance characteristics of gas-lubricated nonconforming pivoted-pad journal bearings and a spiral-groove thrust bearing designed for the Brayton cycle rotating unit (BRU). Hydrostatic, hybrid (simultaneously hydrostatic and hydrodynamic), and hydrodynamic tests were conducted in argon gas at ambient pressure and temperature ranges representative of hydrostatic operation up to the 10.5-kWe BRU power-generating level. Performance of the gas lubricated bearings is presented, including hydrostatic gas flow rates, bearing clearances, bearing temperatures, and transient performance.

Performance of gas-lubricated cruciform-mounted tilting-pad journal bearings and a damped flexibly mounted spiral-groove thrust bearing

NASA Technical Reports Server (NTRS)

Ream, L. W.

1974-01-01

A test program was conducted to determine the performance characteristics of gas-lubricated cruciform-mounted tilting-pad journal bearings and a damped spiral-groove thrust bearing designed for the Brayton cycle rotating unit (BRU). Hydrostatic, hybrid (simultaneously hydrostatic and hydrodynamic), and hydrodynamic tests were conducted in argon gas at ambient pressure and temperature ranges representative of operation to the 10.5 kWe BRU power-generating level. Performance of the gas lubricated bearings is presented including hydrostatic gas flow rates, bearing clearances, bearing temperatures, and transient performance.

Isolated Performance at Mach Numbers From 0.60 to 2.86 of Several Expendable Nozzle Concepts for Supersonic Applications

NASA Technical Reports Server (NTRS)

Re, Richard J.; Berrier, Bobby L.; Abeyounis, William K.

2001-01-01

Investigations have been conducted in the Langley 16-Foot Transonic Tunnel (at Mach numbers from 0.60 to 1.25) and in the Langley Unitary Plan Wind Tunnel (at Mach numbers from 2.16 to 2.86) at an angle of attack of 0 deg to determine the isolated performance of several expendable nozzle concepts for supersonic nonaugmented turbojet applications. The effects of centerbody base shape, shroud length, shroud ventilation, cruciform shroud expansion ratio, and cruciform shroud flap vectoring were investigated. The nozzle pressure ratio range, which was a function of Mach number, was between 1.9 and 11.8 in the 16-Foot Transonic Tunnel and between 7.9 and 54.9 in the Unitary Plan Wind Tunnel. Discharge coefficient, thrust-minus-drag, and the forces and moments generated by vectoring the divergent shroud flaps (for Mach numbers of 0.60 to 1.25 only) of a cruciform nozzle configuration were measured. The shortest nozzle had the best thrust-minus-drag performance at Mach numbers up to 0.95 but was approached in performance by other configurations at Mach numbers of 1.15 and 1.25. At Mach numbers above 1.25, the cruciform nozzle configuration having the same expansion ratio (2.64) as the fixed geometry nozzles had the best thrust-minus-drag performance. Ventilation of the fixed geometry divergent shrouds to the nozzle external boattail flow generally improved thrust-minus-drag performance at Mach numbers from 0.60 to 1.25, but decreased performance above a Mach number of 1.25.

Development of circulation control technology for powered-lift STOL aircraft

NASA Technical Reports Server (NTRS)

Englar, Robert J.

1987-01-01

The flow entraining capabilities of the Circulation Control Wing high lift system were employed to provide an even stronger STOL potential when synergistically combined with upper surface mounted engines. The resulting configurations generate very high supercirculation lift in addition to a vertical component of the pneumatically deflected engine thrust. A series of small scale wind tunnel tests and full scale static thrust deflection tests are discussed which provide a sufficient data base performance. These tests results show thrust deflections of greater than 90 deg produced pneumatically by nonmoving aerodynamic surfaces, and the ability to maintain constant high lift while varying the propulsive force from high thrust recovery required for short takeoff to high drag generation required for short low speed landings.

Investigation of Pneumatic Inlet and Diffuser Blowing on a Ducted Fan Propulsor in Static Thrust Operation

NASA Technical Reports Server (NTRS)

Kondor, Shayne; Englar, Robert J.; Lee, Warren J.

2003-01-01

Tilting ducted fans present a solution for the lifting and forward flight propulsion requirements of VTOL aircraft. However, the geometry of the duct enshrouding the propeller has great a effect on the efficiency of the fan in various flight modes. Shroud geometry controls the velocity and pressure at the face of the fan, while maintaining a finite loading out at the tips of the fan blades. A duct tailored for most efficient generation of static lifting thrust will generally suffer from performance deficiencies in forward flight. The converse is true as well, leaving the designer with a difficult trade affecting the overall performance and sizing of the aircraft. Ideally, the shroud of a vertical lifting fan features a generous bell mouth inlet promoting acceleration of flow into the face of the fan, and terminating in a converging nozzle at the exit. Flow entering the inlet is accelerated into the fan by the circulation about the shroud, resulting in an overall increase in thrust compared to an open propeller operating under the same conditions . The accelerating shroud design is often employed in lifting ducted fans to benefit from the thrust augmentation; however, such shroud designs produce significant drag penalties in axial flight, thus are unsuitable for efficient forward flight applications. Decelerating, or diffusing, duct designs are employed for higher speed forward flight configurations. The lower circulation on the shroud tends to decelerate the flow into the face of the fan, which is detrimental to static thrust development; however, net thrust is developed on the shroud while the benefits of finite blade loading are retained. With judicious shroud design for intended flight speeds, a net increase in efficiency can be obtained over an open propeller. In this experiment, conducted under contract to NASA LaRC (contract NAG-1-02093) circulation control is being applied to a mildly diffusing shroud design, intended for improved forward flight performance, to generate circulation in the sense of an accelerating duct design. The intent is to improve static thrust performance of a ducted fan tailored for high speed axial flight, while at the same time significantly reduce the pressure signature on the ground plane. Circulation control on the fan shroud is achieved by the Coanda effect.

Performance and acoustic prediction of counterrotating propeller configurations

NASA Technical Reports Server (NTRS)

Denner, B. W.; Korkan, K. D.

1989-01-01

The Davidson (1981) numerical method is used to predict the performance of a counterrotating propeller configuration over a range of different front and back disk rotation speeds with constant-speed propellers; this has yielded such overall performance parameters as integrated thrust, torque, and power, as well as the radial variation of blade torque and thrust. Since the unsteady component of the noise from a counterrotating propeller configuration is minimal in the plane of the propeller disk, this approach is restricted to noise-level predictions for observer locations in this region.

Cranial tibial wedge osteotomy: a technique for eliminating cranial tibial thrust in cranial cruciate ligament repair.

PubMed

Slocum, B; Devine, T

1984-03-01

Cranial tibial wedge osteotomy, surgical technique for cranial cruciate ligament rupture, was performed on 19 stifles in dogs. This procedure leveled the tibial plateau, thus causing weight-bearing forces to be compressive and eliminating cranial tibial thrust. Without cranial tibial thrust, which was antagonistic to the cranial cruciate ligament and its surgical reconstruction, cruciate ligament repairs were allowed to heal without constant loads. This technique was meant to be used as an adjunct to other cranial cruciate ligament repair techniques.

Effects of an in-flight thrust reverser on the stability and control characteristics of a single-engine fighter airplane model

NASA Technical Reports Server (NTRS)

Mercer, C. E.; Maiden, D. L.

1972-01-01

The changes in thrust minus drag performance as well as longitudinal and directional stability and control characteristics of a single-engine jet aircraft attributable to an in-flight thrust reverser of the blocker-deflector door type were investigated in a 16-foot transonic wind tunnel. The longitudinal and directional stability data are presented. Test conditions simulated landing approach conditions as well as high speed maneuvering such as may be required for combat or steep descent from high altitude.

A comparative study of scramjet injection strategies for high Mach numbers flows

NASA Technical Reports Server (NTRS)

Riggins, D. W.; Mcclinton, C. R.; Rogers, R. C.; Bittner, R. D.

1992-01-01

A simple method for predicting the axial distribution of supersonic combustor thrust potential is described. A complementary technique for illustrating the spatial evolution and distribution of thrust potential and loss mechanisms in reacting flows is developed. Wall jet cases and swept ramp injector cases for Mach 17 and Mach 13.5 flight enthalpy inflow conditions are numerically modeled and analyzed using these techniques. The visualization of thrust potential in the combustor for the various cases examined provides a unique tool for increasing understanding of supersonic combustor performance potential.

An Experiment on Repetitive Pulse Operation of Microwave Rocket

DOE Office of Scientific and Technical Information (OSTI.GOV)

Oda, Yasuhisa; Shibata, Teppei; Komurasaki, Kimiya

2008-04-28

Microwave Rocket was operated with repetitive pulses. The microwave rocket model with forced breathing system was used. The pressure history in the thruster was measured and the thrust impulse was deduced. As a result, the impulse decreased at second pulse and impulses at latter pulses were constant. The dependence of the thrust performance on the partial filling rate of the thruster was compared to the thrust generation model based on the shock wave driven by microwave plasma. The experimental results showed good agreement to the predicted dependency.

Performance and environmental impact assessment of pulse detonation based engine systems

NASA Astrophysics Data System (ADS)

Glaser, Aaron J.

Experimental research was performed to investigate the feasibility of using pulse detonation based engine systems for practical aerospace applications. In order to carry out this work a new pulse detonation combustion research facility was developed at the University of Cincinnati. This research covered two broad areas of application interest. The first area is pure PDE applications where the detonation tube is used to generate an impulsive thrust directly. The second focus area is on pulse detonation based hybrid propulsion systems. Within each of these areas various studies were performed to quantify engine performance. Comparisons of the performance between detonation and conventional deflagration based engine cycles were made. Fundamental studies investigating detonation physics and flow dynamics were performed in order to gain physical insight into the observed performance trends. Experimental studies were performed on PDE-driven straight and diverging ejectors to determine the system performance. Ejector performance was quantified by thrust measurements made using a damped thrust stand. The effects of PDE operating parameters and ejector geometric parameters on thrust augmentation were investigated. For all cases tested, the maximum thrust augmentation is found to occur at a downstream ejector placement. The optimum ejector geometry was determined to have an overall length of LEJECT/DEJECT =5.61, including an intermediate-straight section length of LSTRT /DEJECT=2, and diverging exhaust section with 4 deg half-angle. A maximum thrust augmentation of 105% was observed while employing the optimized ejector geometry and operating the PDE at a fill-fraction of 0.6 and a frequency of 10 Hz. When operated at a fill-fraction of 1.0 and a frequency of 30 Hz, the thrust augmentation of the optimized PDE-driven ejector system was observed to be 71%. Static pressure was measured along the interior surface of the ejector, including the inlet and exhaust sections. The diverging ejector pressure distribution shows that the diverging section acts as a subsonic diffuser. To provide a better explanation of the observed performance trends, shadowgraph images of the detonation wave and starting vortex interacting with the ejector inlet were obtained. The acoustic signature of a pulse detonation engine was characterized in both the near-field and far-field regimes. Experimental measurements were performed in an anechoic test facility designed for jet noise testing. Both shock strength and speed were mapped as a function of radial distance and direction from the PDE exhaust plane. It was found that the PDE generated pressure field can be reasonably modeled by a theoretical point-source explosion. The effect of several exit nozzle configurations on the PDE acoustic signature was studies. These included various chevron nozzles, a perforated nozzle, and a set of proprietary noise attenuation mufflers. Experimental studies were carried out to investigate the performance of a hybrid propulsion system integrating an axial flow turbine with multiple pulse detonation combustors. The integrated system consisted of a circular array of six pulse detonation combustor (PDC) tubes exhausting through an axial flow turbine. Turbine component performance was quantified by measuring the amount of power generated by the turbine section. Direct comparisons of specific power output and turbine efficiency between a PDC-driven turbine and a turbine driven by steady-flow combustors were made. It was found that the PDC-driven turbine had comparable performance to that of a steady-burner-driven turbine across the operating map of the turbine.

A proposed definition for a pitch attitude target for the microburst escape maneuver

NASA Technical Reports Server (NTRS)

Bray, Richard S.

1990-01-01

The Windshear Training Aid promulgated by the Federal Aviation Administration (FAA) defines the practical recovery maneuver following a microburst encounter as application of maximum thrust accompanied by rotation to an aircraft-specific target pitch attitude. In search of a simple method of determining this target, appropriate to a variety of aircraft types, a computer simulation was used to explore the suitability of a pitch target equal in numerical value to that of the angle of attack associated with stall warning. For the configurations and critical microburst shears simulated, this pitch target was demonstrated to be close to optimum.

Deorbit targeting

NASA Technical Reports Server (NTRS)

Tempelman, W. H.

1973-01-01

The navigation and control of the space shuttle during atmospheric entry are discussed. A functional flow diagram presenting the basic approach to the deorbit targeting problem is presented. The major inputs to be considered are: (1) vehicle state vector, (2) landing site location, (3) entry interface parameters, (4) earliest desired time of landing, and (5) maximum cross range. Mathematical models of the navigational procedures based on controlled thrust times are developed.

Simple Sensitivity Analysis for Orion GNC

NASA Technical Reports Server (NTRS)

Pressburger, Tom; Hoelscher, Brian; Martin, Rodney; Sricharan, Kumar

2013-01-01

The performance of Orion flight software, especially its GNC software, is being analyzed by running Monte Carlo simulations of Orion spacecraft flights. The simulated performance is analyzed for conformance with flight requirements, expressed as performance constraints. Flight requirements include guidance (e.g. touchdown distance from target) and control (e.g., control saturation) as well as performance (e.g., heat load constraints). The Monte Carlo simulations disperse hundreds of simulation input variables, for everything from mass properties to date of launch.We describe in this paper a sensitivity analysis tool (Critical Factors Tool or CFT) developed to find the input variables or pairs of variables which by themselves significantly influence satisfaction of requirements or significantly affect key performance metrics (e.g., touchdown distance from target). Knowing these factors can inform robustness analysis, can inform where engineering resources are most needed, and could even affect operations. The contributions of this paper include the introduction of novel sensitivity measures, such as estimating success probability, and a technique for determining whether pairs of factors are interacting dependently or independently. The tool found that input variables such as moments, mass, thrust dispersions, and date of launch were found to be significant factors for success of various requirements. Examples are shown in this paper as well as a summary and physics discussion of EFT-1 driving factors that the tool found.

Piloted Simulation Study of a Dual Thrust-Cutback Procedure for Reducing High-Speed Civil Transport Takeoff Noise Levels

NASA Technical Reports Server (NTRS)

Riley, Donald R.; Glaab, Louis J.; Brandon, Jay M.; Person, Lee H., Jr.; Glaab, Patricia C.

1999-01-01

A piloted simulation study was performed for the purpose of indicating the noise reduction benefits and piloting performance that could occur for a typical 4-engine high-Speed Civil Transport (HSCT) configuration during takeoff when a dual thrust-cutback procedure was employed with throttle operation under direct computer control. Two thrust cutbacks were employed with the first cutback performed while the vehicle was accelerating on the run-way and the second cutback performed at a distance farther downrange. Added vehicle performance improvements included the incorporation of high-lift increments into the aerodynamic database of the vehicle and the use of limited engine oversizing. Four single-stream turbine bypass engines that had no noise suppression of any kind were used with this configuration. This approach permitted establishing the additional noise suppression level that was needed to meet Federal Air Regulation Part 36 Stage 3 noise levels for subsonic commercial jet aircraft. Noise level results were calculated with the jet mixing and shock noise modules of the Aircraft Noise Prediction Program (ANOPP).

Full-scale hingeless rotor performance and loads

NASA Technical Reports Server (NTRS)

Peterson, Randall L.

1995-01-01

A full-scale BO-105 hingeless rotor system was tested in the NASA Ames 40- by 80-Foot Wind Tunnel on the rotor test apparatus. Rotor performance, rotor loads, and aeroelastic stability as functions of both collective and cyclic pitch, tunnel velocity, and shaft angle were investigated. This test was performed in support of the Rotor Data Correlation Task under the U.S. Army/German Memorandum of Understanding on Cooperative Research in the Field of Helicopter Aeromechanics. The primary objective of this test program was to create a data base for full-scale hingeless rotor performance and structural blade loads. A secondary objective was to investigate the ability to match flight test conditions in the wind tunnel. This data base can be used for the experimental and analytical studies of hingeless rotor systems over large variations in rotor thrust and tunnel velocity. Rotor performance and structural loads for tunnel velocities from hover to 170 knots and thrust coefficients (C(sub T)/sigma) from 0.0 to 0.12 are presented in this report. Thrust sweeps at tunnel velocities of 10, 20, and 30 knots are also included in this data set.

Development of a Transient Thrust Stand with Sub-Millisecond Resolution

NASA Astrophysics Data System (ADS)

Spells, Corbin Fraser

The transient thrust stand has been developed to offer 0.1 ms time resolved thrust measurements for the characterization of mono-propellant thrusters for spacecraft applications. Results demonstrated that the system was capable of obtaining dynamic thrust profiles within 5 % and 0.1 ms. Measuring and improving the thrust performance of mono-propellant thrusters will require 1 ms time resolved forces to observe shot-to-shot variations, oscillations, and minimum impulse bits. To date, no thrust stand is capable of measuring up to 22 N forces with a time response of up to 10 kHz. Calibration forces up to 22 N with a frequency response greater than 0.1 ms were obtained using voice coil actuators. Steady state and low frequency measurements were obtained using displacement and velocity sensors and were combined with high frequency vibration modes measured using several accelerometers along the thrust stand arm. The system uses a predictor-based subspace algorithm to obtain a high order state space model of the thrust stand capable of defining the high frequency vibration modes. The high frequency vibration modes are necessary to provide the time response of 0.1 ms. Thruster forces are estimated using an augmented Kalman filter to combine sensor traces from four accelerometers, a velocity sensor, and displacement transducer. Combining low frequency displacement data with high frequency acceleration measurements provides accurate force data across a broad time domain. The transient thrust stand uses a torsional pendulum configuration to minimize influence from external vibration and achieve high force resolution independent of thruster weight.

Thrust Measurement of Dielectric Barrier Discharge (DBD) Plasma Actuators

NASA Astrophysics Data System (ADS)

Ashpis, David E.; Laun, Matthew C.

2013-11-01

DBD plasma actuators generate a wall-jet that can be used for active flow control. We used an analytical balance to measure the thrust generated by the actuator, it is a common metric of its performance without external flow. We found that the measured force is afflicted by several problems; it drifts in time, not always repeatable, is unstable, and depends on the manner the voltage is applied. We report results of investigations of these issues. Tests were conducted on an actuator constructed of 1/4 inch thick high-density polyethylene (HDPE) dielectric with 100 mm long offset electrodes, with applied voltages up to 48 kV p-p and frequencies from 32 Hz to 2.5 kHz, and pure Sine and Trapezoidal waveforms. The relative humidity was in the range of 51-55%, corresponding to moisture range of 10,500 to13,000 ppm mass. Force readings were up to 500 mg, (approximately 50 mN/m). We found that the measured force is the net of the positive thrust generated by the wall-jet and an ``anti-thrust'' acting in the opposite direction. We propose a correction procedure that yields the plasma-generated thrust. The correction is based on voltage-dependent anti-thrust measured in the low frequency range of 20-40 Hz. We found that adjacent objects in a test setup affect the measured thrust, and verified it by comparing experiments with and without a metal enclosure, grounded and ungrounded. Uncorrected thrust varied by up to approximately +/-100%, and the corrected thrust variations were up to approximately 30%. Supported by NASA's FAP/Aerospace Sciences Project.

A study of unmanned mission opportunities to comets and asteroids

NASA Technical Reports Server (NTRS)

Mann, F. I.; Horsewood, J. L.; Bjorkman, W.

1974-01-01

Several unmanned multiple-target mission opportunities to comets and asteroids were studied. The targets investigated include Grigg-Skjellerup, Giacobini-Zinner, Tuttle-Giacobini-Kresak, Borrelly, Halley, Schaumasse, Geographos, Eros, Icarus, and Toro, and the trajectories consist of purely ballistic flight, except that powered swingbys and deep space burns are employed when necessary. Optimum solar electric rendezvous trajectories to the comets Giacobini-Zinner/85, Borrelly/87, and Temple (2)/83 and /88 employing the 8.67 kw Sert III spacecraft modified for interplanetary flight were also investigated. The problem of optimizing electric propulsion heliocentric trajectories, including the effects of geocentric launch asymptote declination on launch vehicle performance capability, was formulated, and a solution developed using variational calculus techniques. Improvements were made to the HILTOP trajectory optimization computer program. An error analysis of high-thrust maneuvers involving spin-stabilized spacecraft was developed and applied to a synchronous meteorological satellite mission.

Performance seeking control program overview

NASA Technical Reports Server (NTRS)

Orme, John S.

1995-01-01

The Performance Seeking Control (PSC) program evolved from a series of integrated propulsion-flight control research programs flown at NASA Dryden Flight Research Center (DFRC) on an F-15. The first of these was the Digital Electronic Engine Control (DEEC) program and provided digital engine controls suitable for integration. The DEEC and digital electronic flight control system of the NASA F-15 were ideally suited for integrated controls research. The Advanced Engine Control System (ADECS) program proved that integrated engine and aircraft control could improve overall system performance. The objective of the PSC program was to advance the technology for a fully integrated propulsion flight control system. Whereas ADECS provided single variable control for an average engine, PSC controlled multiple propulsion system variables while adapting to the measured engine performance. PSC was developed as a model-based, adaptive control algorithm and included four optimization modes: minimum fuel flow at constant thrust, minimum turbine temperature at constant thrust, maximum thrust, and minimum thrust. Subsonic and supersonic flight testing were conducted at NASA Dryden covering the four PSC optimization modes and over the full throttle range. Flight testing of the PSC algorithm, conducted in a series of five flight test phases, has been concluded at NASA Dryden covering all four of the PSC optimization modes. Over a three year period and five flight test phases 72 research flights were conducted. The primary objective of flight testing was to exercise each PSC optimization mode and quantify the resulting performance improvements.

The effects of compressor seventh-stage bleed air extraction on performance of the F100-PW-220 afterburning turbofan engine

NASA Technical Reports Server (NTRS)

Evans, Alison B.

1991-01-01

A study was conducted to determine the effects of seventh-stage compressor bleed on the performance of the F100 afterburning turbofan engine. The effects of bleed on thrust, specific fuel consumption, fan turbine inlet temperature, bleed total pressure, and bleed total temperature were obtained from the engine manufacturer's status deck computer simulation. These effects were determined for power settings of intermediate, partial afterburning, and maximum afterburning for Mach numbers between 0.6 and 2.2 and for altitudes of 30,000, 40,000, and 50,000 ft. It was found that thrust loss and specific fuel consumption increase were approximately linear functions of bleed flow and, based on a percent-thrust change basis, were approximately independent of power setting.

In-Flight Operation of the Dawn Ion Propulsion System Through Survey Science Orbit at Ceres

NASA Technical Reports Server (NTRS)

Garner, Charles E.; Rayman, Marc D.

2015-01-01

The Dawn mission, part of NASA's Discovery Program, has as its goal the scientific exploration of the two most massive main-belt objects, Vesta and Ceres. The Dawn spacecraft was launched from the Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H- 9.5 (Delta-II Heavy) rocket that placed the 1218-kg spacecraft onto an Earth-escape trajectory. On-board the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory which will provide a total delta V of 11 km/s for the heliocentric transfer to Vesta, orbit capture at Vesta, transfer between Vesta science orbits, departure and escape from Vesta, heliocentric transfer to Ceres, orbit capture at Ceres, and transfer between Ceres science orbits. Full-power thrusting from December 2007 through October 2008 was used to successfully target a Mars gravity assist flyby in February 2009 that provided an additional delta V of 2.6 km/s. Deterministic thrusting for the heliocentric transfer to Vesta resumed in June 2009 and concluded with orbit capture at Vesta on July 16, 2011. From July 2011 through September 2012 the IPS was used to transfer to all the different science orbits at Vesta and to escape from Vesta orbit. Cruise for a rendezvous with Ceres began in September 2012 and concluded with the start of the approach to Ceres phase on December 26, 2015, leading to orbit capture on March 6, 2015. Deterministic thrusting continued during approach to place the spacecraft in its first science orbit, called RC3, which was achieved on April 23, 2015. Following science operations at RC3 ion thrusting was resumed for twenty-five days leading to arrival to the next science orbit, called survey orbit, on June 3, 2015. The IPS will be used for all subsequent orbit transfers and trajectory correction maneuvers until completion of the primary mission in approximately June 2016. To date the IPS has been operated for over 46,774 hours, consumed approximately 393 kg of xenon, and provided a delta V of over 10.8 km/s to the spacecraft. The IPS performance characteristics are very close to the expected performance based on analysis and testing performed pre-launch. This paper provides an overview of Dawn's mission objectives and the results of Dawn IPS mission operations through arrival at the second science orbit at Ceres.

DC-9 Flight Demonstration Program with Refanned JT8D Engines. Volume 3; Performance and Analysis

NASA Technical Reports Server (NTRS)

1975-01-01

The JT8D-109 engine has a sea level static, standard day bare engine takeoff thrust of 73,840 N. At sea level standard day conditions the additional thrust of the JT8D-109 results in 2,040 kg additional takeoff gross weight capability for a given field length. Range loss of the DC-9 Refan airplane for long range cruise was determined. The Refan airplane demonstrated stall, static longitudinal stability, longitudinal control, longitudinal trim, minimum control speeds, and directional control characteristics similar to the DC-9-30 production airplane and complied with airworthiness requirements. Cruise, climb, and thrust reverser performance were evaluated. Structural and dynamic ground test, flight test and analytical results substantiate Refan Program requirements that the nacelle, thrust reverser hardware, and the airplane structural modifications are flightworthy and certifiable and that the airplane meets flutter speed margins. Estimated unit cost of a DC-9 Refan retrofit program is 1.338 million in mid-1975 dollars with about an equal split in cost between airframe and engine.

A Simple Model of Pulsed Ejector Thrust Augmentation

NASA Technical Reports Server (NTRS)

Wilson, Jack; Deloof, Richard L. (Technical Monitor)

2003-01-01

A simple model of thrust augmentation from a pulsed source is described. In the model it is assumed that the flow into the ejector is quasi-steady, and can be calculated using potential flow techniques. The velocity of the flow is related to the speed of the starting vortex ring formed by the jet. The vortex ring properties are obtained from the slug model, knowing the jet diameter, speed and slug length. The model, when combined with experimental results, predicts an optimum ejector radius for thrust augmentation. Data on pulsed ejector performance for comparison with the model was obtained using a shrouded Hartmann-Sprenger tube as the pulsed jet source. A statistical experiment, in which ejector length, diameter, and nose radius were independent parameters, was performed at four different frequencies. These frequencies corresponded to four different slug length to diameter ratios, two below cut-off, and two above. Comparison of the model with the experimental data showed reasonable agreement. Maximum pulsed thrust augmentation is shown to occur for a pulsed source with slug length to diameter ratio equal to the cut-off value.

Identification of human-generated forces on wheelchairs during total-body extensor thrusts.

PubMed

Hong, Seong-Wook; Patrangenaru, Vlad; Singhose, William; Sprigle, Stephen

2006-10-01

Involuntary extensor thrust experienced by wheelchair users with neurological disorders may cause injuries via impact with the wheelchair, lead to the occupant sliding out of the seat, and also damage the wheelchair. The concept of a dynamic seat, which allows movement of a seat with respect to the wheelchair frame, has been suggested as a potential solution to provide greater freedom and safety. Knowledge of the human-generated motion and forces during unconstrained extensor thrust events is of great importance in developing more comfortable and effective dynamic seats. The objective of this study was to develop a method to identify human-generated motions and forces during extensor thrust events. This information can be used to design the triggering system for a dynamic seat. An experimental system was developed to automatically track the motions of the wheelchair user using a video camera and also measure the forces at the footrest. An inverse dynamic approach was employed along with a three-link human body model and the experimental data to predict the human-generated forces. Two kinds of experiments were performed: the first experiment validated the proposed model and the second experiment showed the effects of the extensor thrust speed, the footrest angle, and the seatback angle. The proposed method was tested using a sensitivity analysis, from which a performance index was deduced to help indicate the robust region of the force identification. A system to determine human-generated motions and forces during unconstrained extensor thrusts was developed. Through experiments and simulations, the effectiveness and reliability of the developed system was established.

Rapid deceleration mode evaluation

NASA Technical Reports Server (NTRS)

Conners, Timothy R.; Nobbs, Steven G.; Orme, John S.

1995-01-01

Aircraft with flight capability above 1.4 normally have an RPM lockup or similar feature to prevent inlet buzz that would occur at low engine airflows. This RPM lockup has the effect of holding the engine thrust level at the intermediate power (maximum non-afterburning). For aircraft such as military fighters or supersonic transports, the need exists to be able to rapidly slow from supersonic to subsonic speeds. For example, a supersonic transport that experiences a cabin decompression needs to be able to slow/descend rapidly, and this requirement may size the cabin environmental control system. For a fighter, there may be a desire to slow/descend rapidly, and while doing so to minimize fuel usage and engine exhaust temperature. Both of these needs can be aided by achieving the minimum possible overall net propulsive force. As the intermediate power thrust levels of engines increase, it becomes even more difficult to slow rapidly from supersonic speeds. Therefore, a mode of the performance seeking control (PSC) system to minimize overall propulsion system thrust has been developed and tested. The rapid deceleration mode reduces the engine airflow consistent with avoiding inlet buzz. The engine controls are trimmed to minimize the thrust produced by this reduced airflow, and moves the inlet geometry to degrade the inlet performance. As in the case of the other PSC modes, the best overall performance (in this case the least net propulsive force) requires an integrated optimization of inlet, engine, and nozzle variables. This paper presents the predicted and measured results for the supersonic minimum thrust mode, including the overall effects on aircraft deceleration.

Evaluation of fins used in underwater swimming.

PubMed

Pendergast, D R; Mollendorf, J; Logue, C; Samimy, S

2003-01-01

Underwater swimmers use fins which augment thrust to overcome drag and propel the diver. The VdotO2 of swimming as a function of speed, velocity as a function of kick frequency, maximal speed (v), maximal oxygen consumption (VdotO2) and the maximal thrust were determined for eight fins in 10 male divers swimming at 1.25 m depth in a 60 m annular pool. A theoretical analysis of fin cycles was also performed. VdotO2 increased as a second order polynomial as a function of velocity; VdotO2 = 0.045 + 1.65B V + 1.66 (2) V2 (r2 = 0.997), VdotO2 = 0.25 + 1.03 V + 1.83 V2 (r2 = 0.997) and VdotO2 = -0.15 + 2.26 V + 1.49 V2 (r2 = 0.997), for least, average and most economical fins respectively. Kick frequency increased linearly with velocity and had a unique movement path (signature), giving theoretical values that agreed with the measured thrust, drag and efficiency. In conclusion, virtually all thrust comes from the downward power stroke, with rigid fins kicked deep (high drag), while flexible fins are kicked less deep but with higher frequency (low efficiency). Kick depth and frequency explain the performance of the eight tested fins, and should be optimized to enhance diver performance.

A Computational Study to Investigate the Effect of Altitude on Deteriorated Engine Performance

NASA Astrophysics Data System (ADS)

Koh, W. C.; Mazlan, N. M.; Rajendran, P.; Ismail, M. A.

2018-05-01

This study presents an investigation on the effect of operational altitudes on the performance of the deteriorated engine. A two-spool high bypass ratio turbofan engine is used as the test subject for this study. The engine is modelled in Gas Turbine Simulation Program (GSP) based on an existing engine model from literature. Real flight data were used for the validation. Deterioration rate of 0.1% per day is applied for all turbofan components engine. The simulation is performed by varying the altitude from sea level until 9000m. Results obtained show reduction in air mass flow rate and engine thrust as altitude increases. The reduction in air mass flow rate is due to the lower air density at higher altitude hence reduces amount of engine thrust. At 1000m to 4000m, thrust specific fuel consumption (TSFC) of the engine is improved compared to sea level. However depleted in TSFC is shown when the aircraft flies at altitude higher than 4000m. At this altitude, the effect of air density is dominant. As a result, the engine is required to burn more fuel to provide a higher thrust to sustain the aircraft speed. More fuel is consumed hence depletion in TSFC is obtained.

Performance of active and passive control of an airfoil using CPFD

NASA Astrophysics Data System (ADS)

Asselin, Daniel; Young, Jay; Williamson, C. H. K.

2016-11-01

Birds and fish employ flapping motions of their wings and fins in order to produce thrust and maneuver in flight and underwater. There is considerable interest in designing aerial and submersible systems that mimic these motions for the purposes of surveillance, environmental monitoring, and search and rescue, among other applications. Flapping motions are typically composed of combined pitch and heave and can provide good thrust and efficiency (Read, et al. 2003). In this study, we examine the performance of an airfoil actuated only in the heave direction. Using a cyber-physical fluid dynamics system (Mackowski & Williamson 2011, 2015, 2016), we simulate the presence of a torsion spring to enable the airfoil to undergo a passively controlled pitching motion. The addition of passive pitching combined with active heaving ("Active-Passive" or AP) provides significantly improved thrust and efficiency compared with heaving alone. In many cases, values of thrust and efficiency are comparable to or better than those obtained with two actively controlled degrees of freedom ("Active-Active" or AA). By using carefully-designed passive dynamics in the pitch direction, we can eliminate one of the two actuators, saving cost, complexity, and weight, while maintaining or improving performance. This work was supported by the Air Force Office of Scientific Research Grant No. FA9550-15-1-0243, monitored by Dr. Douglas Smith.

Low Thrust, Deep Throttling, US/CIS Integrated NTRE

NASA Astrophysics Data System (ADS)

Culver, Donald W.; Kolganov, Vyacheslav; Rochow, Richard F.

1994-07-01

In 1993 our international team performed a follow-on ``Nuclear Thermal Rocket Engine (NTRE) Extended Life Feasibility Assessment'' study for the Nuclear Propulsion Office (NPO) at NASAs Lewis Research Center. The main purpose of this study was to complete the 1992 study matrix to assess NTRE designs at thrust levels of 22.5, 11.3, and 6.8 tonnes, using Commonwealth of Independent States (CIS) reactor technology. An additional Aerojet goal was to continue improving the NTRE concept we had generated. Deep throttling, mission performance optimized engine design parametrics, and reliability/cost enhancing engine system simplifications were studied, because they seem to be the last three basic design improvements sorely needed by post-NERVA NTRE. Deep throttling improves engine life by eliminating damaging thermal and mechanical shocks caused by after-cooling with pulsed coolant flow. Alternately, it improves mission performance with steady flow after-cooling by minimizing reactor over-cooling. Deep throttling also provides a practical transition from high pressures and powers of the high thrust power cycle to the low pressures and powers of our electric power generating mode. Two deep throttling designs are discussed; a workable system that was studied and a simplified system that is recommended for future study. Mission-optimized engine thrust/weight (T/W) and Isp predictions are included along with system flow schemes and concept sketches.

Hybrid Propulsion Technology Program

NASA Technical Reports Server (NTRS)

Jensen, G. E.; Holzman, A. L.

1990-01-01

Future launch systems of the United States will require improvements in booster safety, reliability, and cost. In order to increase payload capabilities, performance improvements are also desirable. The hybrid rocket motor (HRM) offers the potential for improvements in all of these areas. The designs are presented for two sizes of hybrid boosters, a large 4.57 m (180 in.) diameter booster duplicating the Advanced Solid Rocket Motor (ASRM) vacuum thrust-time profile and smaller 2.44 m (96 in.), one-quater thrust level booster. The large booster would be used in tandem, while eight small boosters would be used to achieve the same total thrust. These preliminary designs were generated as part of the NASA Hybrid Propulsion Technology Program. This program is the first phase of an eventual three-phaes program culminating in the demonstration of a large subscale engine. The initial trade and sizing studies resulted in preferred motor diameters, operating pressures, nozzle geometry, and fuel grain systems for both the large and small boosters. The data were then used for specific performance predictions in terms of payload and the definition and selection of the requirements for the major components: the oxidizer feed system, nozzle, and thrust vector system. All of the parametric studies were performed using realistic fuel regression models based upon specific experimental data.

Fin Ray Stiffness and Fin Morphology Control Ribbon-Fin-Based Propulsion.

PubMed

Liu, Hanlin; Taylor, Bevan; Curet, Oscar M

2017-06-01

Ribbon-fin-based propulsion has rich locomotor capabilities that can enhance the mobility and performance of underwater vehicles navigating in complex environments. Bony fishes using this type of propulsion send one or multiple traveling waves along an elongated fin with the actuation of highly flexible rays that are interconnected by an elastic membrane. In this work, we study how the use of flexible rays and different morphology can affect the performance of ribbon-fin propulsion. We developed a physical model composed of 15 rays that are interconnected with an elastic membrane. We tested four different ray flexural stiffness and four aspect ratios. The robotic model was tested in a low-turbulence flume under two flow conditions ([Formula: see text] wavelength/s). In two experimental sets, we measured fin kinematics, net surge forces, and power consumption. Using these data, we perform a thrust and power analysis of the undulating fin. We present the thrust coefficient, power coefficient, and propulsive efficiency. We find that the thrust generation was linear with the enclosed area swept by the fin, and square of the relative velocity between the incoming flow and traveling wave. The thrust coefficient levels off around 0.5. In addition, for our parameter range, we find that the power consumption scales by the cube of the effective tangential velocity of the rays [Formula: see text] (A is the amplitude of the ray oscillating motion, and [Formula: see text] is the angular velocity). We show that a decay in stiffness decreases both thrust production and power consumption. However, for rays with high flexural stiffness, the difference in thrust compared with rigid rays is minimal. Moreover, our results show that flexible rays can improve the propulsive efficiency compared with a rigid counterpart. Finally, we find that the morphology of ribbon fin affects its propulsive efficiency. For the aspect ratio considered in our experiments, [Formula: see text] was the most efficient compared with [Formula: see text]. Our results suggest that there could be an optimal morphology for a given ribbon fin kinematics. Therefore, both natural swimmers and underwater vehicles using ribbon-fin-based propulsion can take advantage of flexible rays and optimal aspect ratio to improve propulsive performance.

Analysis of a flare-director concept for an externally blown flap STOL aircraft

NASA Technical Reports Server (NTRS)

Middleton, D. B.

1974-01-01

A flare-director concept involving a thrust-required flare-guidance equation was developed and tested on a moving-base simulator. The equation gives a signal to command thrust as a linear function of the errors between the variables thrust, altitude, and altitude rate and corresponding values on a desired reference flare trajectory. During the simulator landing tests this signal drove either the horizontal command bar of the aircraft's flight director or a thrust-command dot on a head-up virtual-image display of a flare director. It was also used as the input to a simple autoflare system. An externally blown flap STOL (short take-off and landing) aircraft (with considerable stability and control augmentation) was modeled for the landing tests. The pilots considered the flare director a valuable guide for executing a proper flare-thrust program under instrument-landing conditions, but were reluctant to make any use of the head-up display when they were performing the landings visually.

Metaheuristic and Machine Learning Models for TFE-731-2, PW4056, and JT8D-9 Cruise Thrust

NASA Astrophysics Data System (ADS)

Baklacioglu, Tolga

2017-08-01

The requirement for an accurate engine thrust model has a major antecedence in airline fuel saving programs, assessment of environmental effects of fuel consumption, emissions reduction studies, and air traffic management applications. In this study, utilizing engine manufacturers' real data, a metaheuristic model based on genetic algorithms (GAs) and a machine learning model based on neural networks (NNs) trained with Levenberg-Marquardt (LM), delta-bar-delta (DBD), and conjugate gradient (CG) algorithms were accomplished to incorporate the effect of both flight altitude and Mach number in the estimation of thrust. For the GA model, the analysis of population size impact on the model's accuracy and effect of number of data on model coefficients were also performed. For the NN model, design of optimum topology was searched for one- and two-hidden-layer networks. Predicted thrust values presented a close agreement with real thrust data for both models, among which LM trained NNs gave the best accuracies.

Static performance tests of a flight-type STOVL ejector

NASA Technical Reports Server (NTRS)

Barankiewicz, Wendy S.

1991-01-01

The design and development of thrust augmenting STOVL ejectors has typically been based on experimental iteration (i.e., trial and error). Static performance tests of a full scale vertical lift ejector were performed at primary flow temperatures up to 1560 R (1100 F). Flow visualization (smoke generators and yarn tufts) were used to view the inlet air flow, especially around the primary nozzle and end plates. Performance calculations are presented for ambient temperatures close to 480 R (20 F) and 535 R (75 F) which simulate seasonal aircraft operating conditions. Resulting thrust augmentation ratios are presented as functions of nozzle pressure ratio and temperature.

Low-thrust chemical rocket engine study

NASA Technical Reports Server (NTRS)

Mellish, J. A.

1981-01-01

Engine data and information are presented to perform system studies on cargo orbit-transfer vehicles which would deliver large space structures to geosynchronous equatorial orbit. Low-thrust engine performance, weight, and envelope parametric data were established, preliminary design information was generated, and technologies for liquid rocket engines were identified. Two major engine design drivers were considered in the study: cooling and engine cycle options. Both film-cooled and regeneratively cooled engines were evaluated. The propellant combinations studied were hydrogen/oxygen, methane/oxygen, and kerosene/oxygen.

Computer program for flat sector thrust bearing performance

NASA Technical Reports Server (NTRS)

Presler, A. F.; Etsion, I.

1977-01-01

A versatile computer program is presented which achieves a rapid, numerical solution of the Reynolds equation for a flat sector thrust pad bearing with either compressible or liquid lubricants. Program input includes a range in values of the geometric and operating parameters of the sector bearing. Performance characteristics are obtained from the calculated bearing pressure distribution. These are the load capacity, center of pressure coordinates, frictional energy dissipation, and flow rates of liquid lubricant across the bearing edges. Two sample problems are described.

Performance seeking control (PSC) for the F-15 highly integrated digital electronic control (HIDEC) aircraft

NASA Technical Reports Server (NTRS)

Orme, John S.

1995-01-01

The performance seeking control algorithm optimizes total propulsion system performance. This adaptive, model-based optimization algorithm has been successfully flight demonstrated on two engines with differing levels of degradation. Models of the engine, nozzle, and inlet produce reliable, accurate estimates of engine performance. But, because of an observability problem, component levels of degradation cannot be accurately determined. Depending on engine-specific operating characteristics PSC achieves various levels performance improvement. For example, engines with more deterioration typically operate at higher turbine temperatures than less deteriorated engines. Thus when the PSC maximum thrust mode is applied, for example, there will be less temperature margin available to be traded for increasing thrust.

On channel interactions in nested Hall thrusters

NASA Astrophysics Data System (ADS)

Cusson, S. E.; Georgin, M. P.; Dragnea, H. C.; Dale, E. T.; Dhaliwal, V.; Boyd, I. D.; Gallimore, A. D.

2018-04-01

Nested Hall thrusters use multiple, concentric discharge channels to increase thrust density. They have shown enhanced performance in multi-channel operation relative to the superposition of individual channels. The X2, a two-channel nested Hall thruster, was used to investigate the mechanism behind this improved performance. It is shown that the local pressure near the thruster exit plane is an order of magnitude higher in two-channel operation. This is due to the increased neutral flow inherent to the multi-channel operation. Due to the proximity of the discharge channels in nested Hall thrusters, these local pressure effects are shown to be responsible for the enhanced production of thrust during multi-channel operation via two mechanisms. The first mechanism is the reduction of the divergence angle due to an upstream shift of the acceleration region. The displacement of the acceleration region was detected using laser induced fluorescence measurements of the ion velocity profile. Analysis of the change in beam divergence indicates that, at an operating condition of 150 V and 30 A, this effect increases the thrust by 8.7 ± 1.2 mN. The second mechanism is neutral ingestion from the adjacent channel resulting in a 2.0 + 0/-0.2 mN increase in thrust. Combined, these mechanisms are shown to explain, within uncertainty, the 17 ± 6.2 mN improvement in thrust during dual channel operation of the X2.

Oil-Free Turbomachinery Research Enhanced by Thrust Bearing Test Capability

NASA Technical Reports Server (NTRS)

Bauman, Steven W.

2003-01-01

NASA Glenn Research Center s Oil-Free Turbomachinery research team is developing aircraft turbine engines that will not require an oil lubrication system. Oil systems are required today to lubricate rolling-element bearings used by the turbine and fan shafts. For the Oil-Free Turbomachinery concept, researchers combined the most advanced foil (air) bearings from industry with NASA-developed high-temperature solid lubricant technology. In 1999, the world s first Oil-Free turbocharger was demonstrated using these technologies. Now we are working with industry to demonstrate Oil-Free turbomachinery technology in a small business jet engine, the EJ-22 produced by Williams International and developed during Glenn s recently concluded General Aviation Propulsion (GAP) program. Eliminating the oil system in this engine will make it simpler, lighter (approximately 15 percent), more reliable, and less costly to purchase and maintain. Propulsion gas turbines will place high demands on foil air bearings, especially the thrust bearings. Up until now, the Oil-Free Turbomachinery research team only had the capability to test radial, journal bearings. This research has resulted in major improvements in the bearings performance, but journal bearings are cylindrical, and can only support radial shaft loads. To counteract axial thrust loads, thrust foil bearings, which are disk shaped, are required. Since relatively little research has been conducted on thrust foil air bearings, their performance lags behind that of journal bearings.

ESCORT: A Pratt & Whitney nuclear thermal propulsion and power system for manned mars missions

NASA Astrophysics Data System (ADS)

Feller, Gerald J.; Joyner, Russell

1999-01-01

The purpose of this paper is to describe the conceptual design of an upgrade to the Pratt & Whitney ESCORT nuclear thermal rocket engine. The ESCORT is a bimodal engine capable of supporting a wide range of vehicle propulsive and electrical power requirements. The ESCORT engine is powered by a fast-spectrum beryllium-reflected CERMET-fueled nuclear reactor. In propulsive mode, the reactor is used to heat hot hydrogen to approximately 2700 K which is expanded through a converging/diverging nozzle to generate thrust. Heat pickup in the nozzle and the radial beryllium reflectors is used to drive the turbomachinery in the ESCORT expander cycle. In electrical mode, the reactor is used to heat a mixture of helium and xenon to drive a closed-loop Brayton cycle in order to generate electrical energy. This closed loop system has the additional function of a decay heat removal system after the propulsive mode operation is discontinued. The original ESCORT design was capable of delivering 4448.2 N (1000 lbf) of thrust at a vacuum impulse level of approximately 900 s. Design Reference Mission requirements (DRM) from NASA Johnson Space Center and NASA Lewis Research Center studies in 1997 and 1998 have detailed upgraded requirements for potential manned Mars missions. The current NASA DRM requires a nuclear thermal propulsion system capable of delivering total mission requirements of 200170 N (45000 lbf) thrust and 50 kWe of spacecraft electrical power. This is met assuming three engines capable of each delivering 66723 N (15000 lbf) of vacuum thrust and 25 kWe of electrical power. The individual engine requirements were developed assuming three out of three engine reliability for propulsion and two out of three engine reliability for spacecraft electrical power. The approximate target vacuum impulse is 925 s. The Pratt & Whitney ESCORT concept was upgraded to meet these requirements. The hexagonal prismatic fuel elements were modified to address the uprated power requirements while maintaining the peak fuel temperature below the 2880 K limit for W-UO2 CERMET fuels. A system integrated performance methodology was developed to assess the sensitivity to weight, thrust and impulse to the DRM requirements. Propellant tanks, shielding, and Brayton cycle power conversion unit requirements were included in this evaluation.

Status of 30 cm mercury ion thruster development

NASA Technical Reports Server (NTRS)

Sovey, J. S.; King, H. J.

1974-01-01

Two engineering model 30-cm ion thrusters were assembled, calibrated, and qualification tested. This paper discusses the thruster design, performance, and power system. Test results include documentation of thrust losses due to doubly charged mercury ions and beam divergence by both direct thrust measurements and beam probes. Diagnostic vibration tests have led to improved designs of the thruster backplate structure, feed system, and harness. Thruster durability is being demonstrated over a thrust range of 97 to 113 mN at a specific impulse of about 2900 seconds. As of August 15, 1974, the thruster has successfully operated for over 4000 hours.

An accurate solution of the gas lubricated, flat sector thrust bearing

NASA Technical Reports Server (NTRS)

Etsion, I.; Fleming, D. P.

1976-01-01

A flat sector shaped pad geometry for gas lubricated thrust bearings is analyzed considering both pitch and roll angles of the pad and the true film thickness distribution. Maximum load capacity is achieved when the pad is tilted so as to create a uniform minimum film thickness along the pad trailing edge. Performance characteristics for various geometries and operating conditions of gas thrust bearings are presented in the form of design curves. A comparison is made with the rectangular slider approximation. It is found that this approximation is unsafe for practical design, since it always overestimates load capacity.

Effect of low-stiffness closeout overwrap on rocket thrust-chamber life

NASA Technical Reports Server (NTRS)

Kasper, H. J.; Nota-Donato, J. J.

1979-01-01

Three rocket thrust chambers with copper liners and a thrust level of 20.9 kN were cyclically test fired to failure. Two of the liners were made from oxygen free, high conductivity (OFHC) copper and from annealed Amzirc. The milled coolant channels were closed out with a thin copper closeout over which a fiberglass composite was wrapped to provide hoop strength only. Experimental data are presented, along with the results of a preliminary analysis that was performed before fabrication to evaluate the life extending potential of a thin copper closeout with a fiberglass overwrap.

Analysis of the gas-lubricated flat-sector-pad thrust bearing

NASA Technical Reports Server (NTRS)

Etsion, I.

1976-01-01

A flat sector-shaped pad geometry for a gas-lubricated thrust bearing is analyzed considering both the pitch and roll of the pad. It is shown that maximum load capacity is achieved when the pad is tilted so as to create uniform minimum film thickness along the pad trailing edge. Performance characteristics for various geometries and operating conditions of gas thrust bearings are presented in the form of design curves, and a comparison is made with the rectangular slider approximation. It is found that this approximation is unsafe for practical design, since it always overestimates load capacity.

Optimal Interception of a Maneuvering Long-range Missile

NASA Astrophysics Data System (ADS)

X. Vinh, Nguyen; T. Kabamba, Pierre; Takehira, Tetsuya

2001-01-01

In a Newtonian central force field, the minimum-fuel interception of a satellite, or a ballistic missile, in elliptic trajectory can be obtained via Lawden's theory of primer vector. To secure interception when the target performs evasive maneuvers, a new control law, with explicit solutions, is implemented. It is shown that by a rotation of coordinate system, the problem of three-dimensional interception is reduced to a planar problem. The general case of planar interception of a long-range ballistic missile is then studied. Examples of interception at a specified time, head-on interception and minimum-fuel interception are presented. In each case, the requirement for the thrust acceleration is expressed explicitly as a function of time.

Associations of Varus Thrust and Alignment with Pain in Knee Osteoarthritis

PubMed Central

Lo, Grace H.; Harvey, William F.; McAlindon, Timothy E.

2012-01-01

Objective To compare associations of varus thrust and varus static alignment with pain in those with knee osteoarthritis (OA). Method This was a cross-sectional study of participants from a randomized controlled trial of vitamin D for knee OA. Participants were video recorded walking and scored for presence of varus thrust. Standard PA knee X-rays were measured for static alignment. Pain questions from the Western Ontario McMasters Osteoarthritis (WOMAC) questionnaire assessed symptoms. We calculated means for total WOMAC pain by varus thrust and varus alignment (i.e. corrected anatomic alignment < 178°). We performed ordinal logistic regressions; outcomes: individual WOMAC pain questions; predictors: varus thrust and varus alignment. Results There were 82 participants, mean age 65.1 (±8.5), mean body mass index 30.2 (±5.4), and 60% female. Total WOMAC pain was 6.3 versus 3.9, p = 0.007 in those with versus without definite varus thrust. For varus alignment, total WOMAC pain was 5.2 versus 4.2, p = 0.30. Odds ratios for pain with walking and standing were 5.5 (95%CI 2.0 – 15.1) and 6.0 (95%CI 2.2 – 16.2) in those with versus without definite varus thrust. There were no significant associations between varus alignment and individual WOMAC pain questions. Sensitivity analyses suggested a more stringent definition of varus might have been associated with walking and standing pain. Conclusion In those with knee OA, varus thrust and possibly varus static alignment, were associated with pain, specifically during weight-bearing activities. Treatment of varus thrust (e.g. via bracing or gait modification) may lead to improvement of symptoms. PMID:22307813

Entrainment and thrust augmentation in pulsatile ejector flows

NASA Technical Reports Server (NTRS)

Sarohia, V.; Bernal, L.; Bui, T.

1981-01-01

This study comprised direct thrust measurements, flow visualization by use of a spark shadowgraph technique, and mean and fluctuating velocity measurements with a pitot tube and linearized constant temperature hot-wire anemometry respectively. A gain in thrust of as much as 10 to 15% was observed for the pulsatile ejector flow as compared to the steady flow configuration. From the velocity profile measurements, it is concluded that this enhanced augmentation for pulsatile flow as compared to a nonpulsatile one was accomplished by a corresponding increased entrainment by the primary jet flow. It is also concluded that the augmentation and total entrainment by a constant area ejector critically depends upon the inlet geometry of the ejector. Experiments were performed to evaluate the influence of primary jet to ejector area ratio, ejector length, and presence of a diffuser on pulsatile ejector performance.

76 FR 27952 - Airworthiness Directives; Eurocopter France Model EC 120B Helicopters

Federal Register 2010, 2011, 2012, 2013, 2014

2011-05-13

... perform a functional test of the cyclic control. (b) To request a different method of compliance or a... control friction device by replacing a certain thrust washer with two thrust washers. This proposed AD is prompted by an incident in which the pilot encountered a sudden restriction of the cyclic control movement...

The Application of Biocybernetic Techniques to Enhance Pilot Performance during Tactical missions.

DTIC Science & Technology

1979-10-01

and observe the altitude indicator ( meatball ) at the end of the runway. At touchdown, the pilot must apply thrust until the arresting hook catches. 103...l PILOT TASES 3 ~ CRITICALI DIFFICULTY J.[ 27 COMMAND TYRUST FOR BOLTER LT SAFETY" D’ENM .....MM................. ...... I.26 MONITOR MEATBALL FLT...THRUST FOR BOLTER ............................................................................ X MONITOR MEATBALL

Unmanned Intratheater Airlift

DTIC Science & Technology

2010-12-01

aircraft use a standard forward- thrust turbojet or turbofan engine that provides forward thrust while in forward flight. However, when performing...Afghanistan and Iraq required resupply by ground convoy, airdrop, or verticaltakeoff- and-landing aircraft . Unfortunately, the Department of Defense (DOD...challenges and requirements, the author suggests the use of remotely piloted aircraft (RPA) as a potential solution. Focusing on existing and quickly

Performance improvements of an F-15 airplane with an integrated engine-flight control system

NASA Technical Reports Server (NTRS)

Myers, Lawrence P.; Walsh, Kevin R.

1988-01-01

An integrated flight and propulsion control system has been developed and flight demonstrated on the NASA Ames-Dryden F-15 research aircraft. The highly integrated digital control (HIDEC) system provides additional engine thrust by increasing engine pressure ratio (EPR) at intermediate and afterburning power. The amount of EPR uptrim is modulated based on airplane maneuver requirements, flight conditions, and engine information. Engine thrust was increased as much as 10.5 percent at subsonic flight conditions by uptrimming EPR. The additional thrust significantly improved aircraft performance. Rate of climb was increased 14 percent at 40,000 ft and the time to climb from 10,000 to 40,000 ft was reduced 13 percent. A 14 and 24 percent increase in acceleration was obtained at intermediate and maximum power, respectively. The HIDEC logic performed fault free. No engine anomalies were encountered for EPR increases up to 12 percent and for angles of attack and sideslip of 32 and 11 deg, respectively.

Performance improvements of an F-15 airplane with an integrated engine-flight control system

NASA Technical Reports Server (NTRS)

Myers, Lawrence P.; Walsh, Kevin R.

1988-01-01

An integrated flight and propulsion control system has been developed and flight demonstrated on the NASA Ames-Dryden F-15 research aircraft. The highly integrated digital control (HIDEC) system provides additional engine thrust by increasing engine pressure ratio (EPR) at intermediate and afterburning power. The amount of EPR uptrim is modulated based on airplane maneuver requirements, flight conditions, and engine information. Engine thrust was increased as much as 10.5 percent at subsonic flight conditions by uptrimming EPR. The additional thrust significantly improved aircraft performance. Rate of climb was increased 14 percent at 40,000 ft and the time to climb from 10,000 to 40,000 ft was reduced 13 percent. A 14 and 24 percent increase in acceleration was obtained at intermediate and maximum power, respectively. The HIDEC logic performed fault free. No engine anomalies were encountered for EPR increases up to 12 percent and for angles of attack and sideslip of 32 and 11 degrees, respectively.

Performance Charts for a Turbojet System

NASA Technical Reports Server (NTRS)

Karp, Irving M.

1947-01-01

Convenient charts are presented for computing the thrust, fuel consumption, and other performance values of a turbojet system. These charts take into account the effects of ram pressure, compressor pressure ratio, ratio of combustion-chamber-outlet temperature to atmospheric temperature, compressor efficiency, turbine efficiency, combustion efficiency, discharge-nozzle coefficient, losses in total pressure in the inlet to the jet-propulsion unit and in the combustion chamber, and variation in specific heats with temperature. The principal performance charts show clearly the effects of the primary variables and correction charts provide the effects of the secondary variables. The performance of illustrative cases of turbojet systems is given. It is shown that maximum thrust per unit mass rate of air flow occurs at a lower compressor pressure ratio than minimum specific fuel consumption. The thrust per unit mass rate of air flow increases as the combustion-chamber discharge temperature increases. For minimum specific fuel consumption, however, an optimum combustion-chamber discharge temperature exists, which in some cases may be less than the limiting temperature imposed by the strength temperature characteristics of present materials.

Developing stochastic model of thrust and flight dynamics for small UAVs

NASA Astrophysics Data System (ADS)

Tjhai, Chandra

This thesis presents a stochastic thrust model and aerodynamic model for small propeller driven UAVs whose power plant is a small electric motor. First a model which relates thrust generated by a small propeller driven electric motor as a function of throttle setting and commanded engine RPM is developed. A perturbation of this model is then used to relate the uncertainty in throttle and engine RPM commanded to the error in the predicted thrust. Such a stochastic model is indispensable in the design of state estimation and control systems for UAVs where the performance requirements of the systems are specied in stochastic terms. It is shown that thrust prediction models for small UAVs are not a simple, explicit functions relating throttle input and RPM command to thrust generated. Rather they are non-linear, iterative procedures which depend on a geometric description of the propeller and mathematical model of the motor. A detailed derivation of the iterative procedure is presented and the impact of errors which arise from inaccurate propeller and motor descriptions are discussed. Validation results from a series of wind tunnel tests are presented. The results show a favorable statistical agreement between the thrust uncertainty predicted by the model and the errors measured in the wind tunnel. The uncertainty model of aircraft aerodynamic coefficients developed based on wind tunnel experiment will be discussed at the end of this thesis.

Single-stage EHD thruster response to several simulation conditions in nitrogen gas

NASA Astrophysics Data System (ADS)

Granados, Victor H.; Pinheiro, Mario J.; Sá, Paulo A.

2017-09-01

We use a numerical model to investigate the influence of pressure from 0.5 Torr (66.7 Pa) to 100 Torr (13.3 kPa) and temperature (190-400 K) on the performance (thrust, fluid velocity, and thrust-to-power-ratio) of a single stage electrohydrodynamic thruster made of a rod anode and funnel-like cathode geometry, using nitrogen as the working gas. The model includes the following nitrogen species: N, N+, N2, N2+ , and N4+ . Additional factors are investigated: (i) the ballast resistance, (ii) the secondary electron emission from the cathode (in the range of 10-5 -10°), and (iii) the influence of the gap between electrodes on the discharge. As expected, higher pressures increase the net thrust, thrust efficiency, and peak gas velocity; however, with increasing temperatures, the trend reverses. We notice that gas flow velocity diminishes for the increasing values of the secondary emission coefficient, and it is possible to identify two working regimes presenting different behaviors: in the first region, for values of the secondary electron emission coefficient between 10-5 and 10-2 , thrust was not affected, and in the second region, between 10-2 and 1, a clear decrease in thrust is observed, accompanied by an increase in the discharge current, an undesired effect for the purpose of thrust production.

Aerodynamics of a Flapping Airfoil with a Flexible Tail

NASA Astrophysics Data System (ADS)

Lai, Alan Kai San

This dissertation presents computational solutions to an airfoil in a oscillatory heaving motion with a aeroelastically flexible tail attachment. An unsteady potential flow solver is coupled to a structural solver to obtain the aeroelastic flow solution over an inviscid fluid to investigate the propulsive performance of such a configuration. The simulation is then extended to a two-dimensional viscous solver by coupling NASA's CFL3D solver to the structural solver to study how the flow is altered by the presence of viscosity. Finally, additional simulations are done in three dimensions over wings with varying aspect ratio to study the three-dimensional effects on the propulsive performance of an airfoil with an aeroelastic tail. The computation reveals that the addition of the aeroelastic trailing edge improved the thrust generated by a heaving airfoil significantly. As the frequency of the heaving motion increases, the thrust generated by the airfoil with the tail increases exponentially. In an inviscid fluid, the increase in thrust is insufficient to overcome the increase in power required to maintain the motion and as a result the overall propulsive efficiency is reduced. When the airfoil is heaving in a viscous fluid, the presence of a suction boundary layer and the appearance of leading edge vortex increase the thrust generated to such an extent that the propulsive efficiency is increased by about 3% when compared to the same airfoil with a rigid tail. The three-dimensional computations shows that the presence of the tip vorticies suppress some of the increase in thrust observed in the two-dimensional viscous computations for short span wings. For large span wings, the overall thrust enhancing capabilities of the aeroelastic tail is preserved.

Flight assessment of the onboard propulsion system model for the Performance Seeking Control algorithm on an F-15 aircraft

NASA Technical Reports Server (NTRS)

Orme, John S.; Schkolnik, Gerard S.

1995-01-01

Performance Seeking Control (PSC), an onboard, adaptive, real-time optimization algorithm, relies upon an onboard propulsion system model. Flight results illustrated propulsion system performance improvements as calculated by the model. These improvements were subject to uncertainty arising from modeling error. Thus to quantify uncertainty in the PSC performance improvements, modeling accuracy must be assessed. A flight test approach to verify PSC-predicted increases in thrust (FNP) and absolute levels of fan stall margin is developed and applied to flight test data. Application of the excess thrust technique shows that increases of FNP agree to within 3 percent of full-scale measurements for most conditions. Accuracy to these levels is significant because uncertainty bands may now be applied to the performance improvements provided by PSC. Assessment of PSC fan stall margin modeling accuracy was completed with analysis of in-flight stall tests. Results indicate that the model overestimates the stall margin by between 5 to 10 percent. Because PSC achieves performance gains by using available stall margin, this overestimation may represent performance improvements to be recovered with increased modeling accuracy. Assessment of thrust and stall margin modeling accuracy provides a critical piece for a comprehensive understanding of PSC's capabilities and limitations.

Study of aerodynamic technology for VSTOL fighter/attack aircraft, phase 1

NASA Technical Reports Server (NTRS)

Driggers, H. H.

1978-01-01

A conceptual design study was performed of a vertical attitude takeoff and landing (VATOL) fighter/attack aircraft. The configuration has a close-coupled canard-delta wing, side two-dimensional ramp inlets, and two augmented turbofan engines with thrust vectoring capability. Performance and sensitivities to objective requirements were calculated. Aerodynamic characteristics were estimated based on contractor and NASA wind tunnel data. Computer simulations of VATOL transitions were performed. Successful transitions can be made, even with series post-stall instabilities, if reaction controls are properly phased. Principal aerodynamic uncertainties identified were post-stall aerodynamics, transonic aerodynamics with thrust vectoring and inlet performance in VATOL transition. A wind tunnel research program was recommended to resolve the aerodynamic uncertainties.

Status of the NEXT Ion Engine Wear Test

NASA Technical Reports Server (NTRS)

Soulas, George C.; Domonkos, Matthew T.; Kamhawi, Hani; Patterson, Michael J.; Gardner, Michael M.

2003-01-01

The status of the NEXT 2000 hour wear test is presented. This test is being conducted with a 40 cm engineering model ion engine, designated EM1, at a beam current higher than listed on the NEXT throttle table. Pretest performance assessments demonstrated that EM1 satisfies all thruster performance requirements. As of 7/3/03, the ion engine has accumulated 406 hours of operation at a thruster input power of 6.9 kW. Overall ion engine performance, which includes thrust, thruster input power, specific impulse, and thrust efficiency, has been steady to date with no indications of performance degradation. Images of the downstream discharge cathode, neutralizer, and accelerator aperture surfaces have exhibited no significant erosion to date.

Augmenting ejector endwall effects. [V/STOL aircraft

NASA Technical Reports Server (NTRS)

Porter, J. L.; Squyers, R. A.

1979-01-01

Rectangular inlet ejectors which had multiple hypermixing nozzles for their primary jets were investigated for the effects of endwall blowing on thrust augmentation performance. The ejector configurations tested had both straight wall and active boundary layer control type diffusers. Endwall flows were energized and controlled by simple blowing jets suitably located in the ejector. Both the endwall and boundary layer control diffuser blowing rates were varied to determine optimum performance. High area ratio diffusers with insufficient endwall blowing showed endwall separation and rapid degradation of thrust performance. Optimized values of diffuser boundary layer control and endwall nozzle blowing rates in an ejector augmenter were shown to achieve high levels of augmentation performance for maximum compactness.

Test Plan. GCPS Task 4, subtask 4.2 thrust structure development

NASA Astrophysics Data System (ADS)

Greenberg, H. S.

1994-09-01

The Single Stage To Orbit (SSTO) vehicle is designed to lift off from a vertical position, go into orbit, return to earth for a horizontal landing, and be reusable for the next mission. (NASA baseline only) In order to meet its performance goals, the SSTO relies on light weight structure and the use of 8 tri-propellant engines. These engines are mounted to the thrust structure. This test plan addresses selection of the material for this structure, and the integrity of the design through testing of elements and a full-scale subcomponent. This test plan supports the development of the design for an advanced composite thrust structure for a Single Stage to Orbit manned, heavy launch vehicle. The thrust structure is designed to transmit very high thrust loads from the engines to the rest of the vehicle (see Figure 1 ). The thrust structure will also be used for primary attachment of the twin vertical tails and possibly act as the aft attach point for the wing. The combination of high loading, high vibration, long service life and high acoustic environments will need to be evaluated by tests. To minimize design risk, a building block approach will be used. We will first screen materials to determine which materials show the most promise for this application. Factors in this screening will be the suitability of these materials for chosen design concepts, particularly concerning specific strength, environmental compatibility and applicability to fabrication processes. Next we will characterize two material systems that will be used in the design; the characterization will allow us to generate preliminary design data that will be used for the analysis. Element testing will be performed to evaluate critical structural locations under load. Final testing on the full scale test article will be performed to verify the design and to demonstrate predictability of the analysis. Additionally, risks associated with fabricating full scale thrust structures will be reduced through testing activities. One of the major concerns that stems from full scale fabrication is the realities of size and the associated complexities of handling, manufacturing, and assembly. The need exists to fabricate, assemble and test_representative joint specimens to achieve_confidence in the design and manufacturing technologies being proposed.

Test Plan. GCPS Task 4, subtask 4.2 thrust structure development

NASA Technical Reports Server (NTRS)

Greenberg, H. S.

1994-01-01

The Single Stage To Orbit (SSTO) vehicle is designed to lift off from a vertical position, go into orbit, return to earth for a horizontal landing, and be reusable for the next mission. (NASA baseline only) In order to meet its performance goals, the SSTO relies on light weight structure and the use of 8 tri-propellant engines. These engines are mounted to the thrust structure. This test plan addresses selection of the material for this structure, and the integrity of the design through testing of elements and a full-scale subcomponent. This test plan supports the development of the design for an advanced composite thrust structure for a Single Stage to Orbit manned, heavy launch vehicle. The thrust structure is designed to transmit very high thrust loads from the engines to the rest of the vehicle (see Figure 1 ). The thrust structure will also be used for primary attachment of the twin vertical tails and possibly act as the aft attach point for the wing. The combination of high loading, high vibration, long service life and high acoustic environments will need to be evaluated by tests. To minimize design risk, a building block approach will be used. We will first screen materials to determine which materials show the most promise for this application. Factors in this screening will be the suitability of these materials for chosen design concepts, particularly concerning specific strength, environmental compatibility and applicability to fabrication processes. Next we will characterize two material systems that will be used in the design; the characterization will allow us to generate preliminary design data that will be used for the analysis. Element testing will be performed to evaluate critical structural locations under load. Final testing on the full scale test article will be performed to verify the design and to demonstrate predictability of the analysis. Additionally, risks associated with fabricating full scale thrust structures will be reduced through testing activities. One of the major concerns that stems from full scale fabrication is the realities of size and the associated complexities of handling, manufacturing, and assembly. The need exists to fabricate, assemble and test_representative joint specimens to achieve_confidence in the design and manufacturing technologies being proposed.

Computing Satellite Maneuvers For A Repeating Ground Track

NASA Technical Reports Server (NTRS)

Shapiro, Bruce

1994-01-01

TOPEX/POSEIDON Ground Track Maintenance Maneuver Targeting Program (GTARG) assists in designing maneuvers to maintain orbit of TOPEX/POSEIDON satellite. Targeting strategies used either maximize time between maneuvers or force control band exit to occur at specified intervals. Runout mode allows for ground-track propagation without targeting. GTARG incorporates analytic mean-element propagation algorithm accounting for all perturbations known to cause significant variations in ground track. Perturbations include oblateness of Earth, luni-solar gravitation, drag, thrusts associated with impulsive maneuvers, and unspecified fixed forces acting on satellite in direction along trajectory. Written in VAX-FORTRAN.

Preliminary tests of the electrostatic plasma accelerator

NASA Technical Reports Server (NTRS)

Aston, G.; Acker, T.

1990-01-01

This report describes the results of a program to verify an electrostatic plasma acceleration concept and to identify those parameters most important in optimizing an Electrostatic Plasma Accelerator (EPA) thruster based upon this thrust mechanism. Preliminary performance measurements of thrust, specific impulse and efficiency were obtained using a unique plasma exhaust momentum probe. Reliable EPA thruster operation was achieved using one power supply.

The XFV-12A Thrust-Augmented Wing (TAW) prototype aircraft

NASA Technical Reports Server (NTRS)

Murphy, R.; Lewis, E. L.

1979-01-01

The XFV-12A, a unique V/STOL technology prototype aircraft being developed for the Navy, is described. The innovative design features a thrust augmented wing and a canard ejector. Structural, functional, and control test performances are discussed. Static tether test results are also discussed. Assessment of test results are given along with projections for future modification areas.

Hovering and Low-Speed Performance and Control Characteristics of the Kaman Helicopter Rotor System as Determined on the Langley Helicopter Tower. TED No. NACA DE 205

NASA Technical Reports Server (NTRS)

Carpenter, Paul J.; Paulnock, Russell S.

1949-01-01

An investigation has been conducted with the Langley helicopter tower to obtain basic performance and control characteristics of the Raman rotor system. Blade-pitch control is obtained in this configuration by utilizing an auxiliary flap to twist the blades. Rotor thrust and power required were measured for the hovering condition and over a range of wind velocities from 0 to 30 miles per hour. The control characteristics and the transient response of the rotor to various control movements were also measured. The hovering-performance data are presented as a survey of the wake velocities and the variation of torque coefficient with thrust coefficient. The power required for the test rotor to hover at a thrust of 1350 pounds and a rotor speed of 240 rpm is approximately 6.5 percent greater than that estimated for a conventional rotor of the same diameter and solidity. It is believed that most of this difference is caused by th e flap servomechanism. The reduction in total power required for sustentation of the single-rotor configuration tested at various wind velocities and at the normal operating rotor thrust was found to be similar to the theoretical and experimental results for ro tors with conventionally actuated pitch. The control effectiveness was determined as a function of rotor speed. Sufficient control was available to give a thrust range of 0 to 1500 pounds and a rotor tilt of plus or minus 7 degrees. The time lag between flap motion and blade-pitch response is approximately 0.02 to 0.03 second. The response of the rotor following the blade-pitch response is similar to that of a rotor with conventionally actuated pitch changes. The over-all characteristics of the rotor investigated indicate that satisfactory performance and control characteristics were obtained.

Design analysis and performance assessment of hybrid magnetic bearings for a rotary centrifugal blood pump.

PubMed

Ren, Zhaohui; Jahanmir, Said; Heshmat, Hooshang; Hunsberger, Andrew Z; Walton, James F

2009-01-01

A hybrid magnetic bearing system was designed for a rotary centrifugal blood pump being developed to provide long-term circulatory support for heart failure patients. This design consists of two compact bearings to suspend the rotor in five degrees-of-freedom with single axis active control. Permanent magnets are used to provide passive radial support and electromagnets to maintain axial stability of the rotor. Characteristics of the passive radial and active thrust magnetic bearing system were evaluated by the electromagnetic finite element analysis. A proportional-integral-derivative controller with force balance algorithm was implemented for closed loop control of the magnetic thrust bearing. The control position is continuously adjusted based on the electrical energy in the bearing coils, and thus passive magnetic forces carry static thrust loads to minimize the bearing current. Performance of the magnetic bearing system with associated control algorithm was evaluated at different operating conditions. The bearing current was significantly reduced with the force balance control method and the power consumption was below 0.5 W under various thrust loads. The bearing parameters predicted by the analysis were validated by the experimental data.

Jet Spreading Increase by Passive Control and Associated Performance Penalty

NASA Technical Reports Server (NTRS)

Zaman, K. B. M. Q.

1999-01-01

This paper reviews the effects of 'screech', 'asymmetric nozzle shaping', 'tabs' and 'overexpansion' on the spreading of free jets. Corresponding thrust penalty for the tabs and overexpanded condition are also evaluated. The asymmetric shapes include rectangular ones with varying aspect ratio. Tabs investigated are triangular shaped 'delta-tabs' placed at the exit of a convergent circular nozzle. The effect of overexpansion is examined with circular convergent-divergent (C-D) nozzles. Tabs and overexpansion are found to yield the largest increase in jet spreading. Each, however, involves a performance penalty, i.e., a loss in thrust coefficient. Variation of the size of four delta-tabs show that there exists an optimum size for which the gain in jet spreading is the maximum per unit loss in thrust coefficient. With the C-D nozzles, the minimum in thrust coefficient is expected near the beginning of the overexpanded regime based on idealized flow calculations. The maximum increase in jet spreading, however, is found to occur at higher pressure ratios well into the overexpanded regime. The optimum benefit with the overexpanded flow, in terms of gain in spreading for unit penalty, is found to be comparable to the optimum tab case.

CVD Rhenium Engines for Solar-Thermal Propulsion Systems

NASA Technical Reports Server (NTRS)

Williams, Brian E.; Fortini, Arthur J.; Tuffias, Robert H.; Duffy, Andrew J.; Tucker, Stephen P.

1999-01-01

Solar-thermal upper-stage propulsion systems have the potential to provide specific impulse approaching 900 seconds, with 760 seconds already demonstrated in ground testing. Such performance levels offer a 100% increase in payload capability compared to state-of-the-art chemical upper-stage systems, at lower cost. Although alternatives such as electric propulsion offer even greater performance, the 6- to 18- month orbital transfer time is a far greater deviation from the state of the art than the one to two months required for solar propulsion. Rhenium metal is the only material that is capable of withstanding the predicted thermal, mechanical, and chemical environment of a solar-thermal propulsion device. Chemical vapor deposition (CVD) is the most well-established and cost-effective process for the fabrication of complex rhenium structures. CVD rhenium engines have been successfully constructed for the Air Force ISUS program (bimodal thrust/electricity) and the NASA Shooting Star program (thrust only), as well as under an Air Force SBIR project (thrust only). The bimodal engine represents a more long-term and versatile approach to solar-thermal propulsion, while the thrust-only engines provide a potentially lower weight/lower cost and more near-term replacement for current upper-stage propulsion systems.

Closed-loop thrust and pressure profile throttling of a nitrous oxide/hydroxyl-terminated polybutadiene hybrid rocket motor

NASA Astrophysics Data System (ADS)

Peterson, Zachary W.

Hybrid motors that employ non-toxic, non-explosive components with a liquid oxidizer and a solid hydrocarbon fuel grain have inherently safe operating characteristics. The inherent safety of hybrid rocket motors offers the potential to greatly reduce overall operating costs. Another key advantage of hybrid rocket motors is the potential for in-flight shutdown, restart, and throttle by controlling the pressure drop between the oxidizer tank and the injector. This research designed, developed, and ground tested a closed-loop throttle controller for a hybrid rocket motor using nitrous oxide and hydroxyl-terminated polybutadiene as propellants. The research simultaneously developed closed-loop throttle algorithms and lab scale motor hardware to evaluate the fidelity of the throttle simulations and algorithms. Initial open-loop motor tests were performed to better classify system parameters and to validate motor performance values. Deep-throttle open-loop tests evaluated limits of stable thrust that can be achieved on the test hardware. Open-loop tests demonstrated the ability to throttle the motor to less than 10% of maximum thrust with little reduction in effective specific impulse and acoustical stability. Following the open-loop development, closed-loop, hardware-in-the-loop tests were performed. The closed-loop controller successfully tracked prescribed step and ramp command profiles with a high degree of fidelity. Steady-state accuracy was greatly improved over uncontrolled thrust.

Adaptation Method for Overall and Local Performances of Gas Turbine Engine Model

NASA Astrophysics Data System (ADS)

Kim, Sangjo; Kim, Kuisoon; Son, Changmin

2018-04-01

An adaptation method was proposed to improve the modeling accuracy of overall and local performances of gas turbine engine. The adaptation method was divided into two steps. First, the overall performance parameters such as engine thrust, thermal efficiency, and pressure ratio were adapted by calibrating compressor maps, and second, the local performance parameters such as temperature of component intersection and shaft speed were adjusted by additional adaptation factors. An optimization technique was used to find the correlation equation of adaptation factors for compressor performance maps. The multi-island genetic algorithm (MIGA) was employed in the present optimization. The correlations of local adaptation factors were generated based on the difference between the first adapted engine model and performance test data. The proposed adaptation method applied to a low-bypass ratio turbofan engine of 12,000 lb thrust. The gas turbine engine model was generated and validated based on the performance test data in the sea-level static condition. In flight condition at 20,000 ft and 0.9 Mach number, the result of adapted engine model showed improved prediction in engine thrust (overall performance parameter) by reducing the difference from 14.5 to 3.3%. Moreover, there was further improvement in the comparison of low-pressure turbine exit temperature (local performance parameter) as the difference is reduced from 3.2 to 0.4%.

Installation effects on performance of multiple model V/STOL lift fans

NASA Technical Reports Server (NTRS)

Diedrich, J. H.; Clough, N.; Lieblein, S.

1972-01-01

An experimental program was performed in which the individual performance of multiple VTOL model lift fans was measured. The model tested consisted of three 5.5 in. diameter tip-turbine driven model VTOL lift fans mounted chordwise in a two-dimensional wing to simulate a pod-type array. The performance data provided significant insight into possible thrust variations and losses caused by the presence of cover doors, adjacent fuselage panels, and adjacent fans. The effect of a partial loss of drive air supply (simulated gas generator failure) on fan performance was also investigated. The results of the tests demonstrated that lift fan installation variables and hardware can have a significant effect on the thrust of the individual fans.

Design and Demonstration of Emergency Control Modes for Enhanced Engine Performance

NASA Technical Reports Server (NTRS)

Liu, Yuan; Litt, Jonathan S.; Guo, Ten-Huei

2013-01-01

A design concept is presented for developing control modes that enhance aircraft engine performance during emergency flight scenarios. The benefits of increased engine performance to overall vehicle survivability during these situations may outweigh the accompanied elevated risk of engine failure. The objective involves building control logic that can consistently increase engine performance beyond designed maximum levels based on an allowable heightened probability of failure. This concept is applied to two previously developed control modes: an overthrust mode that increases maximum engine thrust output and a faster response mode that improves thrust response to dynamic throttle commands. This paper describes the redesign of these control modes and presents simulation results demonstrating both enhanced engine performance and robust maintenance of the desired elevated risk level.

Mission Design Analysis for the Martian Moon Phobos: Close Flybys, Missed Thrusts, and Other In-Flight Entertainment

NASA Technical Reports Server (NTRS)

Stuart, Jeffrey; McElrath, Tim; Petropoulos, Anastassios

2015-01-01

A robotic mission to the Martian moons Phobos and Deimos would offer a wealth of scientific information and serve as a useful precursor to potential human missions. In this paper, we investigate a prospective mission enabled by solar electric propulsion that would explore Phobos via a series of flybys followed by capture into orbit around the moon. Of particular interest are low-cost options for capture and walkdown to the target science orbits aided by multi-body effects due to the mutual gravitational interaction of Phobos and Mars. We also consider contingency operations in the event of missed thrust or maneuver execution errors.

Experiment Pamir-3. Coplanar emission of high energy gamma-quanta at interaction of hadrons with nuclei of air atoms at energies above 10 to the 7th power GeV

NASA Technical Reports Server (NTRS)

Asatiani, T. L.; Genina, L. E.; Zatsepin, G. T.

1985-01-01

A systematic analysis of large gamma families, detected in X-ray emulsion chambers, cases of multicore halos have been observed, and among them five events in which the halo is divided into three of four separate cores with their alignment observed in the target diagram (coplanarity of axes of corresponding electron photon cascades). The halo alignment (tendency to the straight line) leads to the aximuthal asymmetry (thrust). The analysis of lateral and momentum distributions of particles in these families shows that they also have thrust that correlates with the direction of the halo core alignment.

Guidance and Navigation Requirements for Unmanned Flyby and Swingby Missions to the Outer Planets. Volume 3; Low Thrust Missions, Phase B

NASA Technical Reports Server (NTRS)

1970-01-01

The guidance and navigation requirements for unmanned missions to the outer planets, assuming constant, low thrust, ion propulsion are discussed. The navigational capability of the ground based Deep Space Network is compared to the improvements in navigational capability brought about by the addition of guidance and navigation related onboard sensors. Relevant onboard sensors include: (1) the optical onboard navigation sensor, (2) the attitude reference sensors, and (3) highly sensitive accelerometers. The totally ground based, and the combination ground based and onboard sensor systems are compared by means of the estimated errors in target planet ephemeris, and the spacecraft position with respect to the planet.

Study on electromagnetic plasma propulsion using rotating magnetic field acceleration scheme

NASA Astrophysics Data System (ADS)

Furukawa, T.; Takizawa, K.; Kuwahara, D.; Shinohara, S.

2017-04-01

As one of the electromagnetic plasma acceleration systems, we have proposed a rotating magnetic field (RMF) acceleration scheme to overcome the present problem of direct plasma-electrode interactions, leading to a short lifetime with a poor plasma performance due to contamination. In this scheme, we generate a plasma by a helicon wave excited by a radio frequency (rf) antenna which has no direct-contact with a plasma. Then, the produced plasma is accelerated by the axial Lorentz force fz = jθ × Br (jθ is an azimuthal current induced by RMF, and Br is an external radial magnetic field). Erosion of electrodes and contamination are not expected in this total system since RMF coils and an rf antenna do not have contact with the plasma directly. Here, we have measured the plasma parameters (electron density ne and axial ion velocity vi) to demonstrate this RMF acceleration scheme by the use of AC currents in two sets of opposing coils to generate a RMF. The maximum increasing rate Δvi /vi was ˜28% (maximum vi of ˜3 km/s), while the density increasing rate of Δne/ne is ˜ 70% in the case of a RMF current frequency fRMF of 3 MHz, which showed a better plasma performance than that with fRMF = 5 MHz. Moreover, thrust characteristics such as a specific impulse and a thrust efficiency were discussed, although a target plasma was not optimized.

A static investigation of yaw vectoring concepts on two-dimensional convergent-divergent nozzles

NASA Technical Reports Server (NTRS)

Berrier, B. L.; Mason, M. L.

1983-01-01

The flow-turning capability and nozzle internal performance of yaw-vectoring nozzle geometries were tested in the NASA Langley 16-ft Transonic wind tunnel. The concept was investigated as a means of enhancing fighter jet performance. Five two-dimensional convergent-divergent nozzles were equipped for yaw-vectoring and examined. The configurations included a translating left sidewall, left and right sidewall flaps downstream of the nozzle throat, left sidewall flaps or port located upstream of the nozzle throat, and a powered rudder. Trials were also run with 20 deg of pitch thrust vectoring added. The feasibility of providing yaw-thrust vectoring was demonstrated, with the largest yaw vector angles being obtained with sidewall flaps downstream of the nozzle primary throat. It was concluded that yaw vector designs that scoop or capture internal nozzle flow provide the largest yaw-vector capability, but decrease the thrust the most.

A new generation of high performance engines for spacecraft propulsion

NASA Technical Reports Server (NTRS)

Rosenberg, Sanders D.; Schoenman, Leonard

1991-01-01

Experimental data validating advanced engine designs at three thrust levels (5, 15, and 100 lbF) is presented. All of the three engine designs considered employ a Moog bipropellant torque motor valve, platelet injector design, and iridium-lined rhenium combustion chamber. Attention is focused on the performance, robustness, duration, and flexibility characteristics of the engines. It is noted that the 5- and 15-lbF thrust engines can deliver a steady state specific impulse in excess of 310 lbF-sec/lbm at an area ratio of 150:1, while the 150-lbF thrust engines deliver a steady state specific impulse of 320 lbF-sec/lbm at an area ratio of 250:1. The hot-fire test results reveal specific impulse improvements of 15 to 25 sec over conventional fuel film cooled columbium chamber designs while operating at maximum chamber temperatures.

A performance comparison of two small rocket nozzles

NASA Technical Reports Server (NTRS)

Arrington, Lynn A.; Reed, Brian D.; Rivera, Angel, Jr.

1996-01-01

An experimental study was conducted on two small rockets (110 N thrust class) to directly compare a standard conical nozzle with a bell nozzle optimized for maximum thrust using the Rao method. In large rockets, with throat Reynolds numbers of greater than 1 x 10(exp 5), bell nozzles outperform conical nozzles. In rockets with throat Reynolds numbers below 1 x 10(exp 5), however, test results have been ambiguous. An experimental program was conducted to test two small nozzles at two different fuel film cooling percentages and three different chamber pressures. Test results showed that for the throat Reynolds number range from 2 x 10(exp 4) to 4 x 10(exp 4), the bell nozzle outperformed the conical nozzle. Thrust coefficients for the bell nozzle were approximately 4 to 12 percent higher than those obtained with the conical nozzle. As expected, testing showed that lowering the fuel film cooling increased performance for both nozzle types.

Experiments on high speed ejectors

NASA Technical Reports Server (NTRS)

Wu, J. J.

1986-01-01

Experimental studies were conducted to investigate the flow and the performance of thrust augmenting ejectors for flight Mach numbers in the range of 0.5 to 0.8, primary air stagnation pressures up to 107 psig (738 kPa), and primary air stagnation temperatures up to 1250 F (677 C). The experiment verified the existence of the second solution ejector flow, where the flow after complete mixing is supersonic. Thrust augmentation in excess of 1.2 was demonstrated for both hot and cold primary jets. The experimental ejector performed better than the corresponding theoretical optimal first solution ejector, where the mixed flow is subsonic. Further studies are required to realize the full potential of the second solution ejector. The research program was started by the Flight Dynamics Research Corporation (FDRC) to investigate the characteristic of a high speed ejector which augments thrust of a jet at high flight speeds.

Extended operating range of the 30-cm ion thruster with simplified power processor requirements

NASA Technical Reports Server (NTRS)

Rawlin, V. K.

1981-01-01

A two grid 30 cm diameter mercury ion thruster was operated with only six power supplies over the baseline J series thruster power throttle range with negligible impact on thruster performance. An analysis of the functional model power processor showed that the component mass and parts count could be reduced considerably and the electrical efficiency increased slightly by only replacing power supplies with relays. The input power, output thrust, and specific impulse of the thruster were then extended, still using six supplies, from 2660 watts, 0.13 newtons, and 2980 seconds to 9130 watts, 0.37 newtons, and 3820 seconds, respectively. Increases in thrust and power density enable reductions in the number of thrusters and power processors required for most missions. Preliminary assessments of the impact of thruster operation at increased thrust and power density on the discharge characteristics, performance, and lifetime of the thruster were also made.

Orbit Transfer Vehicle Engine Study. Phase A, extension 1: Advanced expander cycle engine optimization

NASA Technical Reports Server (NTRS)

Mellish, J. A.

1979-01-01

The performance optimization of expander cycle engines at vacuum thrust levels of 10K, 15K, and 20K lb is discussed. The optimization is conducted for a maximum engine length with an extendible nozzle in the retracted position of 60 inches and an engine mixture ratio of 6.0:1. The thrust chamber geometry and cycle analyses are documented. In addition, the sensitivity of a recommended baseline expander cycle to component performance variations is determined and chilldown/start propellant consumptions are estimated.

Monomethylhydrazine versus hydrazine fuels - Test results using a 100 pound thrust bipropellant rocket engine

NASA Technical Reports Server (NTRS)

Smith, J. A.; Stechman, R. C.

1981-01-01

A test program was performed to evaluate hydrazine (N2H4) as a fuel for a 445 Newton (100 lbf) thrust bipropellant rocket engine. Results of testing with an identical thruster utilizing monomethylhydrazine (MMH) are included for comparison. Engine performance with hydrazine fuel was essentially identical to that experienced with monomethylhydrazine although higher combustor wall temperatures (approximately 400 F) were obtained with hydrazine. Results are presented which indicate that hydrazine as a fuel is compatible with Marquardt bipropellant rocket engines which use monomethylhydrazine as a baseline fuel.

A new concept for a hubless rotary jet

NASA Technical Reports Server (NTRS)

Garris, C. A.; Toh, K. H.; Xie, L.

1991-01-01

The 'hubless rotary-jet' thrust augmentor primary-stream configuration is proposed as a way of overcoming problems encountered with hubbed rotary-jet ejectors while retaining sufficient geometrical simplicity to facilitate performance characterization and prediction via such commercially available CFD methods as FLOTRAN, as well as verification via LDV mapping. The present discussion is conducted with a view to the contributions of more realistic performance models incorporating turbulent mixing and compressibility. Attention is given to thrust augmentation ratio vs. spin angles for various area ratios with and without mixing.

Effect of rotor design tip seed on aerodynamic performance of a model VTOL lift fan under static and crossflow conditions

NASA Technical Reports Server (NTRS)

Stockman, N. O.; Loeffler, I. J.; Lieblein, S.

1973-01-01

Results are presented for a wind tunnel investigation of three single VTOL lift fan stages designed for the same overall total pressure ratio at different rotor tip speeds. The stages were tested in a model lift fan installed in a wing pod. The three stages had essentially the same aerodynamic performance along the operating line. However, differences in stage thrust characteristics were obtained when a variation in back pressure was imposed on the stages by cross-flow effects and thrust-vectoring louvers.

Performance of 4600-pound-thrust centrifugal-flow-type turbojet engine with water-alcohol injection at inlet

NASA Technical Reports Server (NTRS)

Glasser, Philip W

1950-01-01

An experimental investigation of the effects of injecting a water-alcohol mixture of 2:1 at the compressor inlet of a centrifugal-flow type turbojet engine was conducted in an altitude test chamber at static sea-level conditions and at an altitude of 20,000 feet with a flight Mach number of 0.78 with an engine operating at rated speed. The net thrust was augmented by 0.16 for both flight conditions with a ratio of injected liquid to air flow of 0.05. Further increases in the liquid-air ratio did not give comparable increases in thrust.

Earth orbital assessment of solar electric and solar sail propulsion systems

NASA Technical Reports Server (NTRS)

Teeter, R. R.

1977-01-01

The earth orbital applications potential of Solar Electric (Ion Drive) and Solar Sail low-thrust propulsion systems are evaluated. Emphasis is placed on mission application in the 1980s. The two low-thrust systems are compared with each other and with two chemical propulsion Shuttle upper stages (the IUS and SSUS) expected to be available in the 1980s. The results indicate limited Earth orbital application potential for the low-thrust systems in the 1980s (primarily due to cost disadvantages). The longer term potential is viewed as more promising. Of the two systems, the Ion Drive exhibits better performance and appears to have better overall application potential.

Developmental Testing of Electric Thrust Vector Control Systems for Manned Launch Vehicle Applications

NASA Technical Reports Server (NTRS)

Bates, Lisa B.; Young, David T.

2012-01-01

This paper describes recent developmental testing to verify the integration of a developmental electromechanical actuator (EMA) with high rate lithium ion batteries and a cross platform extensible controller. Testing was performed at the Thrust Vector Control Research, Development and Qualification Laboratory at the NASA George C. Marshall Space Flight Center. Electric Thrust Vector Control (ETVC) systems like the EMA may significantly reduce recurring launch costs and complexity compared to heritage systems. Electric actuator mechanisms and control requirements across dissimilar platforms are also discussed with a focus on the similarities leveraged and differences overcome by the cross platform extensible common controller architecture.

Feasibility of rotating fluidized bed reactor for rocket propulsion

NASA Technical Reports Server (NTRS)

Ludewig, H.; Manning, A. J.; Raseman, C. J.

1974-01-01

The rotating fluidized bed reactor concept is outlined, and its application to rocket propulsion is discussed. Experimental results obtained indicate that minimum fluidization correlations commonly in use for 1-g beds can also be applied to multiple-g beds. It was found that for a low thrust system (20,000 lbf) the fuel particle size and/or particle stress play a limiting role on performance. The superiority of U-233 as a fuel for this type of rocket engine is clearly demonstrated in the analysis. The maximum thrust/weight ratio for a 90,000N thrust engine was found to be approximately 65N/kg.

NASA Orbit Transfer Rocket Engine Technology Program

NASA Technical Reports Server (NTRS)

1984-01-01

The advanced expander cycle engine with a 15,000 lb thrust level and a 6:1 mixture ratio and optimized performance was used as the baseline for a design study of the hydrogen/oxgyen propulsion system for the orbit transfer vehicle. The critical components of this engine are the thrust chamber, the turbomachinery, the extendible nozzle system, and the engine throttling system. Turbomachinery technology is examined for gears, bearing, seals, and rapid solidification rate turbopump shafts. Continuous throttling concepts are discussed. Components of the OTV engine described include the thrust chamber/nozzle assembly design, nozzles, the hydrogen regenerator, the gaseous oxygen heat exchanger, turbopumps, and the engine control valves.

Automatic control of a primary electric thrust subsystem

NASA Technical Reports Server (NTRS)

Macie, T. W.; Macmedan, M. L.

1975-01-01

A concept for automatic control of the thrust subsystem has been developed by JPL and participating NASA Centers. This paper reports on progress in implementing the concept at JPL. Control of the Thrust Subsystem (TSS) is performed by the spacecraft computer command subsystem, and telemetry data is extracted by the spacecraft flight data subsystem. The Data and Control Interface Unit, an element of the TSS, provides the interface with the individual elements of the TSS. The control philosophy and implementation guidelines are presented. Control requirements are listed, and the control mechanism, including the serial digital data intercommunication system, is outlined. The paper summarizes progress to Fall 1974.

Development of a CFRP Engine Thrust Frame for the Next Generation Launchers

NASA Astrophysics Data System (ADS)

Fatemi, Javad; van der Bas, Finn; Cruijssen, Henk

2012-07-01

This paper addresses the activities related to the development of technologies for a composite Engine Thrust Frame (ETF) for the next generation launchers. In particular, the design and analyses of a full Carbon Fibre Reinforced Plastic (CFRP) engine thrust frame are presented in more detail. The ETF concept is composed of three main parts, i.e. an aluminium top-ring which connects the ETF to the upper-stage tank, a CFRP cone, and a CFRP cone-cap which connects the Vinci engine to the ETF. The main challenging requirements for development of a CFRP ETF are recalled. The ETF concept and its mechanical performances are assessed.

Orbit Transfer Vehicle Engine Study. Phase A, extension 1: Alternate low-thrust capability task report

NASA Technical Reports Server (NTRS)

Mellish, J. A.

1980-01-01

The feasibility and design impact of a requirement for the advanced expander cycle engine to be adaptable to extended low thrust operation of approximately 1K to 2K lb is assessed. It is determined that the orbit transfer vehicle point design engine can be reduced in thrust with minor injector modifications from 15K to 1K without significantly affecting combustion performance efficiency or injector face/chamber wall thermal compatibility. Likewise, high frequency transverse mode combustion instability is not expected to be detrimentally affected. Primarily, the operational limitations consist of feed system chugging instabilities and potential coupling of the injector response with the chamber longitudinal mode resonances under certain operating conditions. The recommended injector modification for low thrust operation is a change in the oxidizer injector element orifice size. Analyses also indicate that chamber coolant flow stability may be a concern below 2K 1bF operation and oxidizer pump stability could be a problem below a 2K thrust level although a recirculation flow could alleviate the problem.

Vector thrust induced lift effects for several ejector exhaust locations on a V/STOL wind tunnel model at forward speed

NASA Technical Reports Server (NTRS)

Sharon, A. D.

1975-01-01

The results and analysis of aerodynamic force data obtained from a small scale model of a V/STOL research vehicle in a low speed wind tunnel are presented. The analysis of the data includes the evaluation of aerodynamic-propulsive lift performance when operating twin ejector nozzles with thrust deflected. Three different types of thrust deflector systems were examined: 90 deg downward deflected nozzle, 90 deg slotted nozzle with boundary layer control, and an externally blown flap configuration. Several nozzle locations were tested, including over and underwing positions. The interference lift of the nacelle and model due to jet exhaust thrust is compared and results show that 90 deg turned nozzles located over the wing (near the trailing edge) produce the largest interference lift increment for an untrimmed aircraft, and that the slotted nozzle located under the wing near the trailing edge (in conjunction with a BLC flap) gives a comparable interference lift in the trimmed condition. The externally blown flap nozzle produced the least interference lift and significantly less total lift due to jet thrust effects.

Numerical Simulation of Cylindrical, Self-field MPD Thrusters with Multiple Propellants

NASA Technical Reports Server (NTRS)

Lapointe, Michael R.

1994-01-01

A two-dimensional, two-temperature, single fluid MHD code was used to predict the performance of cylindrical, self-field magnetoplasmadynamic (MPD) thrusters operated with argon, lithium, and hydrogen propellants. A thruster stability equation was determined relating maximum stable J(sup 2)/m values to cylindrical thruster geometry and propellant species. The maximum value of J(sup 2)/m was found to scale as the inverse of the propellant molecular weight to the 0.57 power, in rough agreement with limited experimental data which scales as the inverse square root of the propellant molecular weight. A general equation which relates total thrust to electromagnetic thrust, propellant molecular weight, and J(sup 2)/m was determined using reported thrust values for argon and hydrogen and calculated thrust values for lithium. In addition to argon, lithium, and hydrogen, the equation accurately predicted thrust for ammonia at sufficiently high J(sup 2)/m values. A simple algorithm is suggested to aid in the preliminary design of cylindrical, self-field MPD thrusters. A brief example is presented to illustrate the use of the algorithm in the design of a low power MPD thruster.

Performance characterization of a permanent-magnet helicon plasma thruster

NASA Astrophysics Data System (ADS)

Takahashi, Kazunori; Charles, Christine; Boswell, Rod

2012-10-01

Helicon plasma thrusters operated at a few kWs of rf power is an active area of an international research. Recent experiments have clarified part of the thrust-generation mechanisms. Thrust components which have been identified include an electron pressure inside the source region and a Lorentz force due to an electron diamagnetic drift current and a radial component of the applied magnetic field. The use of permanent magnets (PMs) instead of solenoids is one of the solutions for improving the thruster efficiency because it does not require electricity for the magnetic nozzle formation. Here the thrust imparted from a permanent-magnet helicon plasma thruster is directly measured using a pendulum thrust balance. The source consists of permanent magnet (PM) arrays, a double turn rf loop antenna powered by a 13.56 MHz rf generator and a glass source tube. The PM arrays provide a magnetic nozzle near the open exit of the source and two configurations, which have maximum field strengths of about 100 and 270 G, are tested. A thrust of 15 mN, specific impulse of 2000 sec and a thrust efficiency of 8 percent are presently obtained for 2 kW of input power, 24 sccm flow rate of argon and the stronger magnetic field configuration.

Investigation on the Effect of Drill Geometry and Pilot Holes on Thrust Force and Burr Height When Drilling an Aluminium/PE Sandwich Material

PubMed Central

Rezende, Bruna Aparecida; Silveira, Michele L.; Vieira, Luciano M. G.; Abrão, Alexandre M.; de Faria, Paulo Eustáquio; Rubio, Juan C. Campos

2016-01-01

Composite materials are widely employed in the naval, aerospace and transportation industries owing to the combination of being lightweight and having a high modulus of elasticity, strength and stiffness. Drilling is an operation generally used in composite materials to assemble the final product. Damages such as the burr at the drill entrance and exit, geometric deviations and delamination are typically found in composites subjected to drilling. Drills with special geometries and pilot holes are alternatives used to improve hole quality as well as to increase tool life. The present study is focused on the drilling of a sandwich composite material (two external aluminum plates bound to a polyethylene core). In order to minimize thrust force and burr height, the influence of drill geometry, the pilot hole and the cutting parameters was assessed. Thrust force and burr height values were collected and used to perform an analysis of variance. The results indicated that the tool and the cutting speed were the parameters with more weight on the thrust force and for burr height they were the tool and the interaction between tool and feed. The results indicated that drilling with a pilot hole of Ø4 mm exhibited the best performance with regard to thrust force but facilitated plastic deformation, thus leading to the elevation of burr height, while the lowest burr height was obtained using the Brad and Spur drill geometry. PMID:28773895

Investigation on the Effect of Drill Geometry and Pilot Holes on Thrust Force and Burr Height When Drilling an Aluminium/PE Sandwich Material.

PubMed

Rezende, Bruna Aparecida; Silveira, Michele L; Vieira, Luciano M G; Abrão, Alexandre M; Faria, Paulo Eustáquio de; Rubio, Juan C Campos

2016-09-13

Composite materials are widely employed in the naval, aerospace and transportation industries owing to the combination of being lightweight and having a high modulus of elasticity, strength and stiffness. Drilling is an operation generally used in composite materials to assemble the final product. Damages such as the burr at the drill entrance and exit, geometric deviations and delamination are typically found in composites subjected to drilling. Drills with special geometries and pilot holes are alternatives used to improve hole quality as well as to increase tool life. The present study is focused on the drilling of a sandwich composite material (two external aluminum plates bound to a polyethylene core). In order to minimize thrust force and burr height, the influence of drill geometry, the pilot hole and the cutting parameters was assessed. Thrust force and burr height values were collected and used to perform an analysis of variance. The results indicated that the tool and the cutting speed were the parameters with more weight on the thrust force and for burr height they were the tool and the interaction between tool and feed. The results indicated that drilling with a pilot hole of Ø4 mm exhibited the best performance with regard to thrust force but facilitated plastic deformation, thus leading to the elevation of burr height, while the lowest burr height was obtained using the Brad and Spur drill geometry.

Experimental investigation of SDBD plasma actuator driven by AC high voltage with a superimposed positive pulse bias voltage

NASA Astrophysics Data System (ADS)

Qi, Xiao-Hua; Yan, Hui-Jie; Yang, Liang; Hua, Yue; Ren, Chun-Sheng

2017-08-01

In this work, a driven voltage consisting of AC high voltage with a superimposed positive pulse bias voltage ("AC+ Positive pulse bias" voltage) is adopted to study the performance of a surface dielectric barrier discharge plasma actuator under atmospheric conditions. To compare the performance of the actuator driven by single-AC voltage and "AC+ Positive pulse bias" voltage, the actuator-induced thrust force and power consumption are measured as a function of the applied AC voltage, and the measured results indicate that the thrust force can be promoted significantly after superimposing the positive pulse bias voltage. The physical mechanism behind the thrust force changes is analyzed by measuring the optical properties, electrical characteristics, and surface potential distribution. Experimental results indicate that the glow-like discharge in the AC voltage half-cycle, next to the cycle where a bias voltage pulse has been applied, is enhanced after applying the positive pulse bias voltage, and this perhaps is the main reason for the thrust force increase. Moreover, surface potential measurement results reveal that the spatial electric field formed by the surface charge accumulation after positive pulse discharge can significantly affect the applied external electric field, and this perhaps can be responsible for the experimental phenomenon that the decrease of thrust force is delayed by pulse bias voltage action after the filament discharge occurs in the glow-like discharge region. The schlieren images further verify that the actuator-induced airflow velocity increases with the positive pulse voltage.

A Method for Integrating Thrust-Vectoring and Actuated Forebody Strakes with Conventional Aerodynamic Controls on a High-Performance Fighter Airplane

NASA Technical Reports Server (NTRS)

Lallman, Frederick J.; Davidson, John B.; Murphy, Patrick C.

1998-01-01

A method, called pseudo controls, of integrating several airplane controls to achieve cooperative operation is presented. The method eliminates conflicting control motions, minimizes the number of feedback control gains, and reduces the complication of feedback gain schedules. The method is applied to the lateral/directional controls of a modified high-performance airplane. The airplane has a conventional set of aerodynamic controls, an experimental set of thrust-vectoring controls, and an experimental set of actuated forebody strakes. The experimental controls give the airplane additional control power for enhanced stability and maneuvering capabilities while flying over an expanded envelope, especially at high angles of attack. The flight controls are scheduled to generate independent body-axis control moments. These control moments are coordinated to produce stability-axis angular accelerations. Inertial coupling moments are compensated. Thrust-vectoring controls are engaged according to their effectiveness relative to that of the aerodynamic controls. Vane-relief logic removes steady and slowly varying commands from the thrust-vectoring controls to alleviate heating of the thrust turning devices. The actuated forebody strakes are engaged at high angles of attack. This report presents the forward-loop elements of a flight control system that positions the flight controls according to the desired stability-axis accelerations. This report does not include the generation of the required angular acceleration commands by means of pilot controls or the feedback of sensed airplane motions.

Simple Sensitivity Analysis for Orion Guidance Navigation and Control

NASA Technical Reports Server (NTRS)

Pressburger, Tom; Hoelscher, Brian; Martin, Rodney; Sricharan, Kumar

2013-01-01

The performance of Orion flight software, especially its GNC software, is being analyzed by running Monte Carlo simulations of Orion spacecraft flights. The simulated performance is analyzed for conformance with flight requirements, expressed as performance constraints. Flight requirements include guidance (e.g. touchdown distance from target) and control (e.g., control saturation) as well as performance (e.g., heat load constraints). The Monte Carlo simulations disperse hundreds of simulation input variables, for everything from mass properties to date of launch. We describe in this paper a sensitivity analysis tool ("Critical Factors Tool" or CFT) developed to find the input variables or pairs of variables which by themselves significantly influence satisfaction of requirements or significantly affect key performance metrics (e.g., touchdown distance from target). Knowing these factors can inform robustness analysis, can inform where engineering resources are most needed, and could even affect operations. The contributions of this paper include the introduction of novel sensitivity measures, such as estimating success probability, and a technique for determining whether pairs of factors are interacting dependently or independently. The tool found that input variables such as moments, mass, thrust dispersions, and date of launch were found to be significant factors for success of various requirements. Examples are shown in this paper as well as a summary and physics discussion of EFT-1 driving factors that the tool found.

Flight Test of the Aerojet 7KS-6000 T-27 Jato Rocket Motor

NASA Technical Reports Server (NTRS)

Bond, Aleck C.; Thibodaux, Joseph G., Jr.

1949-01-01

A flight test of the Aero jet Engineering Corporation's 7KS-6000 T-27 Jato rocket motor was conducted at the Langley Pilotless Aircraft Research Station at Wallops Island, Va, to determine the flight performance characteristics of the motor. The flight test imposed an absolute longitudinal acceleration of 9.8 g upon the rocket motor at 2.8 seconds after launching. The total impulse developed by the motor was 43,400 pound-seconds, and the thrusting time was 7.58 seconds. The maximum thrust was 7200 pounds and occurred at 4.8 seconds after launching. No thrust irregularities attributable to effects of the flight longitudinal acceleration were observed. Certain small thrust irregularities occurred in the flight test which appear to correspond to irregularities observed in static tests conducted elsewhere. A hypothesis regarding the origin of these small irregularities is presented.

A zonal computational procedure adapted to the optimization of two-dimensional thrust augmentor inlets

NASA Technical Reports Server (NTRS)

Lund, T. S.; Tavella, D. A.; Roberts, L.

1985-01-01

A viscous-inviscid interaction methodology based on a zonal description of the flowfield is developed as a mean of predicting the performance of two-dimensional thrust augmenting ejectors. An inviscid zone comprising the irrotational flow about the device is patched together with a viscous zone containing the turbulent mixing flow. The inviscid region is computed by a higher order panel method, while an integral method is used for the description of the viscous part. A non-linear, constrained optimization study is undertaken for the design of the inlet region. In this study, the viscous-inviscid analysis is complemented with a boundary layer calculation to account for flow separation from the walls of the inlet region. The thrust-based Reynolds number as well as the free stream velocity are shown to be important parameters in the design of a thrust augmentor inlet.

X-31 quasi-tailless flight demonstration

NASA Technical Reports Server (NTRS)

Huber, Peter; Schellenger, Harvey G.

1994-01-01

The primary objective of the quasi-tailless flight demonstration is to demonstrate the feasibility of using thrust vectoring for directional control of an unstable aircraft. By using this low-cost, low-risk approach it is possible to get information about required thrust vector control power and deflection rates from an inflight experiment as well as insight in low-power thrust vectoring issues. The quasi-tailless flight demonstration series with the X-31 began in March 1994. The demonstration flight condition was Mach 1.2 at 37,500 feet. A series of basic flying quality maneuvers, doublets, bank to bank rolls, and wind-up-turns have been performed with a simulated 100% vertical tail reduction. Flight test and supporting simulation demonstrated that the quasi-tailless approach is effective in representing the reduced stability of tailless configurations. The flights also demonstrated that thrust vectoring could be effectively used to stabilize a directionally unstable configuration and provide control power for maneuver coordination.

An analysis of thrust of a realistic solar sail with focus on a flight validation mission in a geocentric orbit

NASA Astrophysics Data System (ADS)

Campbell, Bruce A.

Several scientifically important space flight missions have been identified that, at this time, can only be practically achieved using a solar sail propulsion system. These missions take advantage of the potentially continuous force on the sail, provided by solar radiation, to produce significant changes in the spacecraft's velocity, in both magnitude and/or direction, without the need for carrying the enormous amount of fuel that conventional propulsion systems would require to provide the same performance. However, to provide thrust levels that would support these missions requires solar sail areas in the (tens of) thousands of square meter sizes. To realize this, many technical areas must be developed further and demonstrated in space before solar sails will be accepted as a viable space mission propulsion system. One of these areas concerns understanding the propulsion performance of a realistic solar sail well enough for mission planning. Without this understanding, solar sail orbits could not be predicted well enough to meet defined mission requirements, such as rendezvous or station-keeping, and solar sail orbit optimization, such as minimizing flight time, could be close to impossible. In most mission studies, either an "ideal" sail's performance is used for mission planning, or some top-level assumptions of certain nonideal sail characteristics are incorporated to give a slightly better estimate of the sail performance. This paper identifies the major sources of solar sail thrust performance uncertainty, and analyzes the most significant ones to provide a more comprehensive understanding of thrust generation by a "realistic" solar sail. With this understanding, mission planners will be able to more confidently and accurately estimate the capabilities of such a system. The first solar sail mission will likely be a system validation mission, using a relatively small sail in a geocentric (Earth-centered) orbit. The author has been involved in conceptual design of such missions, and through this became aware of the current status in solar sail system development, and the need for a better understanding of the thrust performance of a "realistic" solar sail. Such a validation mission is significantly different than most of the "operational" science missions envisioned to utilize a solar sail propulsion system. These future missions will likely use very large, very light sails in heliocentric orbits far away from major gravity fields like planets, have very long mission lifetimes (years), and will conduct relatively minor and slow orbital and attitude control maneuvers. Nonetheless, most of the capabilities of later systems can be gleaned from a small geocentric validation mission. This paper is a significant step toward understanding the thrust characteristics and performance of a realistic solar sail, and provides insight to the methods by which this understanding can be corroborated by a solar sail validation mission.

Feasibility of Reusable Continuous Thrust Spacecraft for Cargo Resupply Missions to Mars

NASA Astrophysics Data System (ADS)

Rabotin, C. B.

Continuous thrust propulsion systems benefit from a much greater efficiency in vacuum than chemical rockets, at the expense of lower instantaneous thrust and high power requirements. The satellite telecommunications industry, known for greatly emphasizing heritage over innovation, now uses electric propulsion for station keeping on a number of spacecraft, and for orbit raising for some smaller satellites, such as the Boeing 702SP platform. Only a few interplanetary missions have relied on continuous thrust for most of their mission, such as ESA's 367 kg SMART-1 and NASA's 1217 kg Dawn mission. The high specific impulse of these continuous thrust engines should make them suitable for transportation of heavy payloads to inner solar system destinations in such a way to limit the dependency on heavy rocket launches. Additionally, such spacecraft should be able to perform orbital insertions at destination in order to deliver the cargo directly in a desired orbit. An example application is designing round-trip missions to Mars to support exploration and eventually colonization. This research investigates the feasibility of return journeys to Mars based on the performance of existing or in-development continuous thrust propulsion systems. In order to determine the business viability of such missions, an emphasis is made on the time of flight during different parts of the mission, the relative velocity with respect to the destination planet, and the fuel requirements. The study looks at the applicability for interplanetary mission design of simple control laws for efficient correction of orbital elements, and of thrusting purely in velocity or anti-velocity direction. The simulations explore different configurations of continuous thrusting technologies using a patched-conics approach. In addition, all simulation scenarios facilitate escape from planetary gravity wells as the initial spacecraft orbit is highly elliptical, both around the Earth and around Mars. This work does not include any optimal trajectory design. For this research, a highly configurable orbit propagation software with SPICE ephemerides was developed from scratch in Go, a modern compiled computer language. The outcome of this research is that simple orbital element control laws do not lead to more efficient or faster interplanetary transfers. In addition, spiraling out of Earth's gravity wells requires a substantial amount of time despite starting from a highly elliptical orbit, and even with clustered high thrust engines like the VASIMR VX-200. Further investigation should look into hybrid solutions with a chemical engine for departing Earth; outbound spirals from Mars take a more reasonable amount of time.

An antiproton driver for ICF propulsion

NASA Technical Reports Server (NTRS)

Chiang, Pi-Ren; Lewis, R. A.; Smith, G. A.; Gazze, C.; Higman, K.; Newton, R.; Chiaverini, M.; Dailey, J.; Surratt, M.; Werthman, W. Lance

1993-01-01

Inertial confinement fusion (ICF) utilizing an anitprotoncatalyzed target is discussed as a possible source of propulsion for rapid interplanetary manned space missions. The relevant compression, ignition, and thrust mechanisms are presented. Progress on an experiment presently in progress at the Phillips Laboratory, Kirtland AFB, NM to demonstrate proof-of-principle is reviewed.

77 FR 51724 - Airworthiness Directives; The Boeing Company

Federal Register 2010, 2011, 2012, 2013, 2014

2012-08-27

... ON Thrust Levers CLOSE Speed Brakes FLIGHT DETENT Target Speed MO/MMO'' Note 1 to paragraphs (j)(2... W. Palmer, Aerospace Engineer, Systems and Equipment Branch, ANM-130S, FAA, Seattle Aircraft... charges the FAA with promoting safe flight of civil aircraft in air commerce by prescribing regulations...

Electromyographic Comparison of Barbell Deadlift, Hex Bar Deadlift, and Hip Thrust Exercises: A Cross-Over Study.

PubMed

Andersen, Vidar; Fimland, Marius S; Mo, Dag-Andrè; Iversen, Vegard M; Vederhus, Torbjørn; Rockland Hellebø, Lars R; Nordaune, Kristina I; Saeterbakken, Atle H

2018-03-01

Andersen, V, Fimland, MS, Mo, D-A, Iversen, VM, Vederhus, T, Rockland Hellebø, LR, Nordaune, KI, and Saeterbakken, AH. Electromyographic comparison of barbell deadlift, hex bar deadlift, and hip thrust exercises: a cross-over study. J Strength Cond Res 32(3): 587-593, 2018-The aim of the study was to compare the muscle activation level of the gluteus maximus, biceps femoris, and erector spinae in the hip thrust, barbell deadlift, and hex bar deadlift; each of which are compound resisted hip extension exercises. After 2 familiarization sessions, 13 resistance-trained men performed a 1 repetition maximum in all 3 exercises in 1 session, in randomized and counterbalanced order. The whole ascending movement (concentric phase), as well as its lower and upper parts (whole movement divided in 2), were analyzed. The hip thrust induced greater activation of the gluteus maximus compared with the hex bar deadlift in the whole (16%, p = 0.025) and the upper part (26%, p = 0.015) of the movement. For the whole movement, the biceps femoris was more activated during barbell deadlift compared with both the hex bar deadlift (28%, p < 0.001) and hip thrust (20%, p = 0.005). In the lower part of the movement, the biceps femoris activation was, respectively, 48% and 26% higher for the barbell deadlift (p < 0.001) and hex bar deadlift (p = 0.049) compared with hip thrust. Biceps femoris activation in the upper part of the movement was 39% higher for the barbell deadlift compared with the hex bar deadlift (p = 0.001) and 34% higher for the hip thrust compared with the hex bar deadlift (p = 0.002). No differences were displayed for the erector spinae activation (p = 0.312-0.859). In conclusion, the barbell deadlift was clearly superior in activating the biceps femoris compared with the hex bar deadlift and hip thrust, whereas the hip thrust provided the highest gluteus maximus activation.

Associations of varus thrust and alignment with pain in knee osteoarthritis.

PubMed

Lo, Grace H; Harvey, William F; McAlindon, Timothy E

2012-07-01

To investigate associations of varus thrust and varus static alignment with pain in patients with knee osteoarthritis (OA). This was a cross-sectional study of participants from a randomized controlled trial of vitamin D treatment for knee OA. Participants were video recorded while walking and scored for presence of varus thrust. Static alignment was measured on standard posteroanterior knee radiographs. Pain questions from the Western Ontario and McMaster Universities Osteoarthritis Index (WOMAC) questionnaire were used to assess symptoms. We calculated means for total WOMAC pain in relation to varus thrust and static varus alignment (i.e., corrected anatomic alignment<178 degrees). Ordinal logistic regressions were performed, with responses on individual WOMAC pain questions as the outcomes and varus thrust and varus alignment as the predictors. There were 82 participants, 60% of whom were female. The mean±SD age was 65.1±8.5 years, and the mean±SD body mass index was 30.2±5.4 kg/m2. The mean total WOMAC pain score was 6.3 versus 3.9, respectively, in those with versus without definite varus thrust (P=0.007) and 5.0 versus 4.2 in those with versus without varus alignment (P=0.36). Odds ratios for pain with walking and standing were 4.7 (95% confidence interval 1.8-11.9) and 5.5 (95% confidence interval 2.2-14.2), respectively, in those with and those without definite varus thrust. There were no significant associations between varus alignment and responses to individual WOMAC pain questions. Sensitivity analyses suggested that varus classified using a more stringent definition might have been associated with pain on walking and standing. In patients with knee OA, varus thrust, and possibly varus static alignment, were associated with pain, specifically during weight-bearing activities. Treatment of varus thrust (e.g., via bracing or gait modification) may lead to improvement of symptoms. Copyright © 2012 by the American College of Rheumatology.

High-Energy Space Propulsion Based on Magnetized Target Fusion

NASA Technical Reports Server (NTRS)

Thio, Y. C. F.; Freeze, B.; Kirkpatrick, R. C.; Landrum, B.; Gerrish, H.; Schmidt, G. R.

1999-01-01

A conceptual study is made to explore the feasibility of applying magnetized target fusion (MTF) to space propulsion for omniplanetary travel. Plasma-jet driven MTF not only is highly amenable to space propulsion, but also has a number of very attractive features for this application: 1) The pulsed fusion scheme provides in situ a very dense hydrogenous liner capable of moderating the neutrons, converting more than 97% of the neutron energy into charged particle energy of the fusion plasma available for propulsion. 2) The fusion yield per pulse can be maintained at an attractively low level (< 1 GJ) despite a respectable gain in excess of 70. A compact, low-weight engine is the result. An engine with a jet power of 25 GW, a thrust of 66 kN, and a specific impulse of 77,000 s, can be achieved with an overall engine mass of about 41 metric tons, with a specific power density of 605 kW/kg, and a specific thrust density of 1.6 N/kg. The engine is rep-rated at 40 Hz to provide this power and thrust level. At a practical rep-rate limit of 200 Hz, the engine can deliver 128 GW jet power and 340 kN of thrust, at specific power and thrust density of 1,141 kW/kg and 3 N/kg respectively. 3) It is possible to operate the magnetic nozzle as a magnetic flux compression generator in this scheme, while attaining a high nozzle efficiency of 80% in converting the spherically radial momentum of the fusion plasma to an axial impulse. 4) A small fraction of the electrical energy generated from the flux compression is used directly to recharge the capacitor bank and other energy storage equipment, without the use of a highvoltage DC power supply. A separate electrical generator is not necessary. 5) Due to the simplicity of the electrical circuit and the components, involving mainly inductors, capacitors, and plasma guns, which are connected directly to each other without any intermediate equipment, a high rep-rate (with a maximum of 200 Hz) appears practicable. 6) All fusion related components are within the current state of the art for pulsed power technology. Experimental facilities with the required pulsed power capabilities already exist. 7) The scheme does not require prefabricated fuel target and liner hardware in any esoteric form or state. All necessary fuel and liner material are introduced into the engine in the form of ordinary matter in gaseous state at room temperature, greatly simplifying their handling on board. They are delivered into the fusion reaction chamber in a completely standoff manner.

Nozzle design study for a quasi-axisymmetric scramjet-powered vehicle at Mach 7.9 flight conditions

NASA Astrophysics Data System (ADS)

Tanimizu, Katsuyoshi; Mee, David J.; Stalker, Raymond J.; Jacobs, Peter A.

2013-09-01

A nozzle shape optimization study for a quasi-axisymmetric scramjet has been performed for a Mach 7.9 operating condition with hydrogen fuel, aiming at the application of a hypersonic airbreathing vehicle. In this study, the nozzle geometry which is parameterized by a set of design variables, is optimized for the single objective of maximum net thrust using an in-house CFD solver for inviscid flowfields with a simple force prediction methodology. The combustion is modelled using a simple chemical reaction code. The effects of the nozzle design on the overall vehicle performance are discussed. For the present geometry, net thrust is achieved for the optimized vehicle design. The results of the nozzle-optimization study show that performance is limited by the nozzle area ratio that can be incorporated into the vehicle without leading to too large a base diameter of the vehicle and increasing the external drag of the vehicle. This study indicates that it is very difficult to achieve positive thrust at Mach 7.9 using the basic geometry investigated.

Oxygen-hydrogen thrusters for Space Station auxiliary propulsion systems

NASA Technical Reports Server (NTRS)

Berkman, D. K.

1984-01-01

The feasibility and technology requirements of a low-thrust, high-performance, long-life, gaseous oxygen (GO2)/gaseous hydrogen (GH2) thruster were examined. Candidate engine concepts for auxiliary propulsion systems for space station applications were identified. The low-thrust engine (5 to 100 lb sub f) requires significant departure from current applications of oxygen/hydrogen propulsion technology. Selection of the thrust chamber material and cooling method needed or long life poses a major challenge. The use of a chamber material requiring a minimum amount of cooling or the incorporation of regenerative cooling were the only choices available with the potential of achieving very high performance. The design selection for the injector/igniter, the design and fabrication of a regeneratively cooled copper chamber, and the design of a high-temperature rhenium chamber were documented and the performance and heat transfer results obtained from the test program conducted at JPL using the above engine components presented. Approximately 115 engine firings were conducted in the JPL vacuum test facility, using 100:1 expansion ratio nozzles. Engine mixture ratio and fuel-film cooling percentages were parametrically investigated for each test configuration.

Transonic Performance Characteristics of Several Jet Noise Suppressors

NASA Technical Reports Server (NTRS)

Schmeer, James W.; Salters, Leland B., Jr.; Cassetti, Marlowe D.

1960-01-01

An investigation of the transonic performance characteristics of several noise-suppressor configurations has been conducted in the Langley 16-foot transonic tunnel. The models were tested statically and over a Mach number range from 0.70 to 1.05 at an angle of attack of 0 deg. The primary jet total-pressure ratio was varied from 1.0 (jet off) to about 4.5. The effect of secondary air flow on the performance of two of the configurations was investigated. A hydrogen peroxide turbojet-engine simulator was used to supply the hot-jet exhaust. An 8-lobe afterbody with centerbody, short shroud, and secondary air had the highest thrust-minus-drag coefficients of the six noise-suppressor configurations tested. The 12-tube and 12-lobe afterbodies had the lowest internal losses. The presence of an ejector shroud partially shields the external pressure distribution of the 8-lobe after-body from the influence of the primary jet. A ring-airfoil shroud increased the static thrust of the annular nozzle but generally decreased the thrust minus drag at transonic Mach numbers.

Preliminary Sizing and Performance Evaluation of Supersonic Cruise Aircraft

NASA Technical Reports Server (NTRS)

Fetterman, D. E., Jr.

1976-01-01

The basic processes of a method that performs sizing operations on a baseline aircraft and determines their subsequent effects on aerodynamics, propulsion, weights, and mission performance are described. The input requirements of the associated computer program are defined and its output listings explained. Results obtained by applying the method to an advanced supersonic technology concept are discussed. These results include the effects of wing loading, thrust-to-weight ratio, and technology improvements on range performance, and possible gains in both range and payload capability that become available through growth versions of the baseline aircraft. The results from an in depth contractual study that confirm the range gain predicted for a particular wing loading, thrust-to-weight ratio combination are also included.

Initial Thrust Measurements of Marshall's Ion-ioN Thruster

NASA Technical Reports Server (NTRS)

Caruso, Natalie R. S.; Scogin, Tyler; Liu, Thomas M.; Walker, Mitchell L. R.; Polzin, Kurt A.; Dankanich, John W.

2015-01-01

Electronegative ion thrusters are a variation of traditional gridded ion thruster technology differentiated by the production and acceleration of both positive and negative ions. Benefits of electronegative ion thrusters include the elimination of lifetime-limiting cathodes from the thruster architecture and the ability to generate appreciable thrust from both charge species. While much progress has been made in the development of electronegative ion thruster technology, direct thrust measurements are required to unambiguously demonstrate the efficacy of the concept and support continued development. In the present work, direct thrust measurements of the thrust produced by the MINT (Marshall's Ion-ioN Thruster) are performed using an inverted-pendulum thrust stand in the High-Power Electric Propulsion Laboratory's Vacuum Test Facility-1 at the Georgia Institute of Technology with operating pressures ranging from 4.8 x 10(exp -5) and 5.7 x 10(exp -5) torr. Thrust is recorded while operating with a propellant volumetric mixture ratio of 5:1 argon to nitrogen with total volumetric flow rates of 6, 12, and 24 sccm (0.17, 0.34, and 0.68 mg/s). Plasma is generated using a helical antenna at 13.56 MHz and radio frequency (RF) power levels of 150 and 350 W. The acceleration grid assembly is operated using both sinusoidal and square waveform biases of +/-350 V at frequencies of 4, 10, 25, 125, and 225 kHz. Thrust is recorded for two separate thruster configurations: with and without the magnetic filter. No thrust is discernable during thruster operation without the magnetic filter for any volumetric flow rate, RF forward Power level, or acceleration grid biasing scheme. For the full thruster configuration, with the magnetic filter installed, a brief burst of thrust of approximately 3.75 mN +/- 3 mN of error is observed at the start of grid operation for a volumetric flow rate of 24 sccm at 350 W RF power using a sinusoidal waveform grid bias at 125 kHz and +/- 350 V. Similar bursts in thrust are observed using a square waveform grid bias at 10 kHz and +/- 350 V for volumetric flow rates of 6, 10, and 12 sccm at 150, 350, and 350 W respectively. The only operating condition that exhibits repeated thrust spikes throughout thruster operation is the 24 sccm condition with a 5:1 mixture ratio at 150 W RF power using the 10 kHz square waveform acceleration grid bias. Thrust spikes for this condition measure 3 mN with an error of +/- 2.5 mN. There are no operating conditions tested that show continuous thrust production.

Directed energy deflection laboratory measurements

NASA Astrophysics Data System (ADS)

Brashears, Travis; Lubin, Phillip; Hughes, Gary B.; Meinhold, Peter; Suen, Jonathan; Batliner, Payton; Motta, Caio; Griswold, Janelle; Kangas, Miikka; Johansson, Isbella; Alnawakhtha, Yusuf; Prater, Kenyon; Lang, Alex; Madajian, Jonathan

2015-09-01

We report on laboratory studies of the effectiveness of directed energy planetary defense as a part of the DESTAR (Directed Energy System for Targeting of Asteroids and exploRation) program. DE-STAR [1][5][6] and DE-STARLITE [2][5][6] are directed energy "stand-off" and "stand-on" programs, respectively. These systems consist of a modular array of kilowatt-class lasers powered by photovoltaics, and are capable of heating a spot on the surface of an asteroid to the point of vaporization. Mass ejection, as a plume of evaporated material, creates a reactionary thrust capable of diverting the asteroid's orbit. In a series of papers, we have developed a theoretical basis and described numerical simulations for determining the thrust produced by material evaporating from the surface of an asteroid [1][2][3][4][5][6]. In the DE-STAR concept, the asteroid itself is used as the deflection "propellant". This study presents results of experiments designed to measure the thrust created by evaporation from a laser directed energy spot. We constructed a vacuum chamber to simulate space conditions, and installed a torsion balance that holds an "asteroid" sample. The sample is illuminated with a fiber array laser with flux levels up to 60 MW/m2 which allows us to simulate a mission level flux but on a small scale. We use a separate laser as well as a position sensitive centroid detector to readout the angular motion of the torsion balance and can thus determine the thrust. We compare the measured thrust to the models. Our theoretical models indicate a coupling coefficient well in excess of 100 μN/Woptical, though we assume a more conservative value of 80 μN/Woptical and then degrade this with an optical "encircled energy" efficiency of 0.75 to 60 μN/Woptical in our deflection modeling. Our measurements discussed here yield about 45 μN/Wabsorbed as a reasonable lower limit to the thrust per optical watt absorbed.

Solar-Thermal Engine Testing

NASA Technical Reports Server (NTRS)

Tucker, Stephen; Salvail, Pat; Haynes, Davy (Technical Monitor)

2001-01-01

A solar-thermal engine serves as a high-temperature solar-radiation absorber, heat exchanger, and rocket nozzle. collecting concentrated solar radiation into an absorber cavity and transferring this energy to a propellant as heat. Propellant gas can be heated to temperatures approaching 4,500 F and expanded in a rocket nozzle, creating low thrust with a high specific impulse (I(sub sp)). The Shooting Star Experiment (SSE) solar-thermal engine is made of 100 percent chemical vapor deposited (CVD) rhenium. The engine 'module' consists of an engine assembly, propellant feedline, engine support structure, thermal insulation, and instrumentation. Engine thermal performance tests consist of a series of high-temperature thermal cycles intended to characterize the propulsive performance of the engines and the thermal effectiveness of the engine support structure and insulation system. A silicone-carbide electrical resistance heater, placed inside the inner shell, substitutes for solar radiation and heats the engine. Although the preferred propellant is hydrogen, the propellant used in these tests is gaseous nitrogen. Because rhenium oxidizes at elevated temperatures, the tests are performed in a vacuum chamber. Test data will include transient and steady state temperatures on selected engine surfaces, propellant pressures and flow rates, and engine thrust levels. The engine propellant-feed system is designed to Supply GN2 to the engine at a constant inlet pressure of 60 psia, producing a near-constant thrust of 1.0 lb. Gaseous hydrogen will be used in subsequent tests. The propellant flow rate decreases with increasing propellant temperature, while maintaining constant thrust, increasing engine I(sub sp). In conjunction with analytical models of the heat exchanger, the temperature data will provide insight into the effectiveness of the insulation system, the structural support system, and the overall engine performance. These tests also provide experience on operational aspects of the engine and associated subsystems, and will include independent variation of both steady slate heat-exchanger temperature prior to thrust operation and nitrogen inlet pressure (flow rate) during thrust operation. Although the Shooting Star engines were designed as thermal-storage engines to accommodate mission parameters, they are fully capable of operating as scalable, direct-gain engines. Tests are conducted in both operational modes. Engine thrust and propellant flow rate will be measured and thereby I(sub sp). The objective of these tests is to investigate the effectiveness of the solar engine as a heat exchanger and a rocket. Of particular interest is the effectiveness of the support structure as a thermal insulator, the integrity of both the insulation system and the insulation containment system, the overall temperature distribution throughout the engine module, and the thermal power required to sustain steady state fluid temperatures at various flow rates.

Limited variance control in statistical low thrust guidance analysis. [stochastic algorithm for SEP comet Encke flyby mission

NASA Technical Reports Server (NTRS)

Jacobson, R. A.

1975-01-01

Difficulties arise in guiding a solar electric propulsion spacecraft due to nongravitational accelerations caused by random fluctuations in the magnitude and direction of the thrust vector. These difficulties may be handled by using a low thrust guidance law based on the linear-quadratic-Gaussian problem of stochastic control theory with a minimum terminal miss performance criterion. Explicit constraints are imposed on the variances of the control parameters, and an algorithm based on the Hilbert space extension of a parameter optimization method is presented for calculation of gains in the guidance law. The terminal navigation of a 1980 flyby mission to the comet Encke is used as an example.

Elimination of Intermediate-Frequency Combustion Instability in the Fastrac Engine Thrust Chamber

NASA Technical Reports Server (NTRS)

Rocker, Marvin; Nesman, Tomas E.; Turner, Jim E. (Technical Monitor)

2001-01-01

A series of tests were conducted to measure the combustion performance of the Fastrac engine thrust chamber. The thrust chamber exhibited benign, yet marginally unstable combustion. The marginally unstable combustion was characterized by chamber pressure oscillations with large amplitudes and a frequency that was too low to be identified as acoustic or high-frequency combustion instability and too high to be identified as chug or low-frequency combustion instability. The source of the buzz or intermediate-frequency combustion instability was traced to the fuel venturi whose violently noisy cavitation caused resonance in the feedline downstream. Combustion was stabilized by increasing the throat diameter of the fuel venturi such that the cavitation would occur more quietly.

Optimal Pitch Thrust-Vector Angle and Benefits for all Flight Regimes

NASA Technical Reports Server (NTRS)

Gilyard, Glenn B.; Bolonkin, Alexander

2000-01-01

The NASA Dryden Flight Research Center is exploring the optimum thrust-vector angle on aircraft. Simple aerodynamic performance models for various phases of aircraft flight are developed and optimization equations and algorithms are presented in this report. Results of optimal angles of thrust vectors and associated benefits for various flight regimes of aircraft (takeoff, climb, cruise, descent, final approach, and landing) are given. Results for a typical wide-body transport aircraft are also given. The benefits accruable for this class of aircraft are small, but the technique can be applied to other conventionally configured aircraft. The lower L/D aerodynamic characteristics of fighters generally would produce larger benefits than those produced for transport aircraft.

Thrust Production and Wake Structure of a Batoid-Inspired Oscillating Fin

NASA Astrophysics Data System (ADS)

Clark, Richard

2005-11-01

Experiments are reported on the hydrodynamic performance of a flexible fin. The fin replicates some features of the pectoral fin of a batoid fish (such as a ray or skate) in that it is actuated in a traveling wave motion, with the amplitude of the motion increasing linearly along the span from root to tip. Thrust is found to increase with non-dimensional frequency, and an optimal oscillatory gait is identified. Power consumption measurements lead to the computation of Froude efficiency, and an optimal efficiency condition is evaluated. Wake visualizations are presented, and a vortex model of the wake near zero net thrust is suggested. Strouhal number effects on the wake topology are also illustrated.

Thrust production and wake structure of a batoid-inspired oscillating fin

NASA Astrophysics Data System (ADS)

Clark, R. P.; Smits, A. J.

2006-09-01

Experiments are reported on the hydrodynamic performance of a flexible fin. The fin replicates some features of the pectoral fin of a batoid fish (such as a ray or skate) in that it is actuated in a travelling wave motion, with the amplitude of the motion increasing linearly along the span from root to tip. Thrust is found to increase with non-dimensional frequency, and an optimal oscillatory gait is identified. Power consumption measurements lead to the computation of propulsive efficiency, and an optimal efficiency condition is evaluated. Wake visualizations are presented, and a vortex model of the wake near zero net thrust is suggested. Strouhal number effects on the wake topology are also illustrated.

Low-thrust Isp sensitivity study

NASA Technical Reports Server (NTRS)

Schoenman, L.

1982-01-01

A comparison of the cooling requirements and attainable specific impulse performance of engines in the 445 to 4448N thrust class utilizing LOX/RP-1, LOX/Hydrogen and LOX/Methane propellants is presented. The unique design requirements for the regenerative cooling of low-thrust engines operating at high pressures (up to 6894 kPa) were explored analytically by comparing single cooling with the fuel and the oxidizer, and dual cooling with both the fuel and the oxidizer. The effects of coolant channel geometry, chamber length, and contraction ratio on the ability to provide proper cooling were evaluated, as was the resulting specific impulse. The results show that larger contraction ratios and smaller channels are highly desirable for certain propellant combinations.

Design of a Low-Energy FARAD Thruster

NASA Technical Reports Server (NTRS)

Polzin, K. A.; Rose, M. F.; Miller, R.; Best, S.; Owens, T.; Dankanich, J.

2007-01-01

The design of an electrodeless thruster that relies on a pulsed, rf-assisted discharge and electromagnetic acceleration using an inductive coil is presented. The thruster design is optimized using known performance,scaling parameters, and experimentally-determined design rules, with design targets for discharge energy, plasma exhaust velocity; and thrust efficiency of 100 J/pulse, 25 km/s, and 50%, respectively. Propellant is injected using a high-speed gas valve and preionized by a pulsed-RF signal supplied by a vector inversion generator, allowing for current sheet formation at lower discharge voltages and energies relative to pulsed inductive accelerators that do not employ preionization. The acceleration coil is designed to possess an inductance of at least 700 nH while the target stray (non-coil) inductance in the circuit is 70 nH. A Bernardes and Merryman pulsed power train or a pulse compression power train provide current to the acceleration coil and solid-state components are used to switch both powertrains.

Performance modelling of plasma microthruster nozzles in vacuum

NASA Astrophysics Data System (ADS)

Ho, Teck Seng; Charles, Christine; Boswell, Rod

2018-05-01

Computational fluid dynamics and plasma simulations of three geometrical variations of the Pocket Rocket radiofrequency plasma electrothermal microthruster are conducted, comparing pulsed plasma to steady state cold gas operation. While numerical limitations prevent plasma modelling in a vacuum environment, results may be obtained by extrapolating from plasma simulations performed in a pressurised environment, using the performance delta from cold gas simulations performed in both environments. Slip regime boundary layer effects are significant at these operating conditions. The present investigation targets a power budget of ˜10 W for applications on CubeSats. During plasma operation, the thrust force increases by ˜30% with a power efficiency of ˜30 μNW-1. These performance metrics represent instantaneous or pulsed operation and will increase over time as the discharge chamber attains thermal equilibrium with the heated propellant. Additionally, the sculpted nozzle geometry achieves plasma confinement facilitated by the formation of a plasma sheath at the nozzle throat, and fast recombination ensures a neutral exhaust plume that avoids the contamination of solar panels and interference with externally mounted instruments.

Developments in REDES: The rocket engine design expert system

NASA Technical Reports Server (NTRS)

Davidian, Kenneth O.

1990-01-01

The Rocket Engine Design Expert System (REDES) is being developed at the NASA-Lewis to collect, automate, and perpetuate the existing expertise of performing a comprehensive rocket engine analysis and design. Currently, REDES uses the rigorous JANNAF methodology to analyze the performance of the thrust chamber and perform computational studies of liquid rocket engine problems. The following computer codes were included in REDES: a gas properties program named GASP, a nozzle design program named RAO, a regenerative cooling channel performance evaluation code named RTE, and the JANNAF standard liquid rocket engine performance prediction code TDK (including performance evaluation modules ODE, ODK, TDE, TDK, and BLM). Computational analyses are being conducted by REDES to provide solutions to liquid rocket engine thrust chamber problems. REDES is built in the Knowledge Engineering Environment (KEE) expert system shell and runs on a Sun 4/110 computer.

Developments in REDES: The Rocket Engine Design Expert System

NASA Technical Reports Server (NTRS)

Davidian, Kenneth O.

1990-01-01

The Rocket Engine Design Expert System (REDES) was developed at NASA-Lewis to collect, automate, and perpetuate the existing expertise of performing a comprehensive rocket engine analysis and design. Currently, REDES uses the rigorous JANNAF methodology to analyze the performance of the thrust chamber and perform computational studies of liquid rocket engine problems. The following computer codes were included in REDES: a gas properties program named GASP; a nozzle design program named RAO; a regenerative cooling channel performance evaluation code named RTE; and the JANNAF standard liquid rocket engine performance prediction code TDK (including performance evaluation modules ODE, ODK, TDE, TDK, and BLM). Computational analyses are being conducted by REDES to provide solutions to liquid rocket engine thrust chamber problems. REDES was built in the Knowledge Engineering Environment (KEE) expert system shell and runs on a Sun 4/110 computer.

Frozen Chemistry Effects on Nozzle Performance Simulations

NASA Technical Reports Server (NTRS)

Yoder, Dennis A.; Georgiadis, Nicholas J.; O'Gara, Michael R.

2009-01-01

Simulations of exhaust nozzle flows are typically conducted assuming the gas is calorically perfect, and typically modeled as air. However the gas inside a real nozzle is generally composed of combustion products whose thermodynamic properties may differ. In this study, the effect of gas model assumption on exhaust nozzle simulations is examined. The three methods considered model the nozzle exhaust gas as calorically perfect air, a calorically perfect exhaust gas mixture, and a frozen exhaust gas mixture. In the latter case the individual non-reacting species are tracked and modeled as a gas which is only thermally perfect. Performance parameters such as mass flow rate, gross thrust, and thrust coefficient are compared as are mean flow and turbulence profiles in the jet plume region. Nozzles which operate at low temperatures or have low subsonic exit Mach numbers experience relatively minor temperature variations inside the nozzle, and may be modeled as a calorically perfect gas. In those which operate at the opposite extreme conditions, variations in the thermodynamic properties can lead to different expansion behavior within the nozzle. Modeling these cases as a perfect exhaust gas flow rather than air captures much of the flow features of the frozen chemistry simulations. Use of the exhaust gas reduces the nozzle mass flow rate, but has little effect on the gross thrust. When reporting nozzle thrust coefficient results, however, it is important to use the appropriate gas model assumptions to compute the ideal exit velocity. Otherwise the values obtained may be an overly optimistic estimate of nozzle performance.

Addition of Passive Dynamics to a Flapping Airfoil to Improve Performance

NASA Astrophysics Data System (ADS)

Asselin, Daniel; Young, Jay; Williamson, C. H. K.

2017-11-01

Animals which fly or swim typically employ flapping motions of their wings and fins in order to produce thrust and to maneuver. Small, unmanned vehicles might also exploit such motions and are of considerable interest for the purposes of surveillance, environmental monitoring, and search and rescue. Flapping refers to a combination of pitch and heave and has been shown to provide good thrust and efficiency (Read, et al. 2003) when both axes are independently controlled (an Active-Active system). In this study, we examine the performance of an airfoil actuated only in the heave direction but allowed to pitch passively under the control of a torsion spring (an Active-Passive system). The presence of the spring is simulated in software using a force-feedback control system called Cyber-Physical Fluid Dynamics, or CPFD (Mackowski & Williamson 2011, 2015, 2016). Adding passive pitch to active heave provides significantly improved thrust and efficiency compared with heaving alone, especially when the torsion spring stiffness is selected so that the system operates near resonance (in an Active-Passive system). In many cases, values of thrust and efficiency are comparable to or better than those obtained with two actively controlled degrees of freedom. By using carefully-designed passive dynamics in the pitch direction, we can eliminate one of the two actuators, saving cost, complexity, and weight, while maintaining performance. This work was supported by the Air Force Office of Scientific Research Grant No. FA9550-15-1-0243, monitored by Dr. Douglas Smith.

Ignition and Performance Tests of Rocket-Based Combined Cycle Propulsion System

NASA Technical Reports Server (NTRS)

Anderson, William E.

2005-01-01

The ground testing of a Rocket Based Combined Cycle engine implementing the Simultaneous Mixing and Combustion scheme was performed at the direct-connect facility of Purdue University's High Pressure Laboratory. The fuel-rich exhaust of a JP-8/H2O2 thruster was mixed with compressed, metered air in a constant area, axisymmetric duct. The thruster was similar in design and function to that which will be used in the flight test series of Dryden's Ducted-Rocket Experiment. The determination of duct ignition limits was made based on the variation of secondary air flow rates and primary thruster equivalence ratios. Thrust augmentation and improvements in specific impulse were studied along with the pressure and temperature profiles of the duct to study mixing lengths and thermal choking. The occurrence of ignition was favored by lower rocket equivalence ratios. However, among ignition cases, better thrust and specific impulse performance were seen with higher equivalence ratios owing to the increased fuel available for combustion. Thrust and specific impulse improvements by factors of 1.2 to 1.7 were seen. The static pressure and temperature profiles allowed regions of mixing and heat addition to be identified. The mixing lengths were found to be shorter at lower rocket equivalence ratios. Total pressure measurements allowed plume-based calculation of thrust, which agreed with load-cell measured values to within 6.5-8.0%. The corresponding Mach Number profile indicated the flow was not thermally choked for the highest duct static pressure case.

Real-time in-flight engine performance and health monitoring techniques for flight research application

NASA Technical Reports Server (NTRS)

Ray, Ronald J.; Hicks, John W.; Wichman, Keith D.

1991-01-01

Procedures for real time evaluation of the inflight health and performance of gas turbine engines and related systems were developed to enhance flight test safety and productivity. These techniques include the monitoring of the engine, the engine control system, thrust vectoring control system health, and the detection of engine stalls. Real time performance techniques were developed for the determination and display of inflight thrust and for aeroperformance drag polars. These new methods were successfully shown on various research aircraft at NASA-Dryden. The capability of NASA's Western Aeronautical Test Range and the advanced data acquisition systems were key factors for implementation and real time display of these methods.

Internal performance predictions for Langley scramjet engine module

NASA Technical Reports Server (NTRS)

Pinckney, S. Z.

1978-01-01

A one dimensional theoretical method for the prediction of the internal performance of a scramjet engine is presented. The effects of changes in vehicle forebody flow parameters and characteristics on predicted thrust for the scramjet engine were evaluated using this method, and results are presented. A theoretical evaluation of the effects of changes in the scramjet engine's internal parameters is also presented. Theoretical internal performance predictions, in terms thrust coefficient and specific impulse, are provided for the scramjet engine for free stream Mach numbers of 5, 6, and 7 free stream dynamic pressure of 23,940 N/sq m forebody surface angles of 4.6 deg to 14.6 deg, and fuel equivalence ratio of 1.0.

Electric sail elliptic displaced orbits with advanced thrust model

NASA Astrophysics Data System (ADS)

Niccolai, Lorenzo; Quarta, Alessandro A.; Mengali, Giovanni

2017-09-01

This paper analyzes the performance of an Electric Solar Wind Sail for generating and maintaining an elliptic, heliocentric, displaced non-Keplerian orbit. In this sense, this paper extends and completes recent studies regarding the performances of an Electric Solar Wind Sail that covers a circular, heliocentric, displaced orbit of given characteristics. The paper presents the general equations that describe the elliptic orbit maintenance in terms of both spacecraft attitude and performance requirements, when a refined thrust model (recently proposed for the preliminary mission design) is taken into account. In particular, the paper also discusses some practical applications on particular mission scenarios in which an analytic solution of the governing equations has been found.

Chasing a Comet with a Solar Sail

NASA Technical Reports Server (NTRS)

Stough, Robert W.; Heaton, Andrew F.; Whorton, Mark S.

2008-01-01

Solar sail propulsion systems enable a wide range of missions that require constant thrust or high delta-V over long mission times. One particularly challenging mission type is a comet rendezvous mission. This paper presents optimal low-thrust trajectory designs for a range of sailcraft performance metrics and mission transit times that enables a comet rendezvous mission. These optimal trajectory results provide a trade space which can be parameterized in terms of mission duration and sailcraft performance parameters such that a design space for a small satellite comet chaser mission is identified. These results show that a feasible space exists for a small satellite to perform a comet chaser mission in a reasonable mission time.

Static internal performance of a two-dimensional convergent-divergent nozzle with thrust vectoring

NASA Technical Reports Server (NTRS)

Bare, E. Ann; Reubush, David E.

1987-01-01

A parametric investigation of the static internal performance of multifunction two-dimensional convergent-divergent nozzles has been made in the static test facility of the Langley 16-Foot Transonic Tunnel. All nozzles had a constant throat area and aspect ratio. The effects of upper and lower flap angles, divergent flap length, throat approach angle, sidewall containment, and throat geometry were determined. All nozzles were tested at a thrust vector angle that varied from 5.60 tp 23.00 deg. The nozzle pressure ratio was varied up to 10 for all configurations.

Regeneratively cooled rocket engine for space storable propellants

NASA Technical Reports Server (NTRS)

Wagner, W. R.

1973-01-01

Analysis, design, fabrication, and test efforts were performed for the existing OF2/B2H6 regeneratively cooled lK (4448 N) thrust chamber to illustrate simultaneous B2H6 fuel and OF2 oxidizer cooling and to provide results for a gaseous propellant condition injected into the combustion chamber. Data derived from performance, thermal and flow measurements confirmed predictions derived from previous test work and from concurrent analytical study. Development data derived from the experimental study were indicated to be sufficient to develop a preflight thrust chamber demonstrator prototype for future space mission objectives.

Performance mapping of a 30 cm engineering model thruster

NASA Technical Reports Server (NTRS)

Poeschel, R. L.; Vahrenkamp, R. P.

1975-01-01

A 30 cm thruster representative of the engineering model design has been tested over a wide range of operating parameters to document performance characteristics such as electrical and propellant efficiencies, double ion and beam divergence thrust loss, component equilibrium temperatures, operational stability, etc. Data obtained show that optimum power throttling, in terms of maximum thruster efficiency, is not highly sensitive to parameter selection. Consequently, considerations of stability, discharge chamber erosion, thrust losses, etc. can be made the determining factors for parameter selection in power throttling operations. Options in parameter selection based on these considerations are discussed.

Performance of an electrically raised, synchronous satellite when subjected to radiation degradation effects

NASA Technical Reports Server (NTRS)

Cake, J. E.; Regetz, J. D., Jr.

1971-01-01

The use of solar electric propulsion to raise a high-power communication satellite from a low altitude, inclined circular orbit of the geosynchronous orbit is evaluated. Since the satellite ascends through the high intensity radiation belts, the power available from the solar array and therefore to the ion thrusters degrades. The performance of the solar electric stage in combination with the thrust augmented Thor/Delta launch vehicle is evaluated for two thrust steering programs. The transfer times and solar array requirements are presented for total geosynchronous payloads from 450 to 1100 kg.

The pasty propellant rocket engine development

NASA Astrophysics Data System (ADS)

Kukushkin, V. I.; Ivanchenko, A. N.

1993-06-01

The paper describes a newly developed pasty propellant rocket engine (PPRE) and the combustion process and presents results of performance tests. It is shown that, compared with liquid propellant rocket engines, the PPREs can regulate the thrust level within a wider range, are safer ecologically, and have better weight characteristics. Compared with solid propellant rocket engines, the PPREs may be produced with lower costs and more safely, are able to regulate thrust performance within a wider range, and are able to offer a greater scope for the variation of the formulation components and propellant characteristics. Diagrams of the PPRE are included.

Advanced engine study program

NASA Technical Reports Server (NTRS)

Masters, A. I.; Galler, D. E.; Denman, T. F.; Shied, R. A.; Black, J. R.; Fierstein, A. R.; Clark, G. L.; Branstrom, B. R.

1993-01-01

A design and analysis study was conducted to provide advanced engine descriptions and parametric data for space transfer vehicles. The study was based on an advanced oxygen/hydrogen engine in the 7,500 to 50,000 lbf thrust range. Emphasis was placed on defining requirements for high-performance engines capable of achieving reliable and versatile operation in a space environment. Four variations on the expander cycle were compared, and the advantages and disadvantages of each were assessed. Parametric weight, envelope, and performance data were generated over a range of 7,500 to 50,000 lb thrust and a wide range of chamber pressure and nozzle expansion ratio.

Rotating and positive-displacement pumps for low-thrust rocket engines. Volume 2: Fabrication and testing

NASA Technical Reports Server (NTRS)

Csomor, A.

1974-01-01

Rotating and positive displacement pumps of various types were studied for pumping liquid fluorine for low thrust high performance rocket engines. Included in the analysis were: centrifugal, pitot, Barske, Tesla, drag, gear, vane, axial piston, radial piston, diaphragm and helirotor pump concepts. The centrifugal and gear pumps were carried through detail design and fabrication. After preliminary testing in Freon 12, the centrifugal pump was selected for further testing and development. It was tested in Freon 12 to obtain the hydrodynamic performance. Tests were also conducted in liquid fluorine to demonstrate chemical compatibility.

Design and performance of compliant thrust bearing with spiral-groove membranes on resilient supports

NASA Technical Reports Server (NTRS)

Licht, L.; Anderson, W. J.; Doroff, S. W.

1980-01-01

Novel thrust bearings with spiral-groove flexible membranes mounted on resilient supports were designed and their performance demonstrated. Advantages of surface compliance were combined with the superior load-capacity of the spiral-groove geometry. Loads of 127-150N were supported on an area 42 sq cm, at speeds of 43,000-45,000 rpm and mean clearances of 15-20 microns. Support-worthiness was proved when tested in conjunction with foil journal-bearings and a 19N rotor, excited in a pitching mode by a total unbalance of 43 micron-N.

Altitude Performance Characteristics of Tail-pipe Burner with Variable-area Exhaust Nozzle

NASA Technical Reports Server (NTRS)

Jansen, Emmert T; Thorman, H Carl

1950-01-01

An investigation was conducted in the NACA Lewis altitude wind tunnel to determine effect of altitude and flight Mach number on performance of tail-pipe burner equipped with variable-area exhaust nozzle and installed on full-scale turbojet engine. At a given flight Mach number, with constant exhaust-gas and turbine-outlet temperatures, increasing altitude lowered the tail-pipe combustion efficiency and raised the specific fuel consumption while the augmented thrust ratio remained approximately constant. At a given altitude, increasing flight Mach number raised the combustion efficiency and augmented thrust ratio and lowered the specific fuel consumption.

Simulated afterburner performance with hydrogen peroxide injection for thrust augmentation

NASA Technical Reports Server (NTRS)

Metzler, Allen J; Grobman, Jack S

1956-01-01

Combustion performance of three afterburner configurations was evaluated at simulated altitude flight conditions with liquid augmentation to the primary combustor. Afterburner combustion efficiency and stability were better with injection of high-strength hydrogen peroxide than with no injection or with water injection. Improvements were observed in afterburner configurations with and without flameholders and in a short-length afterburner. At a peroxide-air ratio of 0.3, combustion was stable and 85 to 90 percent efficient in all configurations tested. Calculated augmented net-thrust ratios for peroxide injection with afterburning were approximately 60 percent greater than those for water injection.

Jet noise suppressor nozzle development for augmentor wing jet STOL research aircraft (C-8A Buffalo)

NASA Technical Reports Server (NTRS)

Harkonen, D. L.; Marks, C. C.; Okeefe, J. V.

1974-01-01

Noise and performance test results are presented for a full-scale advanced design rectangular array lobe jet suppressor nozzle (plain wall and corrugated). Flight design and installation considerations are also discussed. Noise data are presented in terms of peak PNLT (perceived noise level, tone corrected) suppression relative to the existing airplane and one-third octave-band spectra. Nozzle performance is presented in terms of velocity coefficient. Estimates of the hot thrust available during emergency (engine out) with the suppressor nozzle installed are compared with the current thrust levels produced by the round convergent nozzles.

Static-thrust Investigation of Full-scale PV-2 Helicopter Rotors Having NACA 0012.6 and 23012.6 Airfoil Sections

NASA Technical Reports Server (NTRS)

Lipson, Stanley

1946-01-01

An investigation was conducted to compare the performance of two 25-ft-diam rotors which had identical dimensions and were similar in construction but different in blade airfoil-sections. Tests were conducted at indicated blade pitch angles from 3 degrees to 11.5 degrees and rotor speeds of 200, 290, and 371 rpm. The 23012.6 rotor required 2 percent less power to hover than the 0012.6. At thrust coefficients above design, the performance of the 23012.6 became better than the 0012.6 rotor.

Hybrid propulsion technology program: Phase 1. Volume 3: Thiokol Corporation Space Operations

NASA Technical Reports Server (NTRS)

Schuler, A. L.; Wiley, D. R.

1989-01-01

Three candidate hybrid propulsion (HP) concepts were identified, optimized, evaluated, and refined through an iterative process that continually forced improvement to the systems with respect to safety, reliability, cost, and performance criteria. A full scale booster meeting Advanced Solid Rocket Motor (ASRM) thrust-time constraints and a booster application for 1/4 ASRM thrust were evaluated. Trade studies and analyses were performed for each of the motor elements related to SRM technology. Based on trade study results, the optimum HP concept for both full and quarter sized systems was defined. The three candidate hybrid concepts evaluated are illustrated.

The immediate effects of manual stretching and cervicothoracic junction manipulation on cervical range of motion and upper trapezius pressure pain thresholds.

PubMed

Hanney, William J; Puentedura, Emilio J; Kolber, Morey J; Liu, Xinliang; Pabian, Patrick S; Cheatham, Scott W

2017-09-22

Myofascial pain is a common impairment treated with various manual interventions including spinal thrust manipulation and stretching; however, the comparative efficacy of each intervention is uncertain. Therefore, the purpose of this investigation was to evaluate thrust manipulation targeting the cervicothoracic junction compared to a manual stretch of the upper trapezius muscle on cervical range of motion and upper trapezius pressure pain thresholds (PPTs). Healthy participants with no significant history of neck pain were randomized into a thrust manipulation group, a stretching group, or a control group. Within group differences were evaluated via a dependent t-test, and group by time interactions were evaluated by a two-way repeated measures ANOVA. One hundred and two participants were recruited to participate. Baseline demographics revealed no significant differences between groups. Significant group by time interactions were found for changes in PPTs for both the right and left upper trapezius. Also, significant differences were found for changes in cervical extension, as well as right and left cervical side bending favoring the treatment groups. This study demonstrates the potential independent effectiveness of spinal thrust manipulation or stretching for reducing PPTs at the upper trapezius. Future research should further evaluate the limitation of PPTs as a measure of muscle sensitivity as well as factors that may contribute to variability in the measurements among individuals seeking care.

Pursuit/evasion in orbit

NASA Technical Reports Server (NTRS)

Kelley, H. J.; Cliff, E. M.; Lutze, F. H.

1981-01-01

Maneuvers available to a spacecraft having sufficient propellant to escape an antisatellite satellite (ASAT) attack are examined. The ASAT and the evading spacecraft are regarded as being in circular orbits, and equations of motion are developed for the ASAT to commence a two-impulse maneuver sequence. The ASAT employs thrust impulses which yield a minimum-time-to-rendezvous, considering available fuel. Optimal evasion is shown to involve only in-plane maneuvers, and begins as soon as the ASAT launch information is gathered and thrust activation can be initiated. A closest approach, along with a maximum evasion by the target spacecraft, is calculated to be 14,400 ft. Further research to account for ASATs in parking orbit and for generalization of a continuous control-modeled differential game is indicated.

Guidance strategies and analysis for low thrust navigation

NASA Technical Reports Server (NTRS)

Jacobson, R. A.

1973-01-01

A low-thrust guidance algorithm suitable for operational use was formulated. A constrained linear feedback control law was obtained using a minimum terminal miss criterion and restricting control corrections to constant changes for specified time periods. Both fixed- and variable-time-of-arrival guidance were considered. The performance of the guidance law was evaluated by applying it to the approach phase of the 1980 rendezvous mission with the comet Encke.

Low-Thrust Bipropellant Engine Technology.

DTIC Science & Technology

1980-08-01

Non-Destructive Testing OD Outside Diameter xv tr. GLOSSARY (cont.J ODE One Dimensional Equilibrium ODK One Dimensional Kinetics Pc Thrust Chamber...performance (280 sec steady- state, 220 sec pulsing) have not yet been collectively achieved, but should be obtainable with further development activities...even at nozzle area ratios up to 400:1. The influence of nozzle kinetics (i.e., equilibrium versus frozen flow and ODK ) are noted to be a much more

Space transportation booster engine thrust chamber technology, large scale injector

NASA Technical Reports Server (NTRS)

Schneider, J. A.

1993-01-01

The objective of the Large Scale Injector (LSI) program was to deliver a 21 inch diameter, 600,000 lbf thrust class injector to NASA/MSFC for hot fire testing. The hot fire test program would demonstrate the feasibility and integrity of the full scale injector, including combustion stability, chamber wall compatibility (thermal management), and injector performance. The 21 inch diameter injector was delivered in September of 1991.

Nuclear Thermal Propulsion: Past, Present, and a Look Ahead

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.

2014-01-01

NTR: High thrust high specific impulse (2 x LOXLH2 chemical) engine uses high power density fission reactor with enriched uranium fuel as thermal power source. Reactor heat is removed using H2 propellant which is then exhausted to produce thrust. Conventional chemical engine LH2 tanks, turbo pumps, regenerative nozzles and radiation-cooled shirt extensions used -- NTR is next evolutionary step in high performance liquid rocket engines.

Noncatalytic hydrazine thruster development - 0.050 to 5.0 pounds thrust

NASA Technical Reports Server (NTRS)

Murch, C. K.; Sackheim, R. L.; Kuenzly, J. D.; Callens, R. A.

1976-01-01

Noncatalytic (thermal-decompositon) hydrazine thrusters can operate in both the pulsing and steady-state modes to meet the propulsive requirements of long-life spacecraft. The thermal decomposition mode yields higher specific impulse than is characteristic of catalytic thrusters at similar thrust levels. This performance gain is the result of higher temperature operation and a lower fraction of ammonia dissociation. Some life limiting factors of catalytic thrusters are eliminated.

NATO Standardization and Licensing Policy-Exploratory Phase. Volume 3. Supplement

DTIC Science & Technology

1976-11-30

Aerospatiale will perform at least half the production of this aircraft.* The concept involves a relatively light- weight advanced interceptor with...a ratio of total thrust to takeoff gross weight of about unity. The technology of the aircraft calls for the use of carbon fiber in selected...presenting a system with operating temperatures, pressure ratios, and thrust-to- weight ratios roughly comparable to U.S. technology. The

Comparison of continuous and discontinuous collisional bumpers: Dimensionally scaled impact experiments into single wire meshes

NASA Technical Reports Server (NTRS)

Hoerz, Friedrich; Cintala, Mark; See, Thomas; Bernhard, Ronald; Cardenas, Frank; Davidson, William; Haynes, Jerry

1992-01-01

An experimental inquiry into the utility of discontinuous bumpers was conducted to investigate the collisional outcomes of impacts into single grid-like targets and to compare the results with more traditional bumper designs that employ continuous sheet stock. We performed some 35 experiments using 6.3 and 3.2 mm diameter spherical soda-lime glass projectiles at low velocities (less than 2.5 km/s) and 13 at velocities between 5 and 6 km/s, using 3.2 mm spheres only. The thrust of the experiments related to the characterization of collisional fragments as a function of target thickness or areal shield mass of both bumper designs. The primary product of these experiments was witness plates that record the resulting population of collisional fragments. Substantial interpretive and predictive insights into bumper performance were obtained. All qualitative observations (on the witness plates) and detailed measurements of displaced masses seem simply and consistently related only to bumper mass available for interaction with the impactor. This renders the grid bumper into the superior shield design. These findings present evidence that discontinuous bumpers are a viable concept for collisional shields, possibly superior to continuous geometries.

Test Stand at the Rocket Engine Test Facility

NASA Image and Video Library

1973-02-21

The thrust stand in the Rocket Engine Test Facility at the National Aeronautics and Space Administration (NASA) Lewis Research Center in Cleveland, Ohio. The Rocket Engine Test Facility was constructed in the mid-1950s to expand upon the smaller test cells built a decade before at the Rocket Laboratory. The $2.5-million Rocket Engine Test Facility could test larger hydrogen-fluorine and hydrogen-oxygen rocket thrust chambers with thrust levels up to 20,000 pounds. Test Stand A, seen in this photograph, was designed to fire vertically mounted rocket engines downward. The exhaust passed through an exhaust gas scrubber and muffler before being vented into the atmosphere. Lewis researchers in the early 1970s used the Rocket Engine Test Facility to perform basic research that could be utilized by designers of the Space Shuttle Main Engines. A new electronic ignition system and timer were installed at the facility for these tests. Lewis researchers demonstrated the benefits of ceramic thermal coatings for the engine’s thrust chamber and determined the optimal composite material for the coatings. They compared the thermal-coated thrust chamber to traditional unlined high-temperature thrust chambers. There were more than 17,000 different configurations tested on this stand between 1973 and 1976. The Rocket Engine Test Facility was later designated a National Historic Landmark for its role in the development of liquid hydrogen as a propellant.

Proven, long-life hydrogen/oxygen thrust chambers for space station propulsion

NASA Technical Reports Server (NTRS)

Richter, G. P.; Price, H. G.

1986-01-01

The development of the manned space station has necessitated the development of technology related to an onboard auxiliary propulsion system (APS) required to provide for various space station attitude control, orbit positioning, and docking maneuvers. A key component of this onboard APS is the thrust chamber design. To develop the required thrust chamber technology to support the Space Station Program, the NASA Lewis Research Center has sponsored development programs under contracts with Aerojet TechSystems Company and with Bell Aerospace Textron Division of Textron, Inc. During the NASA Lewis sponsored program with Aerojet TechSystems, a 25 lb sub f hydrogen/oxygen thruster has been developed and proven as a viable candidate to meet the needs of the Space Station Program. Likewise, during the development program with Bell Aerospace, a 50 lb sub f hydrogen/oxygen Thrust Chamber has been developed and has demonstrated reliable, long-life expectancy at anticipated space station operating conditions. Both these thrust chambers were based on design criteria developed in previous thruster programs and successfully verified in experimental test programs. Extensive thermal analyses and models were used to design the thrusters to achieve total impulse goals of 2 x 10 to the 6th power lb sub f-sec. Test data for each thruster will be compared to the analytical predictions for the performance and heat transfer characteristics. Also, the results of thrust chamber life verification tests will be presented.

Demand thrust pumped propulsion with automatic warm gas valving

NASA Astrophysics Data System (ADS)

Whitehead, J. C.

1992-06-01

Operation of a thrust-on-demand, monopropellant rocket propulsion system which uses lightweight low-pressure tankage, free-piston pumps, and a small high-pressure thrust chamber, is explained. The pump intake-exhaust valves use warm gas pneumatic signals to ensure that two reciprocating pumps are alternately pressurized, with overlap during switchover to permit uninterrupted propellant flow. Experiments demonstrate that the miniature pumps operate at any speed depending on downstream demand, and can deliver nearly their own mass in hydrazine per second, at 7 MPa (1000 psi). The valves, which use the alternating layers of metal and graphite to mitigate the effects of differential thermal expansion, have been warm-gas tested for thousands of cycles. For biopropellant operation, a pair of reciprocating oxidizer pumps would be slaved to the fuel pumps' pneumatic oscillator, to provide for pulsed or continuous demand-driven flow of both liquids. Mass ratios and thrust-to-weight ratios of demand-thrust pumped propulsion systems compare quite favorably to those of pressure-fed and turbo-pumped systems. Due to the relatively high densities of storable propellants, liquid mass fractions greater than 0.95 are attainable with these novel pumps, with thrust/weight ratios above 10. The high performance potential of small propulsion systems which use reciprocating pumps suggests that this technology can significantly increase the capability of many types of small spacecraft.