OPTRAN- OPTIMAL LOW THRUST ORBIT TRANSFERS

NASA Technical Reports Server (NTRS)

Breakwell, J. V.

1994-01-01

OPTRAN is a collection of programs that solve the problem of optimal low thrust orbit transfers between non-coplanar circular orbits for spacecraft with chemical propulsion systems. The programs are set up to find Hohmann-type solutions, with burns near the perigee and apogee of the transfer orbit. They will solve both fairly long burn-arc transfers and "divided-burn" transfers. Program modeling includes a spherical earth gravity model and propulsion system models for either constant thrust or constant acceleration. The solutions obtained are optimal with respect to fuel use: i.e., final mass of the spacecraft is maximized with respect to the controls. The controls are the direction of thrust and the thrust on/off times. Two basic types of programs are provided in OPTRAN. The first type is for "exact solution" which results in complete, exact tkme-histories. The exact spacecraft position, velocity, and optimal thrust direction are given throughout the maneuver, as are the optimal thrust switch points, the transfer time, and the fuel costs. Exact solution programs are provided in two versions for non-coplanar transfers and in a fast version for coplanar transfers. The second basic type is for "approximate solutions" which results in approximate information on the transfer time and fuel costs. The approximate solution is used to estimate initial conditions for the exact solution. It can be used in divided-burn transfers to find the best number of burns with respect to time. The approximate solution is useful by itself in relatively efficient, short burn-arc transfers. These programs are written in FORTRAN 77 for batch execution and have been implemented on a DEC VAX series computer with the largest program having a central memory requirement of approximately 54K of 8 bit bytes. The OPTRAN program were developed in 1983.

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.

SSME structural computer program development. Volume 3: BOPACE demonstration analysis of the SSME thrust chamber liner

NASA Technical Reports Server (NTRS)

Spring, A. H.

1973-01-01

The application of a structural computer program for analysis of a thrust chamber liner is discussed. Two objectives were accomplished as follows: (1) exercise of the full capabilities of the computer program and (2) definition of thermal and mechanical boundary conditions to reflect the emergency power level operating conditions for the SSME 47OK engine at a station just upstream of the thrust chamber throat. Creep information on the thrust chamber is presented as a reference curve of creep strain versus time for various temperatures. Contour plots of the effective plastic strain, effective stress, and effective creep strain are developed.

Saturn Apollo Program

NASA Image and Video Library

1960-01-01

H-1 engine characteristics: The H-1 engine was developed under the management of the Marshall Space Flight Center (MSFC). The cluster of eight H-1 engines was used to power the first stage of the Saturn I (S-I stage) and Saturn IB (S-IVB stage) launch vehicles, and produced 188,00 pounds of thrust, a combined thrust of 1,500,000 pounds, later uprated to 205,000 pounds of thrust and a combined total thrust of 1,650,000 pounds for the Saturn IB program.

Thrust Chamber Modeling Using Navier-Stokes Equations: Code Documentation and Listings. Volume 2

NASA Technical Reports Server (NTRS)

Daley, P. L.; Owens, S. F.

1988-01-01

A copy of the PHOENICS input files and FORTRAN code developed for the modeling of thrust chambers is given. These copies are contained in the Appendices. The listings are contained in Appendices A through E. Appendix A describes the input statements relevant to thrust chamber modeling as well as the FORTRAN code developed for the Satellite program. Appendix B describes the FORTRAN code developed for the Ground program. Appendices C through E contain copies of the Q1 (input) file, the Satellite program, and the Ground program respectively.

A users manual for a computer program which calculates time optical geocentric transfers using solar or nuclear electric and high thrust propulsion

NASA Technical Reports Server (NTRS)

Sackett, L. L.; Edelbaum, T. N.; Malchow, H. L.

1974-01-01

This manual is a guide for using a computer program which calculates time optimal trajectories for high-and low-thrust geocentric transfers. Either SEP or NEP may be assumed and a one or two impulse, fixed total delta V, initial high thrust phase may be included. Also a single impulse of specified delta V may be included after the low thrust state. The low thrust phase utilizes equinoctial orbital elements to avoid the classical singularities and Kryloff-Boguliuboff averaging to help insure more rapid computation time. The program is written in FORTRAN 4 in double precision for use on an IBM 360 computer. The manual includes a description of the problem treated, input/output information, examples of runs, and source code listings.

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.

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.

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.

Tailoff thrust and impulse imbalance between pairs of Space Shuttle solid rocket motors

NASA Technical Reports Server (NTRS)

Jacobs, E. P.; Yeager, J. M.

1975-01-01

The tailoff thrust and impulse imbalance between pairs of solid rocket motors is of particular interest for the Space Shuttle Vehicle because of the potential control problems that exist with this asymmetric configuration. Although a similar arrangement of solid rocket motors was utilized for the Titan Program, they produced less than one-half the thrust level of the Space Shuttle at web action time, and the overall vehicle was symmetric. Since the Titan Program does provide the most applicable actual test data, 23 flight pairs were analyzed to determine the actual tailoff thrust and impulse imbalance experienced. The results were scaled up using the predicted web action time thrust and tailoff time to arrive at values for the Space Shuttle. These values were then statistically treated to obtain a prediction of the maximum imbalance one could expect to experience during the Shuttle Program.

Small-scale test program to develop a more efficient swivel nozzle thrust deflector for V/STOL lift/cruise engines

NASA Technical Reports Server (NTRS)

Schlundt, D. W.

1976-01-01

The installed performance degradation of a swivel nozzle thrust deflector system obtained during increased vectoring angles of a large-scale test program was investigated and improved. Small-scale models were used to generate performance data for analyzing selected swivel nozzle configurations. A single-swivel nozzle design model with five different nozzle configurations and a twin-swivel nozzle design model, scaled to 0.15 size of the large-scale test hardware, were statically tested at low exhaust pressure ratios of 1.4, 1.3, 1.2, and 1.1 and vectored at four nozzle positions from 0 deg cruise through 90 deg vertical used for the VTOL mode.

Aerospike thrust chamber program. [cumulative damage and maintenance of structural members in hydrogen oxygen engines

NASA Technical Reports Server (NTRS)

Campbell, J., Jr.; Cobb, S. M.

1976-01-01

An existing, but damaged, 25,000-pound thrust, flightweight, oxygen/hydrogen aerospike rocket thrust chamber was disassembled and partially repaired. A description is presented of the aerospike chamber configuration and of the damage it had suffered. Techniques for aerospike thrust chamber repair were developed, and are described, covering repair procedures for lightweight tubular nozzles, titanium thrust structures, and copper channel combustors. Effort was terminated prior to completion of the repairs and conduct of a planned hot fire test program when it was found that the copper alloy walls of many of the thrust chamber's 24 combustors had been degraded in strength and ductility during the initial fabrication of the thrust chamber. The degradation is discussed and traced to a reaction between oxygen and/or oxides diffused into the copper alloy during fabrication processes and the hydrogen utilized as a brazing furnace atmosphere during the initial assembly operation on many of the combustors. The effects of the H2/O2 reaction within the copper alloy are described.

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.

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"

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.

Low-thrust solar electric propulsion navigation simulation program

NASA Technical Reports Server (NTRS)

Hagar, H. J.; Eller, T. J.

1973-01-01

An interplanetary low-thrust, solar electric propulsion mission simulation program suitable for navigation studies is presented. The mathematical models for trajectory simulation, error compensation, and tracking motion are described. The languages, input-output procedures, and subroutines are included.

Engineering Research and Development and Technology thrust area report FY92

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

Langland, R.T.; Minichino, C.

1993-03-01

The mission of the Engineering Research, Development, and Technology Program at Lawrence Livermore National Laboratory (LLNL) is to develop the technical staff and the technology needed to support current and future LLNL programs. To accomplish this mission, the Engineering Research, Development, and Technology Program has two important goals: (1) to identify key technologies and (2) to conduct high-quality work to enhance our capabilities in these key technologies. To help focus our efforts, we identify technology thrust areas and select technical leaders for each area. The thrust areas are integrated engineering activities and, rather than being based on individual disciplines, theymore » are staffed by personnel from Electronics Engineering, Mechanical Engineering, and other LLNL organizations, as appropriate. The thrust area leaders are expected to establish strong links to LLNL program leaders and to industry; to use outside and inside experts to review the quality and direction of the work; to use university contacts to supplement and complement their efforts; and to be certain that we are not duplicating the work of others. This annual report, organized by thrust area, describes activities conducted within the Program for the fiscal year 1992. Its intent is to provide timely summaries of objectives, theories, methods, and results. The nine thrust areas for this fiscal year are: Computational Electronics and Electromagnetics; Computational Mechanics; Diagnostics and Microelectronics; Emerging Technologies; Fabrication Technology; Materials Science and Engineering; Microwave and Pulsed Power; Nondestructive Evaluation; and Remote Sensing and Imaging, and Signal Engineering.« less

Optimal low thrust geocentric transfer. [mission analysis computer program

NASA Technical Reports Server (NTRS)

Edelbaum, T. N.; Sackett, L. L.; Malchow, H. L.

1973-01-01

A computer code which will rapidly calculate time-optimal low thrust transfers is being developed as a mission analysis tool. The final program will apply to NEP or SEP missions and will include a variety of environmental effects. The current program assumes constant acceleration. The oblateness effect and shadowing may be included. Detailed state and costate equations are given for the thrust effect, oblateness effect, and shadowing. A simple but adequate model yields analytical formulas for power degradation due to the Van Allen radiation belts for SEP missions. The program avoids the classical singularities by the use of equinoctial orbital elements. Kryloff-Bogoliuboff averaging is used to facilitate rapid calculation. Results for selected cases using the current program are given.

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.

Perceived Noise Analysis for Offset Jets Applied to Commercial Supersonic Aircraft

NASA Technical Reports Server (NTRS)

Huff, Dennis L.; Henderson, Brenda S.; Berton, Jeffrey J.; Seidel, Jonathan A.

2016-01-01

A systems analysis was performed with experimental jet noise data, engine/aircraft performance codes and aircraft noise prediction codes to assess takeoff noise levels and mission range for conceptual supersonic commercial aircraft. A parametric study was done to identify viable engine cycles that meet NASAs N+2 goals for noise and performance. Model scale data from offset jets was used as input to the aircraft noise prediction code to determine the expected sound levels for the lateral certification point where jet noise dominates over all other noise sources. The noise predictions were used to determine the optimal orientation of the offset nozzles to minimize the noise at the lateral microphone location. An alternative takeoff procedure called programmed lapse rate was evaluated for noise reduction benefits. Results show there are two types of engines that provide acceptable range performance; one is a standard mixed-flow turbofan with a single-stage fan, and the other is a three-stream variable-cycle engine with a multi-stage fan. The engine with a single-stage fan has a lower specific thrust and is 8 to 10 EPNdB quieter for takeoff. Offset nozzles reduce the noise directed toward the thicker side of the outer flow stream, but have less benefit as the core nozzle pressure ratio is reduced and the bypass-to-core area ratio increases. At the systems level for a three-engine N+2 aircraft with full throttle takeoff, there is a 1.4 EPNdB margin to Chapter 3 noise regulations predicted for the lateral certification point (assuming jet noise dominates). With a 10 reduction in thrust just after takeoff rotation, the margin increases to 5.5 EPNdB. Margins to Chapter 4 and Chapter 14 levels will depend on the cumulative split between the three certification points, but it appears that low specific thrust engines with a 10 reduction in thrust (programmed lapse rate) can come close to meeting Chapter 14 noise levels. Further noise reduction is possible with additional reduction in takeoff thrust using programmed lapse rate, but studies are needed to investigate the practical limits for safety and takeoff regulations.

Development of a real-time aeroperformance analysis technique for the X-29A advanced technology demonstrator

NASA Technical Reports Server (NTRS)

Ray, R. J.; Hicks, J. W.; Alexander, R. I.

1988-01-01

The X-29A advanced technology demonstrator has shown the practicality and advantages of the capability to compute and display, in real time, aeroperformance flight results. This capability includes the calculation of the in-flight measured drag polar, lift curve, and aircraft specific excess power. From these elements many other types of aeroperformance measurements can be computed and analyzed. The technique can be used to give an immediate postmaneuver assessment of data quality and maneuver technique, thus increasing the productivity of a flight program. A key element of this new method was the concurrent development of a real-time in-flight net thrust algorithm, based on the simplified gross thrust method. This net thrust algorithm allows for the direct calculation of total aircraft drag.

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.

Path to a Research Plan

NASA Technical Reports Server (NTRS)

Chiaramonte, Fran

2003-01-01

This viewgraph presentation discusses the status and goals for the NASA OBPR Physical Science Research Program. The following text was used to summarize the presentation. The OBPR Physical Sciences Research program has been comprehensively reviewed and endorsed by National Research Council. The value and need for the research have been re-affirmed. The research program has been prioritized and resource re-allocations have been carried out through an OBPR-wide process. An increasing emphasis on strategic, mission-oriented research is planned. The program will strive to maintain a balance between strategic and fundamental research. A feasible ISS flight research program fitting within the budgetary and ISS resource envelopes has been formulated for the near term (2003-2007). The current ISS research program will be significantly strengthened starting 2005 by using discipline dedicated research facility racks. A research re-planning effort has been initiated and will include active participation from the research community in the next few months. The research re-planning effort will poise PSR to increase ISS research utilization for a potential enhancement beyond ISS IP Core Complete. The Physical Sciences research program readily integrates the cross-disciplinary requirements of the NASA and OBPR strategic objectives. Each fundamental research thrust will develop a roadmap through technical workshops and Discipline Working Groups (DWGs). Most fundamental research thrusts will involve cross-disciplinary efforts. A Technology Roadmap will guide the Strategic Research for Exploration thrust. The Research Plan will integrate and coordinate fundamental Research Thrusts Roadmaps with the Technology Roadmap. The Technology Roadmap will be developed in coordination with other OBPR programs as well as other Enterprise (R,S,M,N). International Partners will contribute to the roadmaps and through research coordination. The research plan will be vetted with the discipline working groups, the BPRAC subcommittees, and with the BPRAC. Recommendations from NRC past and current committees will be implemented whenever appropriate.Proposed theme element content will be "missionized" around planned content and potential new projects (facilities, modules, initiatives) on approximately a five-year horizon, with the approval of PSRD management. Center/science working group teams will develop descriptions of "mission" objectives, value, and requirements. Purpose is to create a competitive environment for concept development and to stimulate community ownership/advocacy. Proposed theme elements reviewed and approved by PSRD management. Strawman roadmaps for themes developed. Program budget and technology requirements verified. Theme elements are prioritized with the input of advisory groups. Integration into program themes (questions) and required technology investments are defined by science and technology roadmaps. Review and assessment by OBPR management.

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.

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.

Relationship between Biomechanical Characteristics of Spinal Manipulation and Neural Responses in an Animal Model: Effect of Linear Control of Thrust Displacement versus Force, Thrust Amplitude, Thrust Duration, and Thrust Rate

PubMed Central

Reed, William R.; Cao, Dong-Yuan; Long, Cynthia R.; Kawchuk, Gregory N.; Pickar, Joel G.

2013-01-01

High velocity low amplitude spinal manipulation (HVLA-SM) is used frequently to treat musculoskeletal complaints. Little is known about the intervention's biomechanical characteristics that determine its clinical benefit. Using an animal preparation, we determined how neural activity from lumbar muscle spindles during a lumbar HVLA-SM is affected by the type of thrust control and by the thrust's amplitude, duration, and rate. A mechanical device was used to apply a linear increase in thrust displacement or force and to control thrust duration. Under displacement control, neural responses during the HVLA-SM increased in a fashion graded with thrust amplitude. Under force control neural responses were similar regardless of the thrust amplitude. Decreasing thrust durations at all thrust amplitudes except the smallest thrust displacement had an overall significant effect on increasing muscle spindle activity during the HVLA-SMs. Under force control, spindle responses specifically and significantly increased between thrust durations of 75 and 150 ms suggesting the presence of a threshold value. Thrust velocities greater than 20–30 mm/s and thrust rates greater than 300 N/s tended to maximize the spindle responses. This study provides a basis for considering biomechanical characteristics of an HVLA-SM that should be measured and reported in clinical efficacy studies to help define effective clinical dosages. PMID:23401713

Design and evaluation of thrust vectored nozzles using a multicomponent thrust stand

NASA Technical Reports Server (NTRS)

Carpenter, Thomas W.; Blattner, Ernest W.; Stagner, Robert E.; Contreras, Juanita; Lencioni, Dennis; Mcintosh, Greg

1990-01-01

Future aircraft with the capability of short takeoff and landing, and improved maneuverability especially in the post-stall flight regime will incorporate exhaust nozzles which can be thrust vectored. In order to conduct thrust vector research in the Mechanical Engineering Department at Cal Poly, a program was planned with two objectives; design and construct a multicomponent thrust stand for the specific purpose of measuring nozzle thrust vectors; and to provide quality low moisture air to the thrust stand for cold flow nozzle tests. The design and fabrication of the six-component thrust stand was completed. Detailed evaluation tests of the thrust stand will continue upon the receipt of one signal conditioning option (-702) for the Fluke Data Acquisition System. Preliminary design of thrust nozzles with air supply plenums were completed. The air supply was analyzed with regard to head loss. Initial flow visualization tests were conducted using dual water jets.

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.

Quiet Clean Short-Haul Experimental Engine (QCSEE) acoustic and aerodynamic tests on a scale model over-the-wing thrust reverser and forward thrust nozzle

NASA Technical Reports Server (NTRS)

Stimpert, D. L.

1978-01-01

An acoustic and aerodynamic test program was conducted on a 1/6.25 scale model of the Quiet, Clean, Short-Haul Experimental Engine (QCSEE) forward thrust over-the-wing (OTW) nozzle and OTW thrust reverser. In reverse thrust, the effect of reverser geometry was studied by parametric variations in blocker spacing, blocker height, lip angle, and lip length. Forward thrust nozzle tests determined the jet noise levels of the cruise and takeoff nozzles, the effect of opening side doors to achieve takeoff thrust, and scrubbing noise of the cruise and takeoff jet on a simulated wing surface. Velocity profiles are presented for both forward and reverse thrust nozzles. An estimate of the reverse thrust was made utilizing the measured centerline turning angle.

Bulk-Flow Analysis of Hybrid Thrust Bearings for Advanced Cryogenic Turbopumps

NASA Technical Reports Server (NTRS)

SanAndres, Luis

1998-01-01

A bulk-flow analysis and computer program for prediction of the static load performance and dynamic force coefficients of angled injection, orifice-compensated hydrostatic/hydrodynamic thrust bearings have been completed. The product of the research is an efficient computational tool for the design of high-speed thrust bearings for cryogenic fluid turbopumps. The study addresses the needs of a growing technology that requires of reliable fluid film bearings to provide the maximum operating life with optimum controllable rotordynamic characteristics at the lowest cost. The motion of a cryogenic fluid on the thin film lands of a thrust bearing is governed by a set of bulk-flow mass and momentum conservation and energy transport equations. Mass flow conservation and a simple model for momentum transport within the hydrostatic bearing recesses are also accounted for. The bulk-flow model includes flow turbulence with fluid inertia advection, Coriolis and centrifugal acceleration effects on the bearing recesses and film lands. The cryogenic fluid properties are obtained from realistic thermophysical equations of state. Turbulent bulk-flow shear parameters are based on Hirs' model with Moody's friction factor equations allowing a simple simulation for machined bearing surface roughness. A perturbation analysis leads to zeroth-order nonlinear equations governing the fluid flow for the thrust bearing operating at a static equilibrium position, and first-order linear equations describing the perturbed fluid flow for small amplitude shaft motions in the axial direction. Numerical solution to the zeroth-order flow field equations renders the bearing flow rate, thrust load, drag torque and power dissipation. Solution to the first-order equations determines the axial stiffness, damping and inertia force coefficients. The computational method uses well established algorithms and generic subprograms available from prior developments. The Fortran9O computer program hydrothrust runs on a Windows 95/NT personal computer. The program, help files and examples are licensed by Texas A&M University Technology License Office. The study of the static and dynamic performance of two hydrostatic/hydrodynamic bearings demonstrates the importance of centrifugal and advection fluid inertia effects for operation at high rotational speeds. The first example considers a conceptual hydrostatic thrust bearing for an advanced liquid hydrogen turbopump operating at 170,000 rpm. The large axial stiffness and damping coefficients of the bearing should provide accurate control and axial positioning of the turbopump and also allow for unshrouded impellers, therefore increasing the overall pump efficiency. The second bearing uses a refrigerant R134a, and its application in oil-free air conditioning compressors is of great technological importance and commercial value. The computed predictions reveal that the LH2 bearing load capacity and flow rate increase with the recess pressure (i.e. increasing orifice diameters). The bearing axial stiffness has a maximum for a recess pressure rati of approx. 0.55. while the axial damping coefficient decreases as the recess pressure ratio increases. The computer results from three flow models are compared. These models are a) inertialess, b) fluid inertia at recess edges only, and c) full fluid inertia at both recess edges and film lands. The full inertia model shows the lowest flow rates, axial load capacity and stiffness coefficient but on the other hand renders the largest damping coefficients and inertia coefficients. The most important findings are related to the reduction of the outflow through the inner radius and the appearance of subambient pressures. The performance of the refrigerant hybrid thrust bearing is evaluated at two operating speeds and pressure drops. The computed results are presented in dimensionless form to evidence consistent trends in the bearing performance characteristics. As the applied axial load increases, the bearing film thickness and flow rate decrease while the recess pressure increases. The axial stiffness coefficient shows a maximum for a certain intermediate load while the damping coefficient steadily increases. The computed results evidence the paramount of centrifugal fluid inertia at low recess pressures (i.e. low loads), and where there is actually an inflow through the bearing inner diameter, accompanied by subambient pressures just downstream of the bearing recess edge. These results are solely due to centrifugal fluid inertia and advection transport effects. Recommendations include the extension of the computer program to handle flexure pivot tilting pad hybrid bearings and the ability to calculate moment coefficients for shaft angular misalignments.

Investigation of Thrust Augmentation of a 1600-pound Thrust Centrifugal-flow-type Turbojet Engine by Injection of Refrigerants at Compressor Inlets

NASA Technical Reports Server (NTRS)

Jones, William L.; Dowman, Harry W.

1947-01-01

Investigations were conducted to determine effectiveness of refrigerants in increasing thrust of turbojet engines. Mixtures of water an alcohol were injected for a range of total flows up to 2.2 lb/sec. Kerosene was injected into inlets covering a range of injected flows up to approximately 30% of normal engine fuel flow. Injection of 2.0 lb/sec of water alone produced an increase in thrust of 35.8% of rate engine conditions and kerosene produced a negligible increase in thrust. Carbon dioxide increased thrust 23.5 percent.

Computer Gaming at Every Age: A Comparative Evaluation of Alice

ERIC Educational Resources Information Center

Seals, Cheryl D.; McMillian, Yolanda; Rouse, Kenneth; Agarwal, Ravikant; Johnson, Andrea Williams; Gilbert, Juan E.; Chapman, Richard

2008-01-01

This research has two thrusts of teaching object oriented programming to very young audiences and of increasing student excitement about computing applications with the long-term goal of increasing involvement in technology classes, in the use of computer applications and interest in technology careers. The goal of this work was to provide…

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.

Analytical investigations in aircraft and spacecraft trajectory optimization and optimal guidance

NASA Technical Reports Server (NTRS)

Markopoulos, Nikos; Calise, Anthony J.

1995-01-01

A collection of analytical studies is presented related to unconstrained and constrained aircraft (a/c) energy-state modeling and to spacecraft (s/c) motion under continuous thrust. With regard to a/c unconstrained energy-state modeling, the physical origin of the singular perturbation parameter that accounts for the observed 2-time-scale behavior of a/c during energy climbs is identified and explained. With regard to the constrained energy-state modeling, optimal control problems are studied involving active state-variable inequality constraints. Departing from the practical deficiencies of the control programs for such problems that result from the traditional formulations, a complete reformulation is proposed for these problems which, in contrast to the old formulation, will presumably lead to practically useful controllers that can track an inequality constraint boundary asymptotically, and even in the presence of 2-sided perturbations about it. Finally, with regard to s/c motion under continuous thrust, a thrust program is proposed for which the equations of 2-dimensional motion of a space vehicle in orbit, viewed as a point mass, afford an exact analytic solution. The thrust program arises under the assumption of tangential thrust from the costate system corresponding to minimum-fuel, power-limited, coplanar transfers between two arbitrary conics. The thrust program can be used not only with power-limited propulsion systems, but also with any propulsion system capable of generating continuous thrust of controllable magnitude, and, for propulsion types and classes of transfers for which it is sufficiently optimal the results of this report suggest a method of maneuvering during planetocentric or heliocentric orbital operations, requiring a minimum amount of computation; thus uniquely suitable for real-time feedback guidance implementations.

Saturn Apollo Program

NASA Image and Video Library

1968-01-09

A cluster of eight H-1 engines were used to thrust the first stage of Saturn I (S-I stage) and Saturn IB (S-IB stage). The engines were arranged in a double pattern. Four engines, located inboard, were fixed in a square pattern around the stage axis, while the remaining four engines were located outboard in a larger square pattern and each outer engine was gimbaled. Each H-1 engine, fueled with liquid oxygen (LOX) and kerosene (RP-1), initially had a thrust of 188,000 pounds each for a combined thrust of over 1,500,000 pounds. Later, the H-1 engine was upgraded to 205,000 pounds of thrust and a combined total thrust of 1,650,000 pounds for the Saturn IB program. This photo depicts a single modified H-1 engine. The H-1 engine was developed under the direction of Marshall Space Flight Center (MSFC).

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.

Structural analysis of cylindrical thrust chambers, volume 3

NASA Technical Reports Server (NTRS)

Pearson, M. L.

1981-01-01

A system of three computer programs is described for use in conjunction with the BOPAGE finite element program. The programs are demonstrated by analyzing cumulative plastic deformation in a regeneratively cooled rocket thrust chamber. The codes provide the capability to predict geometric and material nonlinear behavior of cyclically loaded structures without performing a cycle-by-cycle analysis over the life of the structure. The program set consists of a BOPACE restart tape reader routine, and extrapolation program and a plot package.

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.

Thrust Area Report, Engineering Research, Development and Technology

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

Langland, R. T.

1997-02-01

The mission of the Engineering Research, Development, and Technology Program at Lawrence Livermore National Laboratory (LLNL) is to develop the knowledge base, process technologies, specialized equipment, tools and facilities to support current and future LLNL programs. Engineering`s efforts are guided by a strategy that results in dual benefit: first, in support of Department of Energy missions, such as national security through nuclear deterrence; and second, in enhancing the nation`s economic competitiveness through our collaboration with U.S. industry in pursuit of the most cost- effective engineering solutions to LLNL programs. To accomplish this mission, the Engineering Research, Development, and Technology Programmore » has two important goals: (1) identify key technologies relevant to LLNL programs where we can establish unique competencies, and (2) conduct high-quality research and development to enhance our capabilities and establish ourselves as the world leaders in these technologies. To focus Engineering`s efforts technology {ital thrust areas} are identified and technical leaders are selected for each area. The thrust areas are comprised of integrated engineering activities, staffed by personnel from the nine electronics and mechanical engineering divisions, and from other LLNL organizations. This annual report, organized by thrust area, describes Engineering`s activities for fiscal year 1996. The report provides timely summaries of objectives, methods, and key results from eight thrust areas: Computational Electronics and Electromagnetics; Computational Mechanics; Microtechnology; Manufacturing Technology; Materials Science and Engineering; Power Conversion Technologies; Nondestructive Evaluation; and Information Engineering. Readers desiring more information are encouraged to contact the individual thrust area leaders or authors. 198 refs., 206 figs., 16 tabs.« less

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.

Some effects of cyclic induced deformation in rocket thrust chambers

NASA Technical Reports Server (NTRS)

Hannum, N. P.; Quentmeyer, R. J.

1979-01-01

A test program to investigate the deformation process observed in the hot gas wall of rocket thrust chambers was conducted using three different liner materials. Five thrust chambers were cycled to failure using hydrogen and oxygen as propellants at a chamber pressure of 4.14 MN/m square (600 psia). The deformation was observed nondestructively at midlife points and destructively after failure occurred. The cyclic life results are presented with an accompanying discussion about the types of failure encountered. Data indicating the deformation of the thrust chamber liner as cycles are accumulated are presented for each of the test thrust chambers.

Static Performance of Six Innovative Thrust Reverser Concepts for Subsonic Transport Applications: Summary of the NASA Langley Innovative Thrust Reverser Test Program

NASA Technical Reports Server (NTRS)

Asbury, Scott C.; Yetter, Jeffrey A.

2000-01-01

The NASA Langley Configuration Aerodynamics Branch has conducted an experimental investigation to study the static performance of innovative thrust reverser concepts applicable to high-bypass-ratio turbofan engines. Testing was conducted on a conventional separate-flow exhaust system configuration, a conventional cascade thrust reverser configuration, and six innovative thrust reverser configurations. The innovative thrust reverser configurations consisted of a cascade thrust reverser with porous fan-duct blocker, a blockerless thrust reverser, two core-mounted target thrust reversers, a multi-door crocodile thrust reverser, and a wing-mounted thrust reverser. Each of the innovative thrust reverser concepts offer potential weight savings and/or design simplifications over a conventional cascade thrust reverser design. Testing was conducted in the Jet-Exit Test Facility at NASA Langley Research Center using a 7.9%-scale exhaust system model with a fan-to-core bypass ratio of approximately 9.0. All tests were conducted with no external flow and cold, high-pressure air was used to simulate core and fan exhaust flows. Results show that the innovative thrust reverser concepts achieved thrust reverser performance levels which, when taking into account the potential for system simplification and reduced weight, may make them competitive with, or potentially more cost effective than current state-of-the-art thrust reverser systems.

Engineering research, development and technology report

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

Langland, R T

1999-02-01

Nineteen ninety-eight has been a transition year for Engineering, as we have moved from our traditional focus on thrust areas to a more focused approach with research centers. These five new centers of excellence collectively comprise Engineering's Science and Technology program. This publication summarizes our formative year under this new structure. Let me start by talking about the differences between a thrust area and a research center. The thrust area is more informal, combining an important technology with programmatic priorities. In contrast, a research center is directly linked to an Engineering core technology. It is the purer model, for itmore » is more enduring yet has the scope to be able to adapt quickly to evolving programmatic priorities. To put it another way, the mission of a thrust area was often to grow the programs in conjunction with a technology, whereas the task of a research center is to vigorously grow our core technologies. By cultivating each core technology, we in turn enable long-term growth of new programs.« less

Implementation of a Low-Thrust Trajectory Optimization Algorithm for Preliminary Design

NASA Technical Reports Server (NTRS)

Sims, Jon A.; Finlayson, Paul A.; Rinderle, Edward A.; Vavrina, Matthew A.; Kowalkowski, Theresa D.

2006-01-01

A tool developed for the preliminary design of low-thrust trajectories is described. The trajectory is discretized into segments and a nonlinear programming method is used for optimization. The tool is easy to use, has robust convergence, and can handle many intermediate encounters. In addition, the tool has a wide variety of features, including several options for objective function and different low-thrust propulsion models (e.g., solar electric propulsion, nuclear electric propulsion, and solar sail). High-thrust, impulsive trajectories can also be optimized.

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 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.

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.

Space shuttle maneuvering engine reusable thrust chamber program. Task 11: Low Epsilon stability test plan

NASA Technical Reports Server (NTRS)

Pauckert, R. P.

1974-01-01

The performance and heat transfer characteristics of a doublet element type injector for the space shuttle orbiter maneuvering engine thrust chamber were investigated. Ths stability characteristics were evaluated over a range of chamber pressures and mixture ratios. The specific objectives of the test were: (1) to determine whether stability has been influenced by injection of boundary layer coolant across the cavity entrance, (2) if the injector is stable, to determine the minimum cavity area required to maintain stability, and (3) if the injector is unstable, to determine the effects of entrance geometry and increased area on stability.

In-flight thrust determination

NASA Technical Reports Server (NTRS)

Abernethy, Robert B.; Adams, Gary R.; Ascough, John C.; Baer-Riedhart, Jennifer L.; Balkcom, George H.; Biesiadny, Thomas

1986-01-01

The major aspects of processes that may be used for the determination of in-flight thrust are reviewed. Basic definitions are presented as well as analytical and ground-test methods for gathering data and calculating the thrust of the propulsion system during the flight development program of the aircraft. Test analysis examples include a single-exhaust turbofan, an intermediate-cowl turbofan, and a mixed-flow afterburning turbofan.

Saturn Apollo Program

NASA Image and Video Library

1969-01-01

Workers at the Michoud Assembly Facility (MAF) near New Orleans, Louisiana, hoist the thrust structure assembly for the Saturn IB S-IB (first) stage. Developed by the Marshall Space Flight Center and built by the Chrysler Corporation at Michoud Assembly Facility (MAF), the S-IB utilized eight H-1 engines and each produced 200,000 pounds of thrust, a combined thrust of 1,600,000 pounds.

Overview of mechanics of materials branch activities in the computational structures area

NASA Technical Reports Server (NTRS)

Poe, C. C., Jr.

1992-01-01

Base programs and system programs are discussed. The base programs include fundamental research of composites and metals for airframes leading to characterization of advanced materials, models of behavior, and methods for predicting damage tolerance. Results from the base programs support the systems programs, which change as NASA's missions change. The National Aerospace Plane (NASP), Advanced Composites Technology (ACT), Airframe Structural Integrity Program (Aging Aircraft), and High Speed Research (HSR) programs are currently being supported. Airframe durability is one of the key issues in each of these system programs. The base program has four major thrusts, which will be reviewed subsequently. Additionally, several technical highlights will be reviewed for each thrust.

RS-88 Pad Abort Demonstrator Thrust Chamber Assembly Testing at NASA Marshall Space Flight Center

NASA Technical Reports Server (NTRS)

Farr, Rebecca A.; Sanders, Timothy M.

1990-01-01

This paper documents the effort conducted to collect hot-tire dynamic and acoustics environments data during 50,000-lb thrust lox-ethanol hot-fire rocket testing at NASA Marshall Space Flight Center (MSFC) in November-December 2003. This test program was conducted during development testing of the Boeing Rocketdyne RS-88 development engine thrust chamber assembly (TCA) in support of the Orbital Space Plane (OSP) Crew Escape System Propulsion (CESP) Program Pad Abort Demonstrator (PAD). In addition to numerous internal TCA and nozzle measurements, induced acoustics environments data were also collected. Provided here is an overview of test parameters, a discussion of the measurements, test facility systems and test operations, and a quality assessment of the data collected during this test program.

Implementation of the Orbital Maneuvering Systems Engine and Thrust Vector Control for the European Service Module

NASA Technical Reports Server (NTRS)

Millard, Jon

2014-01-01

The European Space Agency (ESA) has entered into a partnership with the National Aeronautics and Space Administration (NASA) to develop and provide the Service Module (SM) for the Orion Multipurpose Crew Vehicle (MPCV) Program. The European Service Module (ESM) will provide main engine thrust by utilizing the Space Shuttle Program Orbital Maneuvering System Engine (OMS-E). Thrust Vector Control (TVC) of the OMS-E will be provided by the Orbital Maneuvering System (OMS) TVC, also used during the Space Shuttle Program. NASA will be providing the OMS-E and OMS TVC to ESA as Government Furnished Equipment (GFE) to integrate into the ESM. This presentation will describe the OMS-E and OMS TVC and discuss the implementation of the hardware for the ESM.

Solar electric propulsion thrust subsystem development

NASA Technical Reports Server (NTRS)

Masek, T. D.

1973-01-01

The Solar Electric Propulsion System developed under this program was designed to demonstrate all the thrust subsystem functions needed on an unmanned planetary vehicle. The demonstration included operation of the basic elements, power matching input and output voltage regulation, three-axis thrust vector control, subsystem automatic control including failure detection and correction capability (using a PDP-11 computer), operation of critical elements in thermal-vacuum-, zero-gravity-type propellant storage, and data outputs from all subsystem elements. The subsystem elements, functions, unique features, and test setup are described. General features and capabilities of the test-support data system are also presented. The test program culminated in a 1500-h computer-controlled, system-functional demonstration. This included simultaneous operation of two thruster/power conditioner sets. The results of this testing phase satisfied all the program goals.

High- and low-thrust propulsion systems for the space station

NASA Technical Reports Server (NTRS)

Jones, R. E.

1987-01-01

The purpose of the Advanced Development program was to investigate propulsion options for the space station. Two options were investigated in detail: a high-thrust system consisting of 25 to 50 lbf gaseous oxygen/hydrogen rockets, and a low-thrust system of 0.1 lbf multipropellant resistojets. An effort is also being conducted to determine the life capability of hydrazine-fueled thrusters. During the course of this program, studies clearly identified the benefits of utilizing waste water and other fluids as propellant sources. The results of the H/O thruster test programs are presented and the plan to determine the life of hydrazine thrusters is discussed. The background required to establish a long-life resistojet is presented and the first design model is shown in detail.

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.

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.

Design and Analysis of an Electromagnetic Thrust Bearing

NASA Technical Reports Server (NTRS)

Banerjee, Bibhuti; Rao, Dantam K.

1996-01-01

A double-acting electromagnetic thrust bearing is normally used to counter the axial loads in many rotating machines that employ magnetic bearings. It essentially consists of an actuator and drive electronics. Existing thrust bearing design programs are based on several assumptions. These assumptions, however, are often violated in practice. For example, no distinction is made between maximum external loads and maximum bearing forces, which are assumed to be identical. Furthermore, it is assumed that the maximum flux density in the air gap occurs at the nominal gap position of the thrust runner. The purpose of this paper is to present a clear theoretical basis for the design of the electromagnetic thrust bearing which obviates such assumptions.

Thrust faults and related structures in the crater floor of Mount St. Helens volcano, Washington

USGS Publications Warehouse

Chadwick, W.W.; Swanson, D.A.

1989-01-01

A lava dome was built in the crater of Mount St. Helens by intermittent intrusion and extrusion of dacite lava between 1980 and 1986. Spectacular ground deformation was associated with the dome-building events and included the development of a system of radial cracks and tangential thrust faults in the surrounding crater floor. These cracks and thrusts, best developed and studied in 1981-1982, formed first and, as some evolved into strike-slip tear faults, influenced the subsequent geometry of thrusting. Once faulting began, deformation was localized near the thrust scarps and their bounding tear faults. The magnitude of displacements systematically increased before extrusions, whereas the azimuth and inclination of displacements remained relatively constant. The thrust-fault scarps were bulbous in profile, lobate in plan, and steepened during continued fault movement. The hanging walls of each thrust were increasingly disrupted as cumulative fault slip increased. -from Authors

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.

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.

SSTAC/ARTS Review of the Draft Integrated Technology Plan (ITP). Volume 2: Propulsion Systems

NASA Technical Reports Server (NTRS)

1991-01-01

The topics addressed are: (1) space propulsion technology program overview; (2) space propulsion technology program fact sheet; (3) low thrust propulsion; (4) advanced propulsion concepts; (5) high-thrust chemical propulsion; (6) cryogenic fluid management; (7) NASA CSTI earth-to-orbit propulsion; (8) advanced main combustion chamber program; (9) earth-to-orbit propulsion turbomachinery; (10) transportation technology; (11) space chemical engines technology; (12) nuclear propulsion; (13) spacecraft on-board propulsion; and (14) low-cost commercial transport.

CFD Code Survey for Thrust Chamber Application

NASA Technical Reports Server (NTRS)

Gross, Klaus W.

1990-01-01

In the quest fo find analytical reference codes, responses from a questionnaire are presented which portray the current computational fluid dynamics (CFD) program status and capability at various organizations, characterizing liquid rocket thrust chamber flow fields. Sample cases are identified to examine the ability, operational condition, and accuracy of the codes. To select the best suited programs for accelerated improvements, evaluation criteria are being proposed.

Comparisons between a high resolution discrete element model and analogue model

NASA Astrophysics Data System (ADS)

LI, C. S.; Yin, H.; WU, C.; Zhang, J.

2017-12-01

A two-dimensional discrete element model (DEM) with high resolution is constructed to simulate the evolution of thrust wedge and an analogue model (AM) experiment is constructed to compare with the DEM results. This efficient parallel DEM program is written in the C language, and it is useful to solve the complex geological problems. More detailed about fold and thrust belts of DEM can be identified with the help of strain field. With non-rotating and non-tensile assumption, dynamic evolution of DEM is highly consistent with AM. Simulations in different scale can compare with each other by conversion formulas in DEM. Our results show that: (1) The overall evolution of DEM and AM is broadly similar. (2) Shortening is accommodated by in-sequence forward propagation of thrusts. The surface slope of the thrust wedge is within the stable field predicted by critical taper theory. (3) Details of thrust spacing, dip angle and number of thrusts vary between DEM and AM for the shortening experiment, but the characteristics of thrusts are similar on the whole. (4) Dip angles of the forward thrusts increased from foreland (ca. 30°) to the mobile wall (ca. 80°) (5) With shortening, both models had not the obvious volume loss. Instead, the volume basic remained unchanged in the whole extrusion processes. (6) Almost all high strain values are within fold-and-thrust belts in DEM, which allows a direct comparison between the fault zone identified on the DEM deformation field and that in the strain field. (7) The first fault initiates at deep depths and propagate down toward the surface. For the maximal volumetric strain focused on the décollement near the mobile wall, strengthening the material and making it for brittle. (8) With non-tensile particles for DEM, contraction is broadly distributed throughout the model and dilation is hardly any, which also leads to a higher efficient computation. (9) High resolution DEM can to first order successfully reproduce structures observed in AM. The comparisons serve to highlight robust features in tectonic modelling of thrust wedges. This approach is very utility in modelling large displacement, complex deformation of analogue and geological materials.

A hydrogen-oxygen rocket engine coolant passage design program (RECOP) for fluid-cooled thrust chambers and nozzles

NASA Technical Reports Server (NTRS)

Tomsik, Thomas M.

1994-01-01

The design of coolant passages in regeneratively cooled thrust chambers is critical to the operation and safety of a rocket engine system. Designing a coolant passage is a complex thermal and hydraulic problem requiring an accurate understanding of the heat transfer between the combustion gas and the coolant. Every major rocket engine company has invested in the development of thrust chamber computer design and analysis tools; two examples are Rocketdyne's REGEN code and Aerojet's ELES program. In an effort to augment current design capabilities for government and industry, the NASA Lewis Research Center is developing a computer model to design coolant passages for advanced regeneratively cooled thrust chambers. The RECOP code incorporates state-of-the-art correlations, numerical techniques and design methods, certainly minimum requirements for generating optimum designs of future space chemical engines. A preliminary version of the RECOP model was recently completed and code validation work is in progress. This paper introduces major features of RECOP and compares the analysis to design points for the first test case engine; the Pratt & Whitney RL10A-3-3A thrust chamber.

A Regeneratively Cooled Thrust Chamber For The Fastrac Engine

NASA Technical Reports Server (NTRS)

Brown, Kendall K.; Sparks, Dave; Woodcock, Gordon

2000-01-01

Abstract This paper presents the development of a low-cost, regeneratively-cooled thrust chamber for the Fastrac engine. The chamber was fabricated using hydraformed copper tubing to form the coolant jacket and wrapped with a fiber reinforced polymer composite Material to form a structural jacket. The thrust chamber design and fabrication approach was based upon Space America. Inc.'s 12,000 lb regeneratively-cooled LOX/kerosene rocket engine. Fabrication of regeneratively cooled thrust chambers by tubewall construction dates back to the early US ballistic missile programs. The most significant innovations in this design was the development of a low-cost process for fabrication from copper tubing (nickel alloy was the usual practice) and use of graphite composite overwrap as the pressure containment, which yields an easily fabricated, lightweight pressure jacket around the copper tubes A regeneratively-cooled reusable thrust chamber can benefit the Fastrac engine program by allowing more efficient (cost and scheduler testing). A proof-of-concept test article has been fabricated and will he tested at Marshall Space Flight Center in the late Summer or Fall of 2000.

Thrust modeling for hypersonic engines

NASA Technical Reports Server (NTRS)

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

1995-01-01

Expressions for the thrust losses of a scramjet engine are developed in terms of irreversible entropy increases and the degree of incomplete combustion. A method is developed which allows the calculation of the lost vehicle thrust due to different loss mechanisms within a given flow-field. This analysis demonstrates clearly the trade-off between mixing enhancement and resultant increased flow losses in scramjet combustors. An engine effectiveness parameter is defined in terms of thrust loss. Exergy and the thrust-potential method are related and compared.

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.

Helicon double layer thruster operation in a low magnetic field mode

NASA Astrophysics Data System (ADS)

Harle, T.; Pottinger, S. J.; Lappas, V. J.

2013-02-01

Direct thrust measurements are made of a helicon double layer thruster operating in a low magnetic field mode. The relationship between the imposed axial magnetic field and generated thrust is investigated for a radio frequency input power range 200-500 W for propellant flow rates of 16.5 and 20 sccm (0.46 and 0.55 mg s-1) of argon. The measured thrust shows a strong dependence on the magnetic field strength, increasing by up to a factor of 5 compared with the minimum thrust level recorded. A peak thrust of 0.4-1.1 mN depending on thruster operating conditions is obtained. This increase is observed to take place over a small range of peak magnetic field strengths in the region of 70-110 G. The magnitude of the thrust and the corresponding magnitude of the magnetic field at which the peak thrust occurs is shown to increase with increasing input power for a given propellant flow rate. The ion current determined using a retarding field energy analyser and the electron number density found using a microwave resonator probe both correlate with the observed trend in thrust as a function of applied magnetic field.

Saturn Apollo Program

NASA Image and Video Library

1969-01-01

In one of the initial assembly steps for the Saturn IB launch vehicle's S-IB (first) stage, workers at the Michoud Assembly Facility (MAF) near New Orleans, Louisiana, position the thrust structure. Developed by the Marshall Space Flight Center and built by the Chrysler Corporation at Michoud Assembly Facility (MAF), the S-IB utilized eight H-1 engines and each produced 200,000 pounds of thrust, a combined thrust of 1,600,000 pounds.

Saturn Apollo Program

NASA Image and Video Library

1969-01-01

In one of the initial assembly steps for the Saturn IB launch vehicle's S-IB (first) stage, workers at the Michoud Assembly Facility (MAF) near New Orleans, Louisiana, complete the thrust structure. Developed by the Marshall Space Flight Center and built by the Chrysler Corporation at Michoud Assembly Facility (MAF), the S-IB utilized eight H-1 engines and each produced 200,000 pounds of thrust, a combined thrust of 1,600,000 pounds.

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.

High Technology in the Vocational Agriculture Classroom.

ERIC Educational Resources Information Center

Camp, William G.

While many aspects of high technology may be directly applicable in the vocational agricultural classroom, the primary thrust of high technology into such programs, at least in the short range, will be centered around the microcomputer. Because of its cost and availability, the microcomputer will continue to play an ever increasing role in…

Space electronics technology summary

NASA Technical Reports Server (NTRS)

1976-01-01

An overview is given of current electronics R and D activities, potential future thrusts, and related NASA payoffs. Major increases in NASA mission return and significant concurrent reductions in mission cost appear possible through a focused, long range electronics technology program. The overview covers: guidance assessments, navigation and control, and sensing and data acquisition processing, storage, and transfer.

Using Action Learning and Critical Thinking Tools to Make Changes in Higher Education

ERIC Educational Resources Information Center

Smith, Janice Witt; Stitts, D. Kathy

2013-01-01

As the reported institution increased admission requirements, added master's and doctoral programs, and revised its strategic thrust, greater emphasis was placed on assessing student learning outcomes and whether the "treatment" of education has had a measurable impact. Consistent with Smith and Clark (2010), we implemented an action…

Evaluation of rotating, incompressibly lubricated, pressurized thrust bearings

NASA Technical Reports Server (NTRS)

Fleming, D. P.

1971-01-01

Program evaluates a series hybrid, fluid film ball bearing consisting of an orifice compensated pressurized thrust bearing in conjunction with a self-acting journal bearing. Oil viscosities corresponding to experimentally measured ball bearing outer-race temperatures were used.

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.

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.

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.

Integrated flight/propulsion control - Adaptive engine control system mode

NASA Technical Reports Server (NTRS)

Yonke, W. A.; Terrell, L. A.; Meyers, L. P.

1985-01-01

The adaptive engine control system mode (ADECS) which is developed and tested on an F-15 aircraft with PW1128 engines, using the NASA sponsored highly integrated digital electronic control program, is examined. The operation of the ADECS mode, as well as the basic control logic, the avionic architecture, and the airframe/engine interface are described. By increasing engine pressure ratio (EPR) additional thrust is obtained at intermediate power and above. To modulate the amount of EPR uptrim and to prevent engine stall, information from the flight control system is used. The performance benefits, anticipated from control integration are shown for a range of flight conditions and power settings. It is found that at higher altitudes, the ADECS mode can increase thrust as much as 12 percent, which is used for improved acceleration, improved turn rate, or sustained turn angle.

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.

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°.

U.S. Geological Survey toxic Waste-Groundwater Contamination Program, fiscal year 1985

NASA Astrophysics Data System (ADS)

Ragone, Stephen E.

1986-09-01

In fiscal year 1982, the U S Geological Survey began an interdisciplinary research thrust entitled Toxic Waste-Groundwater Contamination Program The objective of the thrust was to provide earth sciences information necessary to evaluate and mitigate existing groundwater contamination problems resulting from the planned or inadvertant disposal of wastes and from certain land-use practices, and to improve future waste disposal and land-use practices The program supports process-oriented and interdisciplinary field research, and regional groundwater quality studies This article provides an overview of the current (Fiscal Year 1985) activities of the Toxic Waste Program

Interaction Behavior between Thrust Faulting and the National Highway No. 3 - Tianliao III bridge as Determined using Numerical Simulation

NASA Astrophysics Data System (ADS)

Li, C. H.; Wu, L. C.; Chan, P. C.; Lin, M. L.

2016-12-01

The National Highway No. 3 - Tianliao III Bridge is located in the southwestern Taiwan mudstone area and crosses the Chekualin fault. Since the bridge was opened to traffic, it has been repaired 11 times. To understand the interaction behavior between thrust faulting and the bridge, a discrete element method-based software program, PFC, was applied to conduct a numerical analysis. A 3D model for simulating the thrust faulting and bridge was established, as shown in Fig. 1. In this conceptual model, the length and width were 50 and 10 m, respectively. Part of the box bottom was moveable, simulating the displacement of the thrust fault. The overburden stratum had a height of 5 m with fault dip angles of 20° (Fig. 2). The bottom-up strata were mudstone, clay, and sand, separately. The uplift was 1 m, which was 20% of the stratum thickness. In accordance with the investigation, the position of the fault tip was set, depending on the fault zone, and the bridge deformation was observed (Fig. 3). By setting "Monitoring Balls" in the numerical model to analyzes bridge displacement, we determined that the bridge deck deflection increased as the uplift distance increased. Furthermore, the force caused by the loading of the bridge deck and fault dislocation was determined to cause a down deflection of the P1 and P2 bridge piers. Finally, the fault deflection trajectory of the P4 pier displayed the maximum displacement (Fig. 4). Similar behavior has been observed through numerical simulation as well as field monitoring data. Usage of the discrete element model (PFC3D) to simulate the deformation behavior between thrust faulting and the bridge provided feedback for the design and improved planning of the bridge.

Sensitivity of Runway Occupancy Time (ROT) to Various Rollout and Turnoff (ROTO) Factors. Volume 1

NASA Technical Reports Server (NTRS)

Goldthorpe, S. H.

1997-01-01

The Terminal Area Productivity (TAP) research program was initiated by NASA to increase the airport capacity for transport aircraft operations. One element of the research program is called Low Visibility Landing and Surface Operations (LVLASO). A goal of the LVLASO research is to develop transport aircraft technologies which reduce Runway Occupancy Time (ROT) so that it does not become the limiting factor in the terminal area operations that determine the capacity of a runway. Under LVLASO, the objective of this study was to determine the sensitivity of ROT to various factors associated with the Rollout and Turnoff (ROTO) operation for transport aircraft. The following operational factors were studied and are listed in the order of decreasing ROT sensitivity: ice/flood runway surface condition, exit entrance ground speed, number of exits, high-speed exit locations and spacing, aircraft type, touchdown ground speed standard deviation, reverse thrust and braking method, accurate exit prediction capability, maximum reverse thrust availability, spiral-arc vs. circle-arc exit geometry, dry/slush/wet/snow runway surface condition, maximum allowed deceleration, auto asymmetric braking on exit, do not stow reverse thrust before the exit, touchdown longitudinal location standard deviation, flap setting, anti-skid efficiency, crosswind conditions, stopping on the exit and touchdown lateral offset.

Sensitivity of Runway Occupancy Time (ROT) to Various Rollout and Turnoff (ROTO) Factors. Volume 2; Complete Set of Plotted Data

NASA Technical Reports Server (NTRS)

Goldthorpe, S. H.

1997-01-01

The Terminal Area Productivity (TAP) research program was initiated by NASA to increase the airport capacity for transport aircraft operations. One element of the research program is called Low Visibility Landing and Surface Operations (LVLASO). A goal of the LVLASO research is to develop transport aircraft technologies which reduce Runway Occupancy Time (ROT) so that it does not become the limiting factor in the terminal area operations that determine the capacity of a runway. Under LVLASO, the objective of this study was to determine the sensitivity of ROT to various factors associated with the Rollout and Turnoff (ROTO) operation for transport aircraft. The following operational factors were studied and are listed in the order of decreasing ROT sensitivity: ice/flood runway surface condition, exit entrance ground speed, number of exits, high-speed exit locations and spacing, aircraft type, touchdown ground speed standard deviation, reverse thrust and braking method, accurate exit prediction capability, maximum reverse thrust availability, spiral-arc vs. circle-arc exit geometry, dry/slush/wet/snow runway surface condition, maximum allowed deceleration, auto asymmetric braking on exit, do not stow reverse thrust before the exit, touchdown longitudinal location standard deviation, flap setting, anti-skid efficiency, crosswind conditions, stopping on the exit and touchdown lateral offset.

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.

Comparative wind tunnel test at high Reynolds numbers of NACA 64 621 airfoils with two aileron configurations

NASA Technical Reports Server (NTRS)

Gregorek, G. M.

1995-01-01

An experimental program to measure the aerodynamic characteristics of the NACA 64-621 airfoil when equipped with plain ailerons of 0.38 chord and 0.30 chord and with 0.38 chord balanced aileron has been conducted in the pressurized O.S.U. 6 x 12 ft High Reynolds Number Wind Tunnel. Surface pressures were measured and integrated to yield lift and pressure drag coefficients for angles of attack from -3 to +42 deg and for selected aileron deflections from 0 to -90 deg at nominal Mach and Reynolds numbers of 0.25 and 5 x 10(exp 6). When resolved into thrust coefficient for wind turbine aerodynamic control applications, the data indicated the anticipated decrease in thrust coefficient with negative aileron deflection at low angles of attack; however, as angle of attack increased, thrust coefficients eventually became positive. All aileron configurations, even at -90 deg deflections showed this trend. Hinge moments for each configuration complete the data set.

Performance improvements of a highly integrated digital electronic control system for an F-15 airplane

NASA Technical Reports Server (NTRS)

Putnam, T. W.; Burcham, F. W., Jr.; Andries, M. G.; Kelly, J. B.

1985-01-01

The NASA highly integrated digital electronic control (HIDEC) program is structured to conduct flight research into the benefits of integrating an aircraft flight control system with the engine control system. A brief description of the HIDEC system installed on an F-15 aircraft is provided. The adaptive engine control system (ADECS) mode is described in detail, together with simulation results and analyses that show the significant excess thrust improvements achievable with the ADECS mode. It was found that this increased thrust capability is accompanied by reduced fan stall margin and can be realized during flight conditions where engine face distortion is low. The results of analyses and simulations also show that engine thrust response is improved and that fuel consumption can be reduced. Although the performance benefits that accrue because of airframe and engine control integration are being demonstrated on an F-15 aircraft, the principles are applicable to advanced aircraft such as the advanced tactical fighter and advanced tactical aircraft.

3D And 4D Cloud Lifecycle Investigations Using Innovative Scanning Radar Analysis Methods. Final report

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

Kollias, Pavlos

2017-04-23

With the vast upgrades to the ARM program radar measurement capabilities in 2010 and beyond, our ability to probe the 3D structure of clouds and associated precipitation has increased dramatically. This project build on the PI's and co-I's expertisein the analysis of radar observations. The first research thrust aims to document the 3D morphological (as depicted by the radar reflectivity structure) and 3D dynamical (cloud$-$scale eddies) structure of boundary layer clouds. Unraveling the 3D dynamical structure of stratocumulus and shallow cumulus clouds requires decomposition of the environmental wind contribution and particle sedimentation velocity from the observed radial Doppler velocity. Themore » second thrust proposes to unravel the mechanism of cumulus entrainment (location, scales) and its impact on microphysics utilizing radar measurements from the vertically pointing and new scanning radars at the ARM sites. The third research thrust requires the development of a cloud$-$tracking algorithm that monitors the properties of cloud.« less

Tests on Thrust Augmenters for Jet Propulsion

NASA Technical Reports Server (NTRS)

Jacobs, Eastman N; Shoemaker, James M

1932-01-01

This series of tests was undertaken to determine how much the reaction thrust of a jet could be increased by the use of thrust augmenters and thus to give some indication as to the feasibility of jet propulsion for airplanes. The tests were made during the first part of 1927 at the Langley Memorial Aeronautical Laboratory. A compressed air jet was used in connection with a series of annular guides surrounding the jet to act as thrust augmenters. The results show that, although it is possible to increase the thrust of a jet, the increase is not large enough to affect greatly the status of the problem of the application of jet propulsion to airplanes.

Hybrid Propulsion Demonstration Program 250K Hybrid Motor

NASA Technical Reports Server (NTRS)

Story, George; Zoladz, Tom; Arves, Joe; Kearney, Darren; Abel, Terry; Park, O.

2003-01-01

The Hybrid Propulsion Demonstration Program (HPDP) program was formed to mature hybrid propulsion technology to a readiness level sufficient to enable commercialization for various space launch applications. The goal of the HPDP was to develop and test a 250,000 pound vacuum thrust hybrid booster in order to demonstrate hybrid propulsion technology and enable manufacturing of large hybrid boosters for current and future space launch vehicles. The HPDP has successfully conducted four tests of the 250,000 pound thrust hybrid rocket motor at NASA's Stennis Space Center. This paper documents the test series.

Fourth National Aeronautics and Space Administration Weather and Climate Program Science Review

NASA Technical Reports Server (NTRS)

Kreins, E. R. (Editor)

1979-01-01

The NASA Weather and Climate Program has two major thrusts. The first involves the development of experimental and prototype operational satellite systems, sensors, and space facilities for monitoring and understanding the atmosphere. The second thrust involves basic scientific investigation aimed at studying the physical and chemical processes which control weather and climate. This fourth science review concentrated on the scientific research rather than the hardware development aspect of the program. These proceedings contain 65 papers covering the three general areas: severe storms and local weather research, global weather, and climate.

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

Final Report, University Research Program in Robotics (URPR), Nuclear Facilities Clean-up

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

Tesar, Delbert; Kapoor, Chetan; Pryor, Mitch

This final report describes the research activity at the University of Texas at Austin with application to EM needs at DOE. This research activity is divided in to two major thrusts and contributes to the overall University Research Program in Robotics (URPR) thrust by providing mechanically oriented robotic solutions based on modularity and generalized software. These thrusts are also the core strengths of the UTA program that has a 40-year history in machine development, 30 years specifically devoted to robotics. Since 1975, much of this effort has been to establish the general analytical and design infrastructure for an open (modular)more » architecture of systems with many degrees of freedom that are able to satisfy a broad range of applications for future production machines. This work has coalesced from two principal areas: standardized actuators and generalized software.« less

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

NASA Technical Reports Server (NTRS)

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

1994-01-01

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

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.

Laser-heated thruster

NASA Technical Reports Server (NTRS)

Kemp, N. H.; Krech, R. H.

1980-01-01

The development of computer codes for the thrust chamber of a rocket of which the propellant gas is heated by a CW laser beam was investigated. The following results are presented: (1) simplified models of laser heated thrusters for approximate parametric studies and performance mapping; (3) computer programs for thrust chamber design; and (3) shock tube experiment to measure absorption coefficients. Two thrust chamber design programs are outlined: (1) for seeded hydrogen, with both low temperature and high temperature seeds, which absorbs the laser radiation continuously, starting at the inlet gas temperature; and (2) for hydrogen seeded with cesium, in which a laser supported combustion wave stands near the gas inlet, and heats the gas up to a temperature at which the gas can absorb the laser energy.

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.

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.

Energy Efficient Engine program advanced turbofan nacelle definition study

NASA Technical Reports Server (NTRS)

Howe, David C.; Wynosky, T. A.

1985-01-01

Advanced, low drag, nacelle configurations were defined for some of the more promising propulsion systems identified in the earlier Benefit/Cost Study, to assess the benefits associated with these advanced technology nacelles and formulate programs for developing these nacelles and low volume thrust reversers/spoilers to a state of technology readiness in the early 1990's. The study results established the design feasibility of advanced technology, slim line nacelles applicable to advanced technology, high bypass ratio turbofan engines. Design feasibility was also established for two low volume thrust reverse/spoiler concepts that meet or exceed the required effectiveness for these engines. These nacelle and thrust reverse/spoiler designs were shown to be applicable in engines with takeoff thrust sizes ranging from 24,000 to 60,000 pounds. The reduced weight, drag, and cost of the advanced technology nacelle installations relative to current technology nacelles offer a mission fuel burn savings ranging from 3.0 to 4.5 percent and direct operating cost plus interest improvements from 1.6 to 2.2 percent.

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.

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.

U.S. Geological Survey toxic Waste-Groundwater Contamination Program, fiscal year 1985

USGS Publications Warehouse

Ragone, S.E.

1986-01-01

In fiscal year 1982, the U S Geological Survey began an interdisciplinary research thrust entitled Toxic Waste-Groundwater Contamination Program The objective of the thrust was to provide earth sciences information necessary to evaluate and mitigate existing groundwater contamination problems resulting from the planned or inadvertant disposal of wastes and from certain land-use practices, and to improve future waste disposal and land-use practices The program supports process-oriented and interdisciplinary field research, and regional groundwater quality studies This article provides an overview of the current (Fiscal Year 1985) activities of the Toxic Waste Program ?? 1986 Springer-Verlag New York Inc.

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.

Human Mars Mission Performance Crew Taxi Profile. Part 1

NASA Technical Reports Server (NTRS)

Duaro, Vince A.

1999-01-01

This timeline was generated on the Integrated Mission Program (IMP). All burn events over 2 seconds are finite with IMP solving a two point boundary value setup for begin burn time, burn time and control angles. Perigee and apogee shown above are mean orbital values. Significant events are listed. Each finite thrust event has two lines. The first is the beginning time showing the initial conditions, thrust and ISP used. The second has the end burn conditions and the delta v and time of burn. This case is an abort from the 750 x 750 phasing abort, using the taxi's main engines. An abort using the Reaction Control System (RCS) was also investigated but required a large increase in RCS propellant and was abandoned.

Test and evaluation of the HIDEC engine uptrim algorithm

NASA Technical Reports Server (NTRS)

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

1986-01-01

The highly integrated digital electronic control (HIDEC) program will demonstrate and evaluate the improvements in performance and mission effectiveness that result from integrated engine-airframe control systems. Performance improvements will result from an adaptive engine stall margin mode, a highly integrated mode that uses the airplane flight conditions and the resulting inlet distortion to continuously compute engine stall margin. When there is excessive stall margin, the engine is uptrimmed for more thrust by increasing engine pressure ratio (EPR). The EPR uptrim logic has been evaluated and implemented into computer simulations. Thrust improvements over 10 percent are predicted for subsonic flight conditions. The EPR uptrim was successfully demonstrated during engine ground tests. Test results verify model predictions at the conditions tested.

The F-18 High Alpha Research Vehicle: A High-Angle-of-Attack Testbed Aircraft

NASA Technical Reports Server (NTRS)

Regenie, Victoria; Gatlin, Donald; Kempel, Robert; Matheny, Neil

1992-01-01

The F-18 High Alpha Research Vehicle is the first thrust-vectoring testbed aircraft used to study the aerodynamics and maneuvering available in the poststall flight regime and to provide the data for validating ground prediction techniques. The aircraft includes a flexible research flight control system and full research instrumentation. The capability to control the vehicle at angles of attack up to 70 degrees is also included. This aircraft was modified by adding a pitch and yaw thrust-vectoring system. No significant problems occurred during the envelope expansion phase of the program. This aircraft has demonstrated excellent control in the wing rock region and increased rolling performance at high angles of attack. Initial pilot reports indicate that the increased capability is desirable although some difficulty in judging the size and timing of control inputs was observed. The aircraft, preflight ground testing and envelope expansion flight tests are described.

To flap or not to flap: continued discussion with particle image velocimetry of the near wake

NASA Astrophysics Data System (ADS)

Martin, Nathan; Roh, Chris; Idrees, Suhail; Gharib, Morteza

2017-11-01

We continue the discussion of which underwater propulsion mechanism is more effective: flapping used by fish or periodic contractions used by jellyfish. The two propulsion mechanisms are simplified into flapping and clapping plate motions, respectively, to allow for a direct comparison. A device is designed to operate in either mode of propulsion between Reynolds numbers 1,880 and 11,260, based on the average tip velocity and the span of the plate. The stroke angle, stroke time, flexibility, and duty cycle are varied to determine their impact on the generated thrust and the required torque. Overall, the clapping mode tends to require significantly more power to generate a similar thrust compared to that from the flapping mode. The performance of the clapping mode is increased by modifying the duty cycle such that the closing motion is faster than the opening motion causing a greater thrust and a similar efficiency to that from the flapping mode. Interestingly, when using rigid plates, the average thrust generated per cycle is similar between the two modes when the overall kinematics are equivalent. Investigation of the near wake of both modes through digital particle image velocimetry provides insight into the cause of this similar thrust. This work was supported by the Charyk Bio-inspired Laboratory at the California Institute of Technology, the National Science Foundation Graduate Research Fellowship under Grant No. DGE-1144469, and the Summer Undergraduate Research Fellowships program.

Orbit transfer rocket engine technology program enhanced heat transfer combustor technology

NASA Technical Reports Server (NTRS)

Brown, William S.

1991-01-01

In order to increase the performance of a high performance, advanced expander-cycle engine combustor, higher chamber pressures are required. In order to increase chamber pressure, more heat energy is required to be transferred to the combustor coolant circuit fluid which drives the turbomachinery. This requirement was fulfilled by increasing the area exposed to the hot-gas by using combustor ribs. A previous technology task conducted 2-d hot air and cold flow tests to determine an optimum rib height and configuration. In task C.5 a combustor calorimeter was fabricated with the optimum rib configuration, 0.040 in. high ribs, in order to determine their enhancing capability. A secondary objective was to determine the effects of mixture ratio changers on the enhancement during hot-fire testing. The program used the Rocketdyne Integrated Component Evaluator (ICE) reconfigured into a thrust chamber only mode. The test results were extrapolated to give a projected enhancement from the ribs for a 16 in. long cylindrical combustor at 15 Klb nominal thrust level. The hot-gas wall ribs resulted in a 58 percent increase in heat transfer. When projected to a full size 15K combustor, it becomes a 46 percent increase. The results of those tests, a comparison with previous 2-d results, the effects of mixture ratio and combustion gas flow on the ribs and the potential ramifications for expander cycle combustors are detailed.

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

Fuel-optimal low-thrust formation reconfiguration via Radau pseudospectral method

NASA Astrophysics Data System (ADS)

Li, Jing

2016-07-01

This paper investigates fuel-optimal low-thrust formation reconfiguration near circular orbit. Based on the Clohessy-Wiltshire equations, first-order necessary optimality conditions are derived from the Pontryagin's maximum principle. The fuel-optimal impulsive solution is utilized to divide the low-thrust trajectory into thrust and coast arcs. By introducing the switching times as optimization variables, the fuel-optimal low-thrust formation reconfiguration is posed as a nonlinear programming problem (NLP) via direct transcription using multiple-phase Radau pseudospectral method (RPM), which is then solved by a sparse nonlinear optimization software SNOPT. To facilitate optimality verification and, if necessary, further refinement of the optimized solution of the NLP, formulas for mass costate estimation and initial costates scaling are presented. Numerical examples are given to show the application of the proposed optimization method. To fix the problem, generic fuel-optimal low-thrust formation reconfiguration can be simplified as reconfiguration without any initial and terminal coast arcs, whose optimal solutions can be efficiently obtained from the multiple-phase RPM at the cost of a slight fuel increment. Finally, influence of the specific impulse and maximum thrust magnitude on the fuel-optimal low-thrust formation reconfiguration is analyzed. Numerical results shown the links and differences between the fuel-optimal impulsive and low-thrust solutions.

A possible explanation for foreland thrust propagation

NASA Astrophysics Data System (ADS)

Panian, John; Pilant, Walter

1990-06-01

A common feature of thin-skinned fold and thrust belts is the sequential nature of foreland directed thrust systems. As a rule, younger thrusts develop in the footwalls of older thrusts, the whole sequence propagating towards the foreland in the transport direction. As each new younger thrust develops, the entire sequence is thickened; particularly in the frontal region. The compressive toe region can be likened to an advancing wave; as the mountainous thrust belt advanced the down-surface slope stresses drive thrusts ahead of it much like a surfboard rider. In an attempt to investigate the stresses in the frontal regions of thrustsheets, a numerical method has been devised from the algorithm given by McTigue and Mei [1981]. The algorithm yields a quickly computed approximate solution of the gravity- and tectonic-induced stresses of a two-dimensional homogeneous elastic half-space with an arbitrarily shaped free surface of small slope. A comparison of the numerical method with analytical examples shows excellent agreement. The numerical method was devised because it greatly facilitates the stress calculations and frees one from using the restrictive, simple topographic profiles necessary to obtain an analytical solution. The numerical version of the McTigue and Mei algorithm shows that there is a region of increased maximum resolved shear stress, τ, directly beneath the toe of the overthrust sheet. Utilizing the Mohr-Coulomb failure criterion, predicted fault lines are computed. It is shown that they flatten and become horizontal in some portions of this zone of increased τ. Thrust sheets are known to advance upon weak decollement zones. If there is a coincidence of increased τ, a weak rock layer, and a potential fault line parallel to this weak layer, we have in place all the elements necessary to initiate a new thrusting event. That is, this combination acts as a nucleating center to initiate a new thrusting event. Therefore, thrusts develop in sequence towards the foreland as a consequence of the stress concentrating abilities of the toe of the thrust sheet. The gravity- and tectonic-induced stresses due to the surface topography (usually ignored in previous analyses) of an advancing thrust sheet play a key role in the nature of shallow foreland thrust propagation.

Thrust Augmentation Measurements Using a Pulse Detonation Engine Ejector

NASA Technical Reports Server (NTRS)

Santoro, Robert J.; Pal, Sibtosh

2003-01-01

The present NASA GRC-funded three-year research project is focused on studying PDE driven ejectors applicable to a hybrid Pulse Detonation/Turbofan Engine. The objective of the study is to characterize the PDE-ejector thrust augmentation. A PDE-ejector system has been designed to provide critical experimental data for assessing the performance enhancements possible with this technology. Completed tasks include demonstration of a thrust stand for measuring average thrust for detonation tube multi-cycle operation, and design of a 72-in.-long, 2.25-in.-diameter (ID) detonation tube and modular ejector assembly. This assembly will allow testing of both straight and contoured ejector geometries. Initial ejectors that have been fabricated are 72-in.-long-constant-diameter tubes (4-, 5-, and 6-in.-diameter) instrumented with high-frequency pressure transducers. The assembly has been designed such that the detonation tube exit can be positioned at various locations within the ejector tube. PDE-ejector system experiments with gaseous ethylene/ nitrogen/oxygen propellants will commence in the very near future. The program benefits from collaborations with Prof. Merkle of University of Tennessee whose PDE-ejector analysis helps guide the experiments. The present research effort will increase the TRL of PDE-ejectors from its current level of 2 to a level of 3.

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.

Divergence thrust loss calculations for convergent-divergent nozzles: Extensions to the classical case

NASA Technical Reports Server (NTRS)

Berton, Jeffrey J.

1991-01-01

The analytical derivations of the non-axial thrust divergence losses for convergent-divergent nozzles are described as well as how these calculations are embodied in the Navy/NASA engine computer program. The convergent-divergent geometries considered are simple classic axisymmetric nozzles, two dimensional rectangular nozzles, and axisymmetric and two dimensional plug nozzles. A simple, traditional, inviscid mathematical approach is used to deduce the influence of the ineffectual non-axial thrust as a function of the nozzle exit divergence angle.

Low heat transfer oxidizer heat exchanger design and analysis

NASA Technical Reports Server (NTRS)

Kanic, P. G.; Kmiec, T. D.; Peckham, R. J.

1987-01-01

The RL10-IIB engine, a derivative of the RLIO, is capable of multi-mode thrust operation. This engine operates at two low thrust levels: tank head idle (THI), which is approximately 1 to 2 percent of full thrust, and pumped idle (PI), which is 10 percent of full thrust. Operation at THI provides vehicle propellant settling thrust and efficient engine thermal conditioning; PI operation provides vehicle tank pre-pressurization and maneuver thrust for log-g deployment. Stable combustion of the RL10-IIB engine at THI and PI thrust levels can be accomplished by providing gaseous oxygen at the propellant injector. Using gaseous hydrogen from the thrust chamber jacket as an energy source, a heat exchanger can be used to vaporize liquid oxygen without creating flow instability. This report summarizes the design and analysis of a United Aircraft Products (UAP) low-rate heat transfer heat exchanger concept for the RL10-IIB rocket engine. The design represents a second iteration of the RL10-IIB heat exchanger investigation program. The design and analysis of the first heat exchanger effort is presented in more detail in NASA CR-174857. Testing of the previous design is detailed in NASA CR-179487.

A 2.5 kW advanced technology ion thruster

NASA Technical Reports Server (NTRS)

Poeschel, R. L.

1974-01-01

A program has been conducted in order to improve the performance characteristics of 30 cm thrusters. This program was divided into three distinct, but related tasks: (1) the discharge chamber and component design modifications proposed for inclusion in the engineering model thruster were evaluated and engineering specifications were verified; (2) thrust losses which result from the contributions of double charged ions and nonaxial ion trajectories to the ion beam current were measured and (3) the specification and verification of power processor and control requirements of the engineering model thruster design were demonstrated. Proven design modifications which provide improved efficiencies are incorporated into the engineering model thruster during a structural re-design without introducing additional delay in schedule or new risks. In addition, a considerable amount of data is generated on the relation of double ion production and beam divergence to thruster parameters. Overall thruster efficiency is increased from 68% to 71% at full power, including corrections for double ion and beam divergence thrust losses.

Flight-determined benefits of integrated flight-propulsion control systems

NASA Technical Reports Server (NTRS)

Stewart, James F.; Burcham, Frank W., Jr.; Gatlin, Donald H.

1992-01-01

Over the last two decades, NASA has conducted several experiments in integrated flight-propulsion control. Benefits have included improved maneuverability; increased thrust, range, and survivability; reduced fuel consumption; and reduced maintenance. This paper presents the basic concepts for control integration, examples of implementation, and benefits. The F-111E experiment integrated the engine and inlet control systems. The YF-12C incorporated an integral control system involving the inlet, autopilot, autothrottle, airdata, navigation, and stability augmentation systems. The F-15 research involved integration of the engine, flight, and inlet control systems. Further extension of the integration included real-time, onboard optimization of engine, inlet, and flight control variables; a self-repairing flight control system; and an engines-only control concept for emergency control. The F-18A aircraft incorporated thrust vectoring integrated with the flight control system to provide enhanced maneuvering at high angles of attack. The flight research programs and the resulting benefits of each program are described.

Thrust imbalance of the Space Shuttle solid rocket motors

NASA Technical Reports Server (NTRS)

Foster, W. A., Jr.; Sforzini, R. H.; Shackelford, B. W., Jr.

1981-01-01

The Monte Carlo statistical analysis of thrust imbalance is applied to both the Titan IIIC and the Space Shuttle solid rocket motors (SRMs) firing in parallel, and results are compared with those obtained from the Space Shuttle program. The test results are examined in three phases: (1) pairs of SRMs selected from static tests of the four developmental motors (DMs 1 through 4); (2) pairs of SRMs selected from static tests of the three quality assurance motors (QMs 1 through 3); (3) SRMs on the first flight test vehicle (STS-1A and STS-1B). The simplified internal ballistic model utilized for computing thrust from head-end pressure measurements on flight tests is shown to agree closely with measured thrust data. Inaccuracies in thrust imbalance evaluation are explained by possible flight test instrumentation errors.

Architecture-Led Safety Process

DTIC Science & Technology

2016-12-01

Action Hazard Guide 42 Table 18: Comparative Table of Safety and Reliability Terms 47 CMU/SEI-2016-TR-012 | SOFTWARE ENGINEERING INSTITUTE...provides too much thrust Engine is slow to pro- vide commanded thrust (increase or de- crease) Engine will not shut- down when com - manded...Thrust level must be provided at the com - manded level H4: Engine is slow to provide commanded thrust SC3: Engine must provide commanded thrust in

NASA Office of Aeronautics and Space Technology Summer Workshop. Volume 1: Data processing and transfer panel

NASA Technical Reports Server (NTRS)

1975-01-01

The data processing and transfer technology areas that need to be developed and that could benefit from space flight experiments are identified. Factors considered include: user requirements, concepts in 'Outlook for Space', and cost reduction. Major program thrusts formulated are an increase in end-to-end information handling and a reduction in life cycle costs.

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.

Episodic growth of fold-thrust belts: Insights from Finite Element Modelling

NASA Astrophysics Data System (ADS)

Yang, Xiaodong; Peel, Frank J.; Sanderson, David J.; McNeill, Lisa C.

2017-09-01

The sequential development of a fold-thrust belt was investigated using 2D Finite Element Modelling (FEM). The new model results show that a thrust system is typically composed of three distinct regions: the thrust wedge, pre-wedge, and undeformed region. The thrust wedge involves growth that repeats episodically and cyclically. A cycle of wedge building starts as frontal accretion occurs, which is accompanied by a rapid increase in wedge width reducing the taper angle below critical. In response to this, the wedge interior (tracked here by the 50 m displacement position) rapidly propagates forwards into a region of incipient folding. The taper angle progressively increases until it obtains a constant apparent critical value (∼10°). During this period, the wedge experiences significant shortening after a new thrust initiates at the failure front, leading to a decrease in wedge width. Successive widening of the wedge and subsequent shortening and thrusting maintain a reasonably constant taper angle. The fold-thrust belt evolves cyclically, through a combination of rapid advancement of the wedge and subsequent gradual, slow wedge growth. The new model results also highlights that there is clear, although minor, deformation (0-10 m horizontal displacement) in front of the thrust wedge.

Target thrust measurement for applied-field magnetoplasmadynamic thruster

NASA Astrophysics Data System (ADS)

Wang, B.; Yang, W.; Tang, H.; Li, Z.; Kitaeva, A.; Chen, Z.; Cao, J.; Herdrich, G.; Zhang, K.

2018-07-01

In this paper, we present a flat target thrust stand which is designed to measure the thrust of a steady-state applied-field magnetoplasmadynamic thruster (AF-MPDT). In our experiments we varied target-thruster distances and target size to analyze their influence on the target thrust measurement results. The obtained thrust-distance curves increase to local maximum and then decreases with the increasing distance, which means that the plume of the AF-MPDT can still accelerate outside the thruster exit. The peak positions are related to the target sizes: larger targets can make the peak positions further from the thruster and decrease the measurement errors. To further improve the reliability of measurement results, a thermal equilibrium assumption combined with Knudsen’s cosine law is adapted to analyze the error caused by the back stream of plume particles. Under the assumption, the error caused by particle backflow is no more than 3.6% and the largest difference between the measured thrust and the theoretical thrust is 14%. Moreover, it was verified that target thrust measurement can disturb the working of the AF-MPD thruster, and the influence on the thrust measurement result is no more than 1% in our experiment.

Development of HIDEC adaptive engine control systems

NASA Technical Reports Server (NTRS)

Landy, R. J.; Yonke, W. A.; Stewart, J. F.

1986-01-01

The purpose of NASA's Highly Integrated Digital Electronic Control (HIDEC) flight research program is the development of integrated flight propulsion control modes, and the evaluation of their benefits aboard an F-15 test aircraft. HIDEC program phases are discussed, with attention to the Adaptive Engine Control System (ADECS I); this involves the upgrading of PW1128 engines for operation at higher engine pressure ratios and the production of greater thrust. ADECS II will involve the development of a constant thrust mode which will significantly reduce turbine operating temperatures.

Modified RS2101 rocket engine study program

NASA Technical Reports Server (NTRS)

1971-01-01

The purpose of the program is to perform design studies and analyses to determine the effects of incorporating a 60:1 expansion area ratio nozzle extension, extended firing time, and modified operating conditions and environments on the MM'71 rocket engine assembly. An injector-to-thrust chamber seal study was conducted to define potential solutions for leakage past this joint. The results and recommendations evolving from the engine thermal analyses, the injector-to-thrust chamber seal studies, and the nozzle extension joint stress analyses are presented.

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.

NASA's Chemical Transfer Propulsion Program for Pathfinder

NASA Technical Reports Server (NTRS)

Hannum, Ned P.; Berkopec, Frank D.; Zurawski, Robert L.

1989-01-01

Pathfinder is a research and technology project, with specific deliverables, initiated by the National Aeronautics and Space Administration (NASA) which will strengthen the technology base of the United States civil space program in preparation for future space exploration missions. Pathfinder begins in Fiscal Year 1989, and is to advance a collection of critical technologies for these missions and ensure technology readiness for future national decisions regarding exploration of the solar system. The four major thrusts of Pathfinder are: surface exploration, in-space operations, humans-in-space, and space transfer. The space transfer thrust will provide the critical technologies needed for transportation to, and return from, the Moon, Mars, and other planets in the solar system, as well as for reliable and cost-effective Earth-orbit operations. A key element of this thrust is the Chemical Transfer Propulsion program which will provide the propulsion technology for high performance, liquid oxygen/liquid hydrogen expander cycle engines which may be operated and maintained in space. Described here are the program overview including the goals and objectives, management, technical plan, and technology transfer for the Chemical Transfer Propulsion element of Pathfinder.

Propulsion system-flight control integration-flight evaluation and technology transition

NASA Technical Reports Server (NTRS)

Burcham, Frank W., Jr.; Gilyard, Glenn B.; Myers, Lawrence P.

1990-01-01

Integration of propulsion and flight control systems and their optimization offering significant performance improvement are assessed. In particular, research programs conducted by NASA on flight control systems and propulsion system-flight control interactions on the YF-12 and F-15 aircraft are addressed; these programs have demonstrated increased thrust, reduced fuel consumption, increased engine life, and improved aircraft performance. Focus is placed on altitude control, speed-Mach control, integrated controller design, as well as flight control systems and digital electronic engine control. A highly integrated digital electronic control program is analyzed and compared with a performance seeking control program. It is shown that the flight evaluation and demonstration of these technologies have been a key part in the transition of the concepts to production and operational use on a timely basis.

Weight and cost estimating relationships for heavy lift airships

NASA Technical Reports Server (NTRS)

Gray, D. W.

1979-01-01

Weight and cost estimating relationships, including additional parameters that influence the cost and performance of heavy-lift airships (HLA), are discussed. Inputs to a closed loop computer program, consisting of useful load, forward speed, lift module positive or negative thrust, and rotors and propellers, are examined. Detail is given to the HLA cost and weight program (HLACW), which computes component weights, vehicle size, buoyancy lift, rotor and propellar thrust, and engine horse power. This program solves the problem of interrelating the different aerostat, rotors, engines and propeller sizes. Six sets of 'default parameters' are left for the operator to change during each computer run enabling slight data manipulation without altering the program.

Romanian MRE Rocket Engines Program - An Early Endeavor

NASA Astrophysics Data System (ADS)

Rugescu, R. E.

2002-01-01

(MRE) was initiated in the years '60 of the past century at the Chair of Aerospace Sciences "Elie Carafoli" from the "Politehnica" University in Bucharest (PUB). Consisting of theoretical and experimental investigations in the form of computational methods and technological solutions for small size MRE-s and the concept of the test stand for these engines, the program ended in the construction of the first Romanian liquid rocket motors. Hermann Oberth and Dorin Pavel, were known from 1923, no experimental practice was yet tempted, at the time level of 1960. It was the intention of the developers at PUB to cover this gap and initiate a feasible, low-cost, demonstrative program of designing and testing experimental models of MRE. The research program was oriented towards future development of small size space carrier vehicles for scientific applications only, as an independent program with no connection to other defense programs imagined by the authorities in Bucharest, at that time. Consequently the entire financial support was assured by "Politehnica" university. computerized methods in the thermochemistry of heterogeneous combustion, for both steady and unsteady flows with chemical reactions and two phase flows. The research was gradually extended to the production of a professional CAD program for steady-state heat transfer simulations and the loading capacity analyses of the double wall, cooled thrust chamber. The resulting computer codes were run on a 360-30 IMB machine, beginning in 1968. Some of the computational methods were first exposed at the 9th International Conference on Applied Mechanics, held in Bucharest between June 23-27, 1969. hot testing of a series of storable propellant, variable thrust, variable geometry, liquid rocket motors, with a maximal thrust of 200N. A remotely controlled, portable test bad, actuated either automatically or manually and consisting of a 6-modules construction was built for this motor series, with a simple 8 analog-channel and 5 digital-channel data measuring and recording system. The first hot test firing of the MRE-1B motor took place successfully on April 9th, 1969 in Bucharest, at the "Elie Carafoli" Chair of UPB. The research program continued with the development of a series of solid, double base propellant rocket and ram-rocket motors, with emphasize on the optimization of the gasdynamic contour of the engine, in order to increase the flight performances. Increments of up to 8% in specific thrust were measured on the test stand, with mass savings and no extra costs. The test firing of the first Romanian, air-breathing ram-rocket engine took place successfully in august 1987 at the Chemical Works in Fagaras, Romania. Astronautics", founded in Bucharest. The principles and history of the "MRE" research program are presented in the proposed paper.

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.

An adaptive guidance algorithm for an aerodynamically assisted orbital plane change maneuver. M.S. Thesis

NASA Technical Reports Server (NTRS)

Blissit, J. A.

1986-01-01

Using analysis results from the post trajectory optimization program, an adaptive guidance algorithm is developed to compensate for density, aerodynamic and thrust perturbations during an atmospheric orbital plane change maneuver. The maneuver offers increased mission flexibility along with potential fuel savings for future reentry vehicles. Although designed to guide a proposed NASA Entry Research Vehicle, the algorithm is sufficiently generic for a range of future entry vehicles. The plane change analysis provides insight suggesting a straight-forward algorithm based on an optimized nominal command profile. Bank angle, angle of attack, and engine thrust level, ignition and cutoff times are modulated to adjust the vehicle's trajectory to achieve the desired end-conditions. A performance evaluation of the scheme demonstrates a capability to guide to within 0.05 degrees of the desired plane change and five nautical miles of the desired apogee altitude while maintaining heating constraints. The algorithm is tested under off-nominal conditions of + or -30% density biases, two density profile models, + or -15% aerodynamic uncertainty, and a 33% thrust loss and for various combinations of these conditions.

Propulsion system-flight control integration and optimization: Flight evaluation and technology transition

NASA Technical Reports Server (NTRS)

Burcham, Frank W., Jr.; Gilyard, Glenn B.; Myers, Lawrence P.

1990-01-01

Integration of propulsion and flight control systems and their optimization offers significant performance improvements. Research programs were conducted which have developed new propulsion and flight control integration concepts, implemented designs on high-performance airplanes, demonstrated these designs in flight, and measured the performance improvements. These programs, first on the YF-12 airplane, and later on the F-15, demonstrated increased thrust, reduced fuel consumption, increased engine life, and improved airplane performance; with improvements in the 5 to 10 percent range achieved with integration and with no changes to hardware. The design, software and hardware developments, and testing requirements were shown to be practical.

Dynamic Positioning Capability Analysis for Marine Vessels Based on A DPCap Polar Plot Program

NASA Astrophysics Data System (ADS)

Wang, Lei; Yang, Jian-min; Xu, Sheng-wen

2018-03-01

Dynamic positioning capability (DPCap) analysis is essential in the selection of thrusters, in their configuration, and during preliminary investigation of the positioning ability of a newly designed vessel dynamic positioning system. DPCap analysis can help determine the maximum environmental forces, in which the DP system can counteract in given headings. The accuracy of the DPCap analysis is determined by the precise estimation of the environmental forces as well as the effectiveness of the thrust allocation logic. This paper is dedicated to developing an effective and efficient software program for the DPCap analysis for marine vessels. Estimation of the environmental forces can be obtained by model tests, hydrodynamic computation and empirical formulas. A quadratic programming method is adopted to allocate the total thrust on every thruster of the vessel. A detailed description of the thrust allocation logic of the software program is given. The effectiveness of the new program DPCap Polar Plot (DPCPP) was validated by a DPCap analysis for a supply vessel. The present study indicates that the developed program can be used in the DPCap analysis for marine vessels. Moreover, DPCap analysis considering the thruster failure mode might give guidance to the designers of vessels whose thrusters need to be safer.

Small Fast Spectrum Reactor Designs Suitable for Direct Nuclear Thermal Propulsion

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

Bruce G. Schnitzler; Stanley K. Borowski

Advancement of U.S. scientific, security, and economic interests through a robust space exploration program requires high performance propulsion systems to support a variety of robotic and crewed missions beyond low Earth orbit. Past studies, in particular those in support of both the Strategic Defense Initiative (SDI) and Space Exploration Initiative (SEI), have shown nuclear thermal propulsion systems provide superior performance for high mass high propulsive delta-V missions. The recent NASA Design Reference Architecture (DRA) 5.0 Study re-examined mission, payload, and transportation system requirements for a human Mars landing mission in the post-2030 timeframe. Nuclear thermal propulsion was again identified asmore » the preferred in-space transportation system. A common nuclear thermal propulsion stage with three 25,000-lbf thrust engines was used for all primary mission maneuvers. Moderately lower thrust engines may also have important roles. In particular, lower thrust engine designs demonstrating the critical technologies that are directly extensible to other thrust levels are attractive from a ground testing perspective. An extensive nuclear thermal rocket technology development effort was conducted from 1955-1973 under the Rover/NERVA Program. Both graphite and refractory metal alloy fuel types were pursued. Reactors and engines employing graphite based fuels were designed, built and ground tested. A number of fast spectrum reactor and engine designs employing refractory metal alloy fuel types were proposed and designed, but none were built. The Small Nuclear Rocket Engine (SNRE) was the last engine design studied by the Los Alamos National Laboratory during the program. At the time, this engine was a state-of-the-art graphite based fuel design incorporating lessons learned from the very successful technology development program. The SNRE was a nominal 16,000-lbf thrust engine originally intended for unmanned applications with relatively short engine operations and the engine and stage design were constrained to fit within the payload volume of the then planned space shuttle. The SNRE core design utilized hexagonal fuel elements and hexagonal structural support elements. The total number of elements can be varied to achieve engine designs of higher or lower thrust levels. Some variation in the ratio of fuel elements to structural elements is also possible. Options for SNRE-based engine designs in the 25,000-lbf thrust range were described in a recent (2010) Joint Propulsion Conference paper. The reported designs met or exceeded the performance characteristics baselined in the DRA 5.0 Study. Lower thrust SNRE-based designs were also described in a recent (2011) Joint Propulsion Conference paper. Recent activities have included parallel evaluation and design efforts on fast spectrum engines employing refractory metal alloy fuels. These efforts include evaluation of both heritage designs from the Argonne National Laboratory (ANL) and General Electric Company GE-710 Programs as well as more recent designs. Results are presented for a number of not-yet optimized fast spectrum engine options.« less

Small Fast Spectrum Reactor Designs Suitable for Direct Nuclear Thermal Propulsion

NASA Technical Reports Server (NTRS)

Schnitzler, Bruce G.; Borowski, Stanley K.

2012-01-01

Advancement of U.S. scientific, security, and economic interests through a robust space exploration program requires high performance propulsion systems to support a variety of robotic and crewed missions beyond low Earth orbit. Past studies, in particular those in support of the Space Exploration Initiative (SEI), have shown nuclear thermal propulsion systems provide superior performance for high mass high propulsive delta-V missions. The recent NASA Design Reference Architecture (DRA) 5.0 Study re-examined mission, payload, and transportation system requirements for a human Mars landing mission in the post-2030 timeframe. Nuclear thermal propulsion was again identified as the preferred in-space transportation system. A common nuclear thermal propulsion stage with three 25,000-lbf thrust engines was used for all primary mission maneuvers. Moderately lower thrust engines may also have important roles. In particular, lower thrust engine designs demonstrating the critical technologies that are directly extensible to other thrust levels are attractive from a ground testing perspective. An extensive nuclear thermal rocket technology development effort was conducted from 1955-1973 under the Rover/NERVA Program. Both graphite and refractory metal alloy fuel types were pursued. Reactors and engines employing graphite based fuels were designed, built and ground tested. A number of fast spectrum reactor and engine designs employing refractory metal alloy fuel types were proposed and designed, but none were built. The Small Nuclear Rocket Engine (SNRE) was the last engine design studied by the Los Alamos National Laboratory during the program. At the time, this engine was a state-of-the-art graphite based fuel design incorporating lessons learned from the very successful technology development program. The SNRE was a nominal 16,000-lbf thrust engine originally intended for unmanned applications with relatively short engine operations and the engine and stage design were constrained to fit within the payload volume of the then planned space shuttle. The SNRE core design utilized hexagonal fuel elements and hexagonal structural support elements. The total number of elements can be varied to achieve engine designs of higher or lower thrust levels. Some variation in the ratio of fuel elements to structural elements is also possible. Options for SNRE-based engine designs in the 25,000-lbf thrust range were described in a recent (2010) Joint Propulsion Conference paper. The reported designs met or exceeded the performance characteristics baselined in the DRA 5.0 Study. Lower thrust SNRE-based designs were also described in a recent (2011) Joint Propulsion Conference paper. Recent activities have included parallel evaluation and design efforts on fast spectrum engines employing refractory metal alloy fuels. These efforts include evaluation of both heritage designs from the Argonne National Laboratory (ANL) and General Electric Company GE-710 Programs as well as more recent designs. Results are presented for a number of not-yet optimized fast spectrum engine options.

Orbit Transfer Vehicle (OTV) engine phase A study, extension 1. Volume 3: Study cost estimates

NASA Technical Reports Server (NTRS)

Christensen, K. L.

1980-01-01

Program cost and planning data based on 1980 technology and shown in 1979 dollars for a 20K lb Thrust Staged Combustion Cycle Engine are presented. These data were compared with those for the Advanced Expander Cycle Engine at 10K lb and 20K lb thrust levels.

Estimating Thruster Impulses From IMU and Doppler Data

NASA Technical Reports Server (NTRS)

Lisano, Michael E.; Kruizinga, Gerhard L.

2009-01-01

A computer program implements a thrust impulse measurement (TIM) filter, which processes data on changes in velocity and attitude of a spacecraft to estimate the small impulsive forces and torques exerted by the thrusters of the spacecraft reaction control system (RCS). The velocity-change data are obtained from line-of-sight-velocity data from Doppler measurements made from the Earth. The attitude-change data are the telemetered from an inertial measurement unit (IMU) aboard the spacecraft. The TIM filter estimates the threeaxis thrust vector for each RCS thruster, thereby enabling reduction of cumulative navigation error attributable to inaccurate prediction of thrust vectors. The filter has been augmented with a simple mathematical model to compensate for large temperature fluctuations in the spacecraft thruster catalyst bed in order to estimate thrust more accurately at deadbanding cold-firing levels. Also, rigorous consider-covariance estimation is applied in the TIM to account for the expected uncertainty in the moment of inertia and the location of the center of gravity of the spacecraft. The TIM filter was built with, and depends upon, a sigma-point consider-filter algorithm implemented in a Python-language computer program.

Effect of blade outlet angle on radial thrust of single-blade centrifugal pump

NASA Astrophysics Data System (ADS)

Nishi, Y.; Fukutomi, J.; Fujiwara, R.

2012-11-01

Single-blade centrifugal pumps are widely used as sewage pumps. However, a large radial thrust acts on a single blade during pump operation because of the geometrical axial asymmetry of the impeller. This radial thrust causes vibrations of the pump shaft, reducing the service life of bearings and shaft seal devices. Therefore, to ensure pump reliability, it is necessary to quantitatively understand the radial thrust and clarify the behavior and generation mechanism. This study investigated the radial thrust acting on two kinds of single-blade centrifugal impellers having different blade outlet angles by experiments and computational fluid dynamics (CFD) analysis. Furthermore, the radial thrust was modeled by a combination of three components, inertia, momentum, and pressure, by applying an unsteady conservation of momentum to this impeller. As a result, the effects of the blade outlet angle on both the radial thrust and the modeled components were clarified. The total head of the impeller with a blade outlet angle of 16 degrees increases more than the impeller with a blade outlet angle of 8 degrees at a large flow rate. In this case, since the static pressure of the circumference of the impeller increases uniformly, the time-averaged value of the radial thrust of both impellers does not change at every flow rate. On the other hand, since the impeller blade loading becomes large, the fluctuation component of the radial thrust of the impeller with the blade outlet angle of 16 degrees increases. If the blade outlet angle increases, the fluctuation component of the inertia component will increase, but the time-averaged value of the inertia component is located near the origin despite changes in the flow rate. The fluctuation component of the momentum component becomes large at all flow rates. Furthermore, although the time-averaged value of the pressure component is almost constant, the fluctuation component of the pressure component becomes large at a large flow rate. In addition to the increase of the fluctuation component of this pressure component, because the fluctuation component of the inertia and momentum components becomes large (as mentioned above), the radial thrust increases at a large flow rate, as is the case for the impeller with a large blade outlet angle.

High-power, null-type, inverted pendulum thrust stand.

PubMed

Xu, Kunning G; Walker, Mitchell L R

2009-05-01

This article presents the theory and operation of a null-type, inverted pendulum thrust stand. The thrust stand design supports thrusters having a total mass up to 250 kg and measures thrust over a range of 1 mN to 5 N. The design uses a conventional inverted pendulum to increase sensitivity, coupled with a null-type feature to eliminate thrust alignment error due to deflection of thrust. The thrust stand position serves as the input to the null-circuit feedback control system and the output is the current to an electromagnetic actuator. Mechanical oscillations are actively damped with an electromagnetic damper. A closed-loop inclination system levels the stand while an active cooling system minimizes thermal effects. The thrust stand incorporates an in situ calibration rig. The thrust of a 3.4 kW Hall thruster is measured for thrust levels up to 230 mN. The uncertainty of the thrust measurements in this experiment is +/-0.6%, determined by examination of the hysteresis, drift of the zero offset and calibration slope variation.

Highly integrated digital engine control system on an F-15 airplane

NASA Technical Reports Server (NTRS)

Burcham, F. W., Jr.; Haering, E. A., Jr.

1984-01-01

The Highly Integrated Digital Electronic Control (HIDEC) program will demonstrate and evaluate the improvements in performance and mission effectiveness that result from integrated engine/airframe control systems. This system is being used on the F-15 airplane. An integrated flightpath management mode and an integrated adaptive engine stall margin mode are implemented into the system. The adaptive stall margin mode is a highly integrated mode in which the airplane flight conditions, the resulting inlet distortion, and the engine stall margin are continuously computed; the excess stall margin is used to uptrim the engine for more thrust. The integrated flightpath management mode optimizes the flightpath and throttle setting to reach a desired flight condition. The increase in thrust and the improvement in airplane performance is discussed.

Test and evaluation of the HIDEC engine uptrim algorithm. [Highly Integrated Digital Electronic Control for aircraft

NASA Technical Reports Server (NTRS)

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

1986-01-01

The highly integrated digital electronic control (HIDEC) program will demonstrate and evaluate the improvements in performance and mission effectiveness that result from integrated engine-airframe control systems. Performance improvements will result from an adaptive engine stall margin mode, a highly integrated mode that uses the airplane flight conditions and the resulting inlet distortion to continuously compute engine stall margin. When there is excessive stall margin, the engine is uptrimmed for more thrust by increasing engine pressure ratio (EPR). The EPR uptrim logic has been evaluated and implemente into computer simulations. Thrust improvements over 10 percent are predicted for subsonic flight conditions. The EPR uptrim was successfully demonstrated during engine ground tests. Test results verify model predictions at the conditions tested.

Dryden/Edwards 1994 Thrust-Vectoring Aircraft Fleet - F-18 HARV, X-31, F-16 MATV

NASA Technical Reports Server (NTRS)

1994-01-01

The three thrust-vectoring aircraft at Edwards, California, each capable of flying at extreme angles of attack, cruise over the California desert in formation during flight in March 1994. They are, from left, NASA's F-18 High Alpha Research Vehicle (HARV), flown by the NASA Dryden Flight Research Center; the X-31, flown by the X-31 International Test Organization (ITO) at Dryden; and the Air Force F-16 Multi-Axis Thrust Vectoring (MATV) aircraft. All three aircraft were flown in different programs and were developed independently. The NASA F-18 HARV was a testbed to produce aerodynamic data at high angles of attack to validate computer codes and wind tunnel research. The X-31 was used to study thrust vectoring to enhance close-in air combat maneuvering, while the F-16 MATV was a demonstration of how thrust vectoring could be applied to operational aircraft.

Pointing and control system enabling technology for future automated space missions

NASA Technical Reports Server (NTRS)

Dahlgren, J. B.

1978-01-01

Future automated space missions present challenging opportunities in the pointing-and-control technology disciplines. The enabling pointing-and-control system technologies for missions from 1985 to the year 2000 were identified and assessed. A generic mission set including Earth orbiter, planetary, and other missions which predominantly drive the pointing-and-control requirements was selected for detailed evaluation. Technology candidates identified were prioritized as planning options for future NASA-OAST advanced development programs. The primary technology thrusts in each candidate program were cited, and advanced development programs in pointing-and-control were recommended for the FY 80 to FY 87 period, based on these technology thrusts.

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.

Management of Guidance, Navigation and Control Technologies for Spacecraft Formations under the NASA Cross-Enterprise Technology Development Program (CETDP)

NASA Technical Reports Server (NTRS)

Hartman, Kathy; Weidow, David; Hadaegh, Fred

1999-01-01

Breakthrough technology development is critical to securing the future of our space industry. The National Aeronautics and Space Administration (NASA) Cross-Enterprise Technology Development Program (CETDP) is developing critical space technologies that enable innovative and less costly missions, and spawn new mission opportunities through revolutionary, long-term, high-risk, high-payoff technology advances. The CETDP is a NASA-wide activity managed by the Advanced Technology and Mission Studies Division (AT&MS) at Headquarters Office of Space Science. Program management for CETDP is distributed across the multiple NASA Centers and draws on expertise throughout the Agency. The technology research activities are organized along Project-level divisions called thrust areas that are directly linked to the Agency's goals and objectives of the Enterprises: Earth Science, Space Science, Human Exploration and Development of Space; and the Office of the Chief Technologist's (OCT) strategic technology areas. Cross-Enterprise technology is defined as long-range strategic technologies that have broad potential to span the needs of more than one Enterprise. Technology needs are identified and prioritized by each of the primary customers. The thrust area manager (TAM) for each division is responsible for the ultimate success of technologies within their area, and can draw from industry, academia, other government agencies, other CETDP thrust areas, and other NASA Centers to accomplish the goals of the thrust area. An overview of the CETDP and description of the future directions of the thrust area called Distributed Spacecraft are presented in this paper. Revolutionary technologies developed within this thrust area will enable the implementation of a spatially distributed network of individual vehicles, or assets, collaborating as a single collective unit, and exhibiting a common system-wide capability to accomplish a shared objective. With such a capability, new Earth and space science measurement concepts become a reality.

Management of Guidance, Navigation, and Control Technologies for Spacecraft Formations Under the NASA Cross Enterprise Technology Development Program (CETDP)

NASA Technical Reports Server (NTRS)

Hartman, Kathy; Weidow, David; Hadaegh, Fred

1999-01-01

Breakthrough technology development is critical to securing the future of our space industry. The National Aeronautics and Space Administration (NASA) Cross-Enterprise Technology Development Program (CETDP) is developing critical space technologies that enable innovative and less costly missions, and spawn new mission opportunities through revolutionary, long-term, high-risk, high-payoff technology advances. The CETDP is a NASA-wide activity managed by the Advanced Technology and Mission Studies Division (AT&MS) at Headquarters Office of Space Science. Program management for CETDP is distributed across the multiple NASA Centers and draws on expertise throughout the Agency. The technology research activities are organized along Project-level divisions called thrust areas that are directly linked to the Agency's goals and objectives of the Enterprises: Earth Science, Space Science, Human Exploration and Development of Space; and the Office of the Chief Technologist's (OCT) strategic technology areas. Cross-Enterprise technology is defined as long-range strategic technologies that have broad potential to span the needs of more than one Enterprise. Technology needs are identified and prioritized by each of the primary customers. The thrust area manager (TAM) for each division is responsible for the ultimate success of technologies within their area, and can draw from industry, academia, other government agencies, other CETDP thrust areas, and other NASA Centers to accomplish the goals of the thrust area. An overview of the CETDP and description of the future directions of the thrust area called Distributed Spacecraft are presented in this paper. Revolutionary technologies developed within this thrust area will enable the implementation of a spatially distributed network of individual vehicles, or assets, collaborating as a single collective unit, and exhibiting a common system-wide capability to accomplish a shared objective. With such a capability, new Earth and space science measurement concepts become a reality.

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.

Hamilton Standard Q-fan demonstrator dynamic pitch change test program, volume 1

NASA Technical Reports Server (NTRS)

Demers, W. J.; Nelson, D. J.; Wainauski, H. S.

1975-01-01

Tests of a full scale variable pitch fan engine to obtain data on the structural characteristics, response times, and fan/core engine compatibility during transient changes in blade angle, fan rpm, and engine power is reported. Steady state reverse thrust tests with a take off nozzle configuration were also conducted. The 1.4 meter diameter, 13 bladed controllable pitch fan was driven by a T55 L 11A engine with power and blade angle coordinated by a digital computer. The tests demonstrated an ability to change from full forward thrust to reverse thrust in less than one (1) second. Reverse thrust was effected through feather and through flat pitch; structural characteristics and engine/fan compatibility were within satisfactory limits.

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.

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.

Heliocentric interplanetary low thrust trajectory optimization program, supplement 1, part 2

NASA Technical Reports Server (NTRS)

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

1978-01-01

The improvements made to the HILTOP electric propulsion trajectory computer program are described. A more realistic propulsion system model was implemented in which various thrust subsystem efficiencies and specific impulse are modeled as variable functions of power available to the propulsion system. The number of operating thrusters are staged, and the beam voltage is selected from a set of five (or less) constant voltages, based upon the application of variational calculus. The constant beam voltages may be optimized individually or collectively. The propulsion system logic is activated by a single program input key in such a manner as to preserve the HILTOP logic. An analysis describing these features, a complete description of program input quantities, and sample cases of computer output illustrating the program capabilities are presented.

The Rocket Engine Advancement Program 2 (REAP2)

NASA Technical Reports Server (NTRS)

Harper, Brent (Technical Monitor); Hawk, Clark W.

2004-01-01

The Rocket Engine Advancement Program (REAP) 2 program is being conducted by a university propulsion consortium consisting of the University of Alabama in Huntsville, Penn State University, Purdue University, Tuskegee University and Auburn University. It has been created to bring their combined skills to bear on liquid rocket combustion stability and thrust chamber cooling. The research team involves well established and known researchers in the propulsion community. The cure team provides the knowledge base, research skills, and commitment to achieve an immediate and continuing impact on present and future propulsion issues. through integrated research teams composed of analysts, diagnosticians, and experimentalists working together in an integrated multi-disciplinary program. This paper provides an overview of the program, its objectives and technical approaches. Research on combustion instability and thrust chamber cooling are being accomplished

14 CFR Appendix I to Part 25 - Installation of an Automatic Takeoff Thrust Control System (ATTCS)

Code of Federal Regulations, 2011 CFR

2011-01-01

... Appendix I to Part 25—Installation of an Automatic Takeoff Thrust Control System (ATTCS) I25.1General. (a... crew to increase thrust or power. I25.2Definitions. (a) Automatic Takeoff Thrust Control System (ATTCS... Control System (ATTCS) I Appendix I to Part 25 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION...

Background and principles of throttles-only flight control

NASA Technical Reports Server (NTRS)

Burcham, Frank W., Jr.

1995-01-01

There have been many cases in which the crew of a multi-engine airplane had to use engine thrust for emergency flight control. Such a procedure is very difficult, because the propulsive control forces are small, the engine response is slow, and airplane dynamics such as the phugoid and dutch roll are difficult to damp with thrust. In general, thrust increases are used to climb, thrust decreases to descend, and differential thrust is used to turn. Average speed is not significantly affected by changes in throttle setting. Pitch control is achieved because of pitching moments due to speed changes, from thrust offset, and from the vertical component of thrust. Roll control is achieved by using differential thrust to develop yaw, which, through the normal dihedral effect, causes a roll. Control power in pitch and roll tends to increase as speed decreases. Although speed is not controlled by the throttles, configuration changes are often available (lowering gear, flaps, moving center-of-gravity) to change the speed. The airplane basic stability is also a significant factor. Fuel slosh and gyroscopic moments are small influences on throttles-only control. The background and principles of throttles-only flight control are described.

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.

ASCENT Program

NASA Technical Reports Server (NTRS)

Brown, Richard; Collier, Gary; Heckenlaible, Richard; Dougherty, Edward; Dolenz, James; Ross, Iain

2012-01-01

The ASCENT program solves the three-dimensional motion and attendant structural loading on a flexible vehicle incorporating, optionally, an active analog thrust control system, aerodynamic effects, and staging of multiple bodies. ASCENT solves the technical problems of loads, accelerations, and displacements of a flexible vehicle; staging of the upper stage from the lower stage; effects of thrust oscillations on the vehicle; a payload's relative motion; the effect of fluid sloshing on vehicle; and the effect of winds and gusts on the vehicle (on the ground or aloft) in a continuous analysis. The ATTACH ASCENT Loads program reads output from the ASCENT flexible body loads program, and calculates the approximate load indicators for the time interval under consideration. It calculates the load indicator values from pre-launch to the end of the first stage.

Human factors technology for America's space program

NASA Technical Reports Server (NTRS)

Montemerlo, M. D.

1982-01-01

NASA is initiating a space human factors research and technology development program in October 1982. The impetus for this program stems from: the frequent and economical access to space provided by the Shuttle, the advances in control and display hardware/software made possible through the recent explosion in microelectronics technology, heightened interest in a space station, heightened interest by the military in space operations, and the fact that the technology for long duration stay times for man in space has received relatively little attention since the Apollo and Skylab missions. The rationale for and issues in the five thrusts of the new program are described. The main thrusts are: basic methodology, crew station design, ground control/operations, teleoperations and extra vehicular activity.

Beam efflux measurements

NASA Technical Reports Server (NTRS)

Komatsu, G. K.; Stellen, J. M., Jr.

1976-01-01

Measurements have been made of the high energy thrust ions, (Group I), high angle/high energy ions (Group II), and high angle/low energy ions (Group IV) of a mercury electron bombardment thruster in the angular divergence range from 0 deg to greater than 90 deg. The measurements have been made as a function of thrust ion current, propellant utilization efficiency, bombardment discharge voltage, screen and accelerator grid potential (accel-decel ratio) and neutralizer keeper potential. The shape of the Group IV (charge exchange) ion plume has remained essentially fixed within the range of variation of the engine operation parameters. The magnitude of the charge exchange ion flux scales with thrust ion current, for good propellant utilization conditions. For fixed thrust ion current, charge exchange ion flux increases for diminishing propellant utilization efficiency. Facility effects influence experimental accuracies within the range of propellant utilization efficiency used in the experiments. The flux of high angle/high energy Group II ions is significantly diminished by the use of minimum decel voltages on the accelerator grid. A computer model of charge exchange ion production and motion has been developed. The program allows computation of charge exchange ion volume production rate, total production rate, and charge exchange ion trajectories for "genuine" and "facilities effects" particles. In the computed flux deposition patterns, the Group I and Group IV ion plumes exhibit a counter motion.

Technical Excellence: A Requirement for Good Engineering

NASA Technical Reports Server (NTRS)

Gill, Paul S.; Vaughan, William W.

2008-01-01

Technical excellence is a requirement for good engineering. Technical excellence has many different ways of expressing itself within engineering. NASA has initiatives that address the enhancement of the Agency's technical excellence and thrust to maintain the associated high level of performance by the Agency on current programs/projects and as it moves into the Constellation Program and the return to the Moon with plans to visit Mars. This paper addresses some of the key initiatives associated with NASA's technical excellence thrust. Examples are provided to illustrate some results being achieved and plans to enhance these initiatives.

Low thrust spacecraft transfers optimization method with the stepwise control structure in the Earth-Moon system in terms of the L1-L2 transfer

NASA Astrophysics Data System (ADS)

Fain, M. K.; Starinova, O. L.

2016-04-01

The paper outlines the method for determination of the locally optimal stepwise control structure in the problem of the low thrust spacecraft transfer optimization in the Earth-Moon system, including the L1-L2 transfer. The total flight time as an optimization criterion is considered. The optimal control programs were obtained by using the Pontryagin's maximum principle. As a result of optimization, optimal control programs, corresponding trajectories, and minimal total flight times were determined.

Episodic Growth of Fold-Thrust Belts: Insights from Finite Element Modelling

NASA Astrophysics Data System (ADS)

Yang, X.; Peel, F.; Sanderson, D. J.; McNeill, L. C.

2016-12-01

The sequential development of an imbricate thrust system was investigated using a set of 2D FEM models. This study provides new insights on how the style and location of thrust activity changes through cycles of thrust accretion by making refined measurements of the thrust system parameters through time and tracking these parameters through each cycle. In addition to conventional wedge parameters (i.e. surface slope, wedge width and height), the overall taper angle is used to determine how the critical taper angle is reached; a particular focus is on the region of outboard minor horizontal displacement provides insights into the forward propagation of material within, and in front of, the thrust wedge; tracking the position of the failure front (where the frontal thrust roots into the basal detachment) reveals the sequence and advancement of the imbricate thrusts. The model results show that a thrust system is generally composed of three deformation components: thrust wedge, pre-wedge and wedge front. A thrust belt involves growth that repeats episodically and cyclically. When a wedge reaches critical taper ( 10°), thrust movement within the wedge slows while the taper angle and wedge width gradually increase. In contrast, the displacement front (tracked here by the location of 0 m displacement) rapidly propagates forward along whilst the wedge height is fast growing. During this period, the wedge experiences a significant shortening after a new thrust initiates at the failure front, leading to an obvious decrease in wedge width. As soon as the critical taper is achieved, wedge interior (tracked here by the location of 50 m displacement) accelerates forward reducing the taper angle below critical. This is accompanied by a sudden increase in wedge width, slow advancement of displacement front, and slow uplift of the fold-thrust belt. The rapid movements within and in front of the wedge occur alternately. The model results also show that there is clear, although minor, activity (5-10 m displacement) in front of the thrust wedge, which distinguishes the failure front from the displacement front throughout the fold-thrust belt development. This spatial and temporal relationship may not have been previously recognized in natural systems.

Effects of thrust amplitude and duration of high velocity low amplitude spinal manipulation on lumbar muscle spindle responses to vertebral position and movement

PubMed Central

Cao, Dong-Yuan; Reed, William R.; Long, Cynthia R.; Kawchuk, Gregory N.; Pickar, Joel G.

2013-01-01

OBJECTIVE Mechanical characteristics of high velocity low amplitude spinal manipulations (HVLA-SM) can be variable. Sustained changes in peripheral neuronal signaling due to altered load transmission to a sensory receptor’s local mechanical environment are often considered a mechanism contributing to the therapeutic effects of spinal manipulation. The purpose of this study was to determine whether an HVLA-SM’s thrust amplitude or duration altered neural responsiveness of lumbar muscle spindles to either vertebral movement or position. METHODS Anesthetized cats (n=112) received L6 HVLA-SMs delivered to the spinous process. Cats were divided into 6 cohorts depending upon the peak thrust force (25%, 55%, 85% body weight) or thrust displacement (1, 2, 3mm) they received. Cats in each cohort received 8 thrust durations (0–250ms). Afferent discharge from 112 spindles was recorded in response to ramp and hold vertebral movement before and after the manipulation. Changes in mean instantaneous frequency (MIF) during the baseline period preceding the ramps (ΔMIFresting), during ramp movements (ΔMIFmovement), and with the vertebra held in the new position (ΔMIFposition) were compared. RESULTS Thrust duration had a small but statistically significant effect on ΔMIFresting at all six thrust amplitudes compared to control (0ms thrust duration). The lowest amplitude thrust displacement (1mm) increased ΔMIFresting at all thrust durations. For all the other thrust displacements and forces, the direction of change in ΔMIFresting was not consistent and the pattern of change was not systematically related to thrust duration. Regardless of thrust force, displacement, or duration, ΔMIFmovement and ΔMIFposition were not significantly different from control. Conclusion Relatively low amplitude thrust displacements applied during an HVLA-SM produced sustained increases in the resting discharge of paraspinal muscle spindles regardless of the duration over which the thrust was applied. However, regardless of the HVLA-SM’s thrust amplitude or duration, the responsiveness of paraspinal muscle spindles to vertebral movement and to a new vertebral position was not affected. PMID:23499141

Saturn Apollo Program

NASA Image and Video Library

1964-11-01

This image shows the Saturn V S-IC-T stage (S-IC static test article) fuel tank being attached to the thrust structure in the vehicle assembly building at the Marshall Space Flight Center (MSFC). The S-IC stage utilized five F-1 engines that used liquid oxygen and kerosene as propellant and provided a combined thrust of 7,500,000 pounds.

Saturn Apollo Program

NASA Image and Video Library

1968-01-01

The Saturn 1B S-IB (first) stage being prepared for shipment at Michoud Assembly Facility (MAF), near New Orleans, Louisiana. Developed by the Marshall Space Flight Center and built by the Chrysler Corporation at MAF, the S-IB stage utilized the eight H-1 engines and each produced 200,000 pounds of thrust, a combined thrust of 1,600,000 pounds.

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.

Nondestructive evaluations

NASA Astrophysics Data System (ADS)

Kulkarni, S.

1993-03-01

This report discusses Nondestructive Evaluation (NDE) thrust area which supports initiatives that advance inspection science and technology. The goal of the NDE thrust area is to provide cutting-edge technologies that have promise of inspection tools three to five years in the future. In selecting projects, the thrust area anticipates the needs of existing and future Lawrence Livermore National Laboratory (LLNL) programs. NDE provides materials characterization inspections, finished parts, and complex objects to find flaws and fabrication defects and to determine their physical and chemical characteristics. NDE also encompasses process monitoring and control sensors and the monitoring of in-service damage. For concurrent engineering, NDE becomes a frontline technology and strongly impacts issues of certification and of life prediction and extension. In FY-92, in addition to supporting LLNL programs and the activities of nuclear weapons contractors, NDE has initiated several projects with government agencies and private industries to study aging infrastructures and to advance manufacturing processes. Examples of these projects are (1) the Aging Airplanes Inspection Program for the Federal Aviation Administration, (2) Signal Processing of Acoustic Signatures of Heart Valves for Shiley, Inc., and (3) Turbine Blade Inspection for the Air Force, jointly with Southwest Research Institute and Garrett. In FY-92, the primary contributions of the NDE thrust area, described in this report, were in fieldable chemical sensor systems, computed tomography, and laser generation and detection of ultrasonic energy.

HIDEC F-15 adaptive engine control system flight test results

NASA Technical Reports Server (NTRS)

Smolka, James W.

1987-01-01

NASA-Ames' Highly Integrated Digital Electronic Control (HIDEC) flight test program aims to develop fully integrated airframe, propulsion, and flight control systems. The HIDEC F-15 adaptive engine control system flight test program has demonstrated that significant performance improvements are obtainable through the retention of stall-free engine operation throughout the aircraft flight and maneuver envelopes. The greatest thrust increase was projected for the medium-to-high altitude flight regime at subsonic speed which is of such importance to air combat. Adaptive engine control systems such as the HIDEC F-15's can be used to upgrade the performance of existing aircraft without resort to expensive reengining programs.

Distinguishing thrust sequences in gravity-driven fold and thrust belts

NASA Astrophysics Data System (ADS)

Alsop, G. I.; Weinberger, R.; Marco, S.

2018-04-01

Piggyback or foreland-propagating thrust sequences, where younger thrusts develop in the footwalls of existing thrusts, are generally assumed to be the typical order of thrust development in most orogenic settings. However, overstep or 'break-back' sequences, where later thrusts develop above and in the hangingwalls of earlier thrusts, may potentially form during cessation of movement in gravity-driven mass transport deposits (MTDs). In this study, we provide a detailed outcrop-based analysis of such an overstep thrust sequence developed in an MTD in the southern Dead Sea Basin. Evidence that may be used to discriminate overstep thrusting from piggyback thrust sequences within the gravity-driven fold and thrust belt includes upright folds and forethrusts that are cut by younger overlying thrusts. Backthrusts form ideal markers that are also clearly offset and cut by overlying younger forethrusts. Portions of the basal detachment to the thrust system are folded and locally imbricated in footwall synclines below forethrust ramps, and these geometries also support an overstep sequence. However, new 'short-cut' basal detachments develop below these synclines, indicating that movement continued on the basal detachment rather than it being abandoned as in classic overstep sequences. Further evidence for 'synchronous thrusting', where movement on more than one thrust occurs at the same time, is provided by displacement patterns on sequences of thrust ramp imbricates that systematically increases downslope towards the toe of the MTD. Older thrusts that initiate downslope in the broadly overstep sequence continue to move and therefore accrue greater displacements during synchronous thrusting. Our study provides a template to help distinguish different thrust sequences in both orogenic settings and gravity-driven surficial systems, with displacement patterns potentially being imaged in seismic sections across offshore MTDs.

WINGDES2 - WING DESIGN AND ANALYSIS CODE

NASA Technical Reports Server (NTRS)

Carlson, H. W.

1994-01-01

This program provides a wing design algorithm based on modified linear theory which takes into account the effects of attainable leading-edge thrust. A primary objective of the WINGDES2 approach is the generation of a camber surface as mild as possible to produce drag levels comparable to those attainable with full theoretical leading-edge thrust. WINGDES2 provides both an analysis and a design capability and is applicable to both subsonic and supersonic flow. The optimization can be carried out for designated wing portions such as leading and trailing edge areas for the design of mission-adaptive surfaces, or for an entire planform such as a supersonic transport wing. This program replaces an earlier wing design code, LAR-13315, designated WINGDES. WINGDES2 incorporates modifications to improve numerical accuracy and provides additional capabilities. A means of accounting for the presence of interference pressure fields from airplane components other than the wing and a direct process for selection of flap surfaces to approach the performance levels of the optimized wing surfaces are included. An increased storage capacity allows better numerical representation of those configurations that have small chord leading-edge or trailing-edge design areas. WINGDES2 determines an optimum combination of a series of candidate surfaces rather than the more commonly used candidate loadings. The objective of the design is the recovery of unrealized theoretical leading-edge thrust of the input flat surface by shaping of the design surface to create a distributed thrust and thus minimize drag. The input consists of airfoil section thickness data, leading and trailing edge planform geometry, and operational parameters such as Mach number, Reynolds number, and design lift coefficient. Output includes optimized camber surface ordinates, pressure coefficient distributions, and theoretical aerodynamic characteristics. WINGDES2 is written in FORTRAN V for batch execution and has been implemented on a CDC CYBER computer operating under NOS 2.7.1 with a central memory requirement of approximately 344K (octal) of 60 bit words. This program was developed in 1984, and last updated in 1990. CDC and CYBER are trademarks of Control Data Corporation.

Analysis of vortex shedding by 2-D numerical simulation for a solid rocket motor and calculations of the nonstationary thrust

NASA Astrophysics Data System (ADS)

Lupoglazoff, N.; Vuillot, F.

Periodic vortex shedding (VS) has been studied by 2-D numerical simulation for the C1 test case in the framework of the ASSM program concerning the stability of the Ariane-5 P230 solid rocket motor. The Flandro method is found to be unsuitable for the type of configuration considered here. The acoustic frequency of VS is a function of the configuration. Calculations of nonstationary thrust indicate that there is no direct relationship between the pressure oscillation amplitudes and the thrust. Secondary injection is found to have a stabilizing effect.

An overview of current Navy programs to develop thrust augmenting ejectors

NASA Technical Reports Server (NTRS)

Green, K. A.

1979-01-01

The primary objective of Navy sponsored research in thrust augmentation is the development of an improved augmenter for V/STOL application. In support of this goal, a data base is being established to provide an accurate prediction capability for use in ejector design. A general technology development of ejectors and associated effects presently is split into the more specific areas of lift and control, since thrust augmenting ejectors may be suitable for both. Research areas examined include advanced diffuser and end wall design; advanced primary nozzles; analytic studies; augmenting reaction controls; and nozzle design.

Linear test bed. Volume 1: Test bed no. 1. [aerospike test bed with segmented combustor

NASA Technical Reports Server (NTRS)

1972-01-01

The Linear Test Bed program was to design, fabricate, and evaluation test an advanced aerospike test bed which employed the segmented combustor concept. The system is designated as a linear aerospike system and consists of a thrust chamber assembly, a power package, and a thrust frame. It was designed as an experimental system to demonstrate the feasibility of the linear aerospike-segmented combustor concept. The overall dimensions are 120 inches long by 120 inches wide by 96 inches in height. The propellants are liquid oxygen/liquid hydrogen. The system was designed to operate at 1200-psia chamber pressure, at a mixture ratio of 5.5. At the design conditions, the sea level thrust is 200,000 pounds. The complete program including concept selection, design, fabrication, component test, system test, supporting analysis and posttest hardware inspection is described.

Research on design method of the full form ship with minimum thrust deduction factor

NASA Astrophysics Data System (ADS)

Zhang, Bao-ji; Miao, Ai-qin; Zhang, Zhu-xin

2015-04-01

In the preliminary design stage of the full form ships, in order to obtain a hull form with low resistance and maximum propulsion efficiency, an optimization design program for a full form ship with the minimum thrust deduction factor has been developed, which combined the potential flow theory and boundary layer theory with the optimization technique. In the optimization process, the Sequential Unconstrained Minimization Technique (SUMT) interior point method of Nonlinear Programming (NLP) was proposed with the minimum thrust deduction factor as the objective function. An appropriate displacement is a basic constraint condition, and the boundary layer separation is an additional one. The parameters of the hull form modification function are used as design variables. At last, the numerical optimization example for lines of after-body of 50000 DWT product oil tanker was provided, which indicated that the propulsion efficiency was improved distinctly by this optimal design method.

F-15B ACTIVE - First supersonic yaw vectoring flight

NASA Technical Reports Server (NTRS)

1996-01-01

On Wednesday, April 24, 1996, the F-15 Advanced Control Technology for Integrated Vehicles (ACTIVE) aircraft achieved its first supersonic yaw vectoring flight at Dryden Flight Research Center, Edwards, California. ACTIVE is a joint NASA, U.S. Air Force, McDonnell Douglas Aerospace (MDA) and Pratt & Whitney (P&W) program. The team will assess performance and technology benefits during flight test operations. Current plans call for approximately 60 flights totaling 100 hours. 'Reaching this milestone is very rewarding. We hope to set some more records before we're through,' stated Roger W. Bursey, P&W's pitch-yaw balance beam nozzle (PYBBN) program manager. A pair of P&W PYBBNs vectored (horizontally side-to-side, pitch is up and down) the thrust for the MDA manufactured F-15 research aircraft. Power to reach supersonic speeds was provided by two high-performance F100-PW-229 engines that were modified with the multi-directional thrust vectoring nozzles. The new concept should lead to significant increases in performance of both civil and military aircraft flying at subsonic and supersonic speeds.

Highly integrated digital engine control system on an F-15 airplane

NASA Technical Reports Server (NTRS)

Burcham, F. W., Jr.; Haering, E. A., Jr.

1984-01-01

The highly integrated digital electronic control (HIDEC) program will demonstrate and evaluate the improvements in performance and mission effectiveness that result from integrated engine-airframe control systems. This system is being used on the F-15 airplane at the Dryden Flight Research Facility of NASA Ames Research Center. An integrated flightpath management mode and an integrated adaptive engine stall margin mode are being implemented into the system. The adaptive stall margin mode is a highly integrated mode in which the airplane flight conditions, the resulting inlet distortion, and the engine stall margin are continuously computed; the excess stall margin is used to uptrim the engine for more thrust. The integrated flightpath management mode optimizes the flightpath and throttle setting to reach a desired flight condition. The increase in thrust and the improvement in airplane performance is discussed in this paper.

Combustion Fundamentals Research

NASA Technical Reports Server (NTRS)

1983-01-01

Increased emphasis is placed on fundamental and generic research at Lewis Research Center with less systems development efforts. This is especially true in combustion research, where the study of combustion fundamentals has grown significantly in order to better address the perceived long term technical needs of the aerospace industry. The main thrusts for this combustion fundamentals program area are as follows: analytical models of combustion processes, model verification experiments, fundamental combustion experiments, and advanced numeric techniques.

Evaluation of propellant tank insulation concepts for low-thrust chemical propulsion systems

NASA Technical Reports Server (NTRS)

Kramer, T.; Brogren, E.; Seigel, B.

1984-01-01

An analytical evaluation of cryogenic propellant tank insulations for liquid oxygen/liquid hydrogen low-thrust 2224N (500 lbf) propulsion systems (LTPS) was conducted. The insulation studied consisted of combinations of N2-purged foam and multilayer insulation (MLI) as well as He-purged MLI-only. Heat leak and payload performance predictions were made for three Shuttle-launched LTPS designed for Shuttle bay packaged payload densities of 56 kg/cu m, 40 kg/cu m and 24 kg/cu m. Foam/MLI insulations were found to increase LTPS payload delivery capability when compared with He-purged MLI-only. An additional benefit of foam/MLI was reduced operational complexity because Orbiter cargo bay N2 purge gas could be used for MLI purging. Maximum payload mass benefit occurred when an enhanced convection, rather than natural convection, heat transfer was specified for the insulation purge enclosure. The enhanced convection environment allowed minimum insulation thickness to be used for the foam/MLI interface temperature selected to correspond to the moisture dew point in the N2 purge gas. Experimental verification of foam/MLI benefits was recommended. A conservative program cost estimate for testing a MLI-foam insulated tank was 2.1 million dollars. It was noted this cost could be reduced significantly without increasing program risk.

Numerical Computation of a Continuous-thrust State Transition Matrix Incorporating Accurate Hardware and Ephemeris Models

NASA Technical Reports Server (NTRS)

Ellison, Donald; Conway, Bruce; Englander, Jacob

2015-01-01

A significant body of work exists showing that providing a nonlinear programming (NLP) solver with expressions for the problem constraint gradient substantially increases the speed of program execution and can also improve the robustness of convergence, especially for local optimizers. Calculation of these derivatives is often accomplished through the computation of spacecraft's state transition matrix (STM). If the two-body gravitational model is employed as is often done in the context of preliminary design, closed form expressions for these derivatives may be provided. If a high fidelity dynamics model, that might include perturbing forces such as the gravitational effect from multiple third bodies and solar radiation pressure is used then these STM's must be computed numerically. We present a method for the power hardward model and a full ephemeris model. An adaptive-step embedded eight order Dormand-Prince numerical integrator is discussed and a method for the computation of the time of flight derivatives in this framework is presented. The use of these numerically calculated derivatieves offer a substantial improvement over finite differencing in the context of a global optimizer. Specifically the inclusion of these STM's into the low thrust missiondesign tool chain in use at NASA Goddard Spaceflight Center allows for an increased preliminary mission design cadence.

Short-term combined effects of thoracic spine thrust manipulation and cervical spine nonthrust manipulation in individuals with mechanical neck pain: a randomized clinical trial.

PubMed

Masaracchio, Michael; Cleland, Joshua A; Hellman, Madeleine; Hagins, Marshall

2013-03-01

Randomized clinical trial. To investigate the short-term effects of thoracic spine thrust manipulation combined with cervical spine nonthrust manipulation (experimental group) versus cervical spine nonthrust manipulation alone (comparison group) in individuals with mechanical neck pain. Research has demonstrated improved outcomes with both nonthrust manipulation directed at the cervical spine and thrust manipulation directed at the thoracic spine in patients with neck pain. Previous studies have not determined if thoracic spine thrust manipulation may increase benefits beyond those provided by cervical nonthrust manipulation alone. Sixty-four participants with mechanical neck pain were randomized into 1 of 2 groups, an experimental or comparison group. Both groups received 2 treatment sessions of cervical spine nonthrust manipulation and a home exercise program consisting of active range-of-motion exercises, and the experimental group received additional thoracic spine thrust manipulations. Outcome measures were collected at baseline and at a 1-week follow-up, and included the numeric pain rating scale, the Neck Disability Index, and the global rating of change. Participants in the experimental group demonstrated significantly greater improvements (P<.001) on both the numeric pain rating scale and Neck Disability Index at the 1-week follow-up compared to those in the comparison group. In addition, 31 of 33 (94%) participants in the experimental group, compared to 11 of 31 participants (35%) in the comparison group, indicated a global rating of change score of +4 or higher at the 1-week follow-up, with an associated number needed to treat of 2. Individuals with neck pain who received a combination of thoracic spine thrust manipulation and cervical spine nonthrust manipulation plus exercise demonstrated better overall short-term outcomes on the numeric pain rating scale, the Neck Disability Index, and the global rating of change.

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.

Saturn Apollo Program

NASA Image and Video Library

1965-03-01

This photograph shows F-1 engines being stored in the F-1 Engine Preparation Shop, building 4666, at the Marshall Space Flight Center. Each F-1 engine produced a thrust of 1,500,000 pounds. A cluster of five engines was mounted on the thrust structure of the S-IC stage of a 364-foot long Saturn V launch vehicle that ultimately took astronauts to the Moon.

Propulsion Flight Research at NASA Dryden From 1967 to 1997

NASA Technical Reports Server (NTRS)

Burcham, Frank W., Jr.; Ray, Ronald J.; Conners, Timothy R.; Walsh, Kevin R.

1997-01-01

From 1967 to 1997, pioneering propulsion flight research activities have been conceived and conducted at the NASA Dryden Flight Research Center. Many of these programs have been flown jointly with the United States Department of Defense, industry, or the Federal Aviation Administration. Propulsion research has been conducted on the XB-70, F-111 A, F-111E, YF-12, JetStar, B-720, MD-11, F-15, F- 104, Highly Maneuverable Aircraft Technology, F-14, F/A-18, SR-71, and the hypersonic X-15 airplanes. Research studies have included inlet dynamics and control, in-flight thrust computation, integrated propulsion controls, inlet and boattail drag, wind tunnel-to-flight comparisons, digital engine controls, advanced engine control optimization algorithms, acoustics, antimisting kerosene, in-flight lift and drag, throttle response criteria, and thrust-vectoring vanes. A computer-controlled thrust system has been developed to land the F-15 and MD-11 airplanes without using any of the normal flight controls. An F-15 airplane has flown tests of axisymmetric thrust-vectoring nozzles. A linear aerospike rocket experiment has been developed and tested on the SR-71 airplane. This paper discusses some of the more unique flight programs, the results, lessons learned, and their impact on current technology.

Test program to provide confidence in liquid oxygen cooling of hydrocarbon fueled rocket thrust chambers

NASA Technical Reports Server (NTRS)

Armstrong, E. S.

1986-01-01

An experimental program has been planned at the NASA Lewis Research Center to build confidence in the feasibility of liquid oxygen cooling for hydrocarbon fueled rocket engines. Although liquid oxygen cooling has previously been incorporated in test hardware, more runtime is necessary to gain confidence in this concept. In the previous tests, small oxygen leaks developed at the throat of the thrust chamber and film cooled the hot-gas side of the chamber wall without resulting in catastrophic failure. However, more testing is necessary to demonstrate that a catastrophic failure would not occur if cracks developed further upstream between the injector and the throat, where the boundary layer has not been established. Since under normal conditions cracks are expected to form in the throat region of the thrust chamber, cracks must be initiated artificially in order to control their location. Several methods of crack initiation are discussed in this report. Four thrust chambers, three with cracks and one without, should be tested. The axial location of the cracks should be varied parametrically. Each chamber should be instrumented to determine the effects of the cracks, as well as the overall performance and durability of the chambers.

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.

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.

Fluid lubricated bearing assembly

DOEpatents

Boorse, Henry A.; Boeker, Gilbert F.; Menke, John R.

1976-01-01

1. A support for a loaded rotatable shaft comprising in combination on a housing having a fluid-tight cavity encasing an end portion of said shaft, a thrust bearing near the open end of said cavity for supporting the axial thrust of said shaft, said thrust bearing comprising a thrust plate mounted in said housing and a thrust collar mounted on said shaft, said thrust plate having a central opening the peripheral portion of which is hermetically sealed to said housing at the open end of said cavity, and means for supplying a fluid lubricant to said thrust bearing, said thrust bearing having a lubricant-conducting path connecting said lubricant supplying means with the space between said thrust plate and collar intermediate the peripheries thereof, the surfaces of said plate and collar being constructed and arranged to inhibit radial flow of lubricant and, on rotation of said thrust collar, to draw lubricant through said path between the bearing surfaces and to increase the pressure therebetween and in said cavity and thereby exert a supporting force on said end portion of said shaft.

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.

Study to document low thrust trajectory optimization programs HILTOP and ASTOP

NASA Technical Reports Server (NTRS)

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

1974-01-01

Detailed documentation of the HILTOP and ASTOP computer programs is presented along with results of the analyses of the possible extension of the HILTOP program and results of an extra-ecliptic mission study performed with HILTOP.

Orbital Transfer Vehicle (OTV) engine study. Phase A: Extension

NASA Technical Reports Server (NTRS)

Sobin, A. J.

1980-01-01

The current Phase A-Extension of the OTV engine study program aims to provide additional expander and staged combustion cycle data that will lead to design definition of the OTV engine. The proposed program effort seeks to optimize the expander cycle engine concept (consistent with identified OTV engine requirements), investigate the feasibility of kitting the staged combustion cycle engine to provide extended thrust operation, and conduct in-depth analysis of development risk, crew safety, and reliability for both cycles. Additional tasks address the costing of a 10/K thrust expander cycle engine and support of OTV systems study contractors.

Low-thrust trajectory analysis for the geosynchronous mission

NASA Technical Reports Server (NTRS)

Jasper, T. P.

1973-01-01

Methodology employed in development of a computer program designed to analyze optimal low-thrust trajectories is described, and application of the program to a Solar Electric Propulsion Stage (SEPS) geosynchronous mission is discussed. To avoid the zero inclination and eccentricity singularities which plague many small-force perturbation techniques, a special set of state variables (equinoctial) is used. Adjoint equations are derived for the minimum time problem and are also free from the singularities. Solutions to the state and adjoint equations are obtained by both orbit averaging and precision numerical integration; an evaluation of these approaches is made.

An automated approach to design of solid rockets utilizing a special internal ballistics model

NASA Technical Reports Server (NTRS)

Sforzini, R. H.

1980-01-01

A pattern search technique is presented, which is utilized in a computer program that minimizes the sum of the squares of the differences, at various times, between a desired thrust-time trace and that calculated with a special mathematical internal ballistics model of a solid propellant rocket motor. The program is demonstrated by matching the thrust-time trace obtained from static tests of the first Space Shuttle SRM starting with input values of 10 variables which are, in general, 10% different from the as-built SRM. It is concluded that an excellent match is obtained.

Spacing of Imbricated Thrust Faults and the Strength of Thrust-Belts and Accretionary Wedges

NASA Astrophysics Data System (ADS)

Ito, G.; Regensburger, P. V.; Moore, G. F.

2017-12-01

The pattern of imbricated thrust blocks is a prominent characteristic of the large-scale structure of thrust-belts and accretionary wedges around the world. Mechanical models of these systems have a rich history from laboratory analogs, and more recently from computational simulations, most of which, qualitatively reproduce the regular patterns of imbricated thrusts seen in nature. Despite the prevalence of these patterns in nature and in models, our knowledge of what controls the spacing of the thrusts remains immature at best. We tackle this problem using a finite difference, particle-in-cell method that simulates visco-elastic-plastic deformation with a Mohr-Coulomb brittle failure criterion. The model simulates a horizontal base that moves toward a rigid vertical backstop, carrying with it an overlying layer of crust. The crustal layer has a greater frictional strength than the base, is cohesive, and is initially uniform in thickness. As the layer contracts, a series of thrust blocks immerge sequentially and form a wedge having a mean taper consistent with that predicted by a noncohesive, critical Coulomb wedge. The widths of the thrust blocks (or spacing between adjacent thrusts) are greatest at the front of the wedge, tend to decrease with continued contraction, and then tend toward a pseudo-steady, minimum width. Numerous experiments show that the characteristic spacing of thrusts increases with the brittle strength of the wedge material (cohesion + friction) and decreases with increasing basal friction for low (<8°) taper angles. These relations are consistent with predictions of the elastic stresses forward of the frontal thrust and at what distance the differential stress exceeds the brittle threshold to form a new frontal thrust. Hence the characteristic spacing of the thrusts across the whole wedge is largely inherited at the very front of the wedge. Our aim is to develop scaling laws that will illuminate the basic physical processes controlling systems, as well as allow researchers to use observations of thrust spacing as an independent constraint on the brittle strength of wedges as well as their bases.

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.

Solid propellant rocket motor internal ballistics performance variation analysis, phase 3

NASA Technical Reports Server (NTRS)

Sforzini, R. H.; Foster, W. A., Jr.; Murph, J. E.; Adams, G. W., Jr.

1977-01-01

Results of research aimed at improving the predictability of off nominal internal ballistics performance of solid propellant rocket motors (SRMs) including thrust imbalance between two SRMs firing in parallel are reported. The potential effects of nozzle throat erosion on internal ballistic performance were studied and a propellant burning rate low postulated. The propellant burning rate model when coupled with the grain deformation model permits an excellent match between theoretical results and test data for the Titan IIIC, TU455.02, and the first Space Shuttle SRM (DM-1). Analysis of star grain deformation using an experimental model and a finite element model shows the star grain deformation effects for the Space Shuttle to be small in comparison to those of the circular perforated grain. An alternative technique was developed for predicting thrust imbalance without recourse to the Monte Carlo computer program. A scaling relationship used to relate theoretical results to test results may be applied to the alternative technique of predicting thrust imbalance or to the Monte Carlo evaluation. Extended investigation into the effect of strain rate on propellant burning rate leads to the conclusion that the thermoelastic effect is generally negligible for both steadily increasing pressure loads and oscillatory loads.

Affordable Development and Demonstration of a Small NTR engine and Stage: A Preliminary NASA, DOE, and Industry Assessment

NASA Technical Reports Server (NTRS)

Borowski, S. K.; Sefcik, R. J.; Fittje, J. E.; McCurdy, D. R.; Qualls, A. L.; Schnitzler, B. G; Werner, J.; Weitzberg, A.; Joyner, C. R.

2015-01-01

In FY'11, Nuclear Thermal Propulsion (NTP) was identified as a key propulsion option under the Advanced In-Space Propulsion (AISP) component of NASA's Exploration Technology Development and Demonstration (ETDD) program A strategy was outlined by GRC and NASA HQ that included 2 key elements -"Foundational Technology Development" followed by specific "Technology Demonstration" projects. The "Technology Demonstration "element proposed ground technology demonstration (GTD) testing in the early 2020's, followed by a flight technology demonstration (FTD) mission by approx. 2025. In order to reduce development costs, the demonstration projects would focus on developing a small, low thrust (approx. 7.5 -16.5 klb(f)) engine that utilizes a "common" fuel element design scalable to the higher thrust (approx. 25 klb(f)) engines used in NASA's Mars DRA 5.0 study(NASA-SP-2009-566). Besides reducing development costs and allowing utilization of existing, flight proven engine hard-ware (e.g., hydrogen pumps and nozzles), small, lower thrust ground and flight demonstration engines can validate the technology and offer improved capability -increased payloads and decreased transit times -valued for robotic science missions identified in NASA's Decadal Study.

737/CFM56-7 Aircraft Engine Systems

NASA Technical Reports Server (NTRS)

Wright, Steve; Shiosaki, Justin

2006-01-01

The configuration of the propulsion system engine externals must meet many airplane requirements such as cost, thrust, weight, range and systems power extraction. On the 737-700 several program requirements also played a major role in the development of the engine externals. These program goals were increased range, same cost as a 1994 737-300, 15% reduction in maintenance costs from the 737-300, and a propulsion package that appeared as if it was designed by one company. This presentation will show how these requirements shaped the design of the engine externals for the 737-700/CFM56-7B.

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.

YF-17/ADEN system study

NASA Technical Reports Server (NTRS)

Gowadia, N. S.; Bard, W. D.; Wooten, W. H.

1979-01-01

The YF-17 aircraft was evaluated as a candidate nonaxisymmetric nozzle flight demonstrator. Configuration design modifications, control system design, flight performance assessment, and program plan and cost we are summarized. Two aircraft configurations were studied. The first was modified as required to install only the augmented deflector exhaust nozzle (ADEN). The second one added a canard installation to take advantage of the full (up to 20 deg) nozzle vectoring capability. Results indicate that: (1) the program is feasible and can be accomplished at reasonable cost and low risk; (2) installation of ADEN increases the aircraft weight by 600 kg (1325 lb); (3) the control system can be modified to accomplish direct lift, pointing capability, variable static margin and deceleration modes of operation; (4) unvectored thrust-minus-drag is similar to the baseline YF-17; and (5) vectoring does not improve maneuvering performance. However, some potential benefits in direct lift, aircraft pointing, handling at low dynamic pressure and takeoff/landing ground roll are available. A 27 month program with 12 months of flight test is envisioned, with the cost estimated to be $15.9 million for the canard equipped aircraft and $13.2 million for the version without canard. The feasiblity of adding a thrust reverser to the YF-17/ADEN was investigated.

Method and apparatus for rapid thrust increases in a turbofan engine

NASA Technical Reports Server (NTRS)

Cornett, J. E.; Corley, R. C.; Fraley, T. O.; Saunders, A. A., Jr. (Inventor)

1980-01-01

Upon a landing approach, the normal compressor stator schedule of a fan speed controlled turbofan engine is temporarily varied to substantially close the stators to thereby increase the fuel flow and compressor speed in order to maintain fan speed and thrust. This running of the compressor at an off-design speed substantially reduces the time required to subsequently advance the engine speed to the takeoff thrust level by advancing the throttle and opening the compressor stators.

Saturn Apollo Program

NASA Image and Video Library

1962-06-07

This photograph depicts the Rocketdyne static firing of the F-1 engine at the towering 76-meter Test Stand 1-C in Area 1-125 of the Edwards Air Force Base in California. The Saturn V S-IC (first) stage utilized five F-1 engines for its thrust. Each engine provided 1,500,000 pounds, for a combined thrust of 7,500,000 pounds with liquid oxygen and kerosene as its propellants.

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.

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.

Saturn Apollo Program

NASA Image and Video Library

1960-01-01

A Cluster of eight H-1 engines were used to thrust the first stage of Saturn I (S-I stage) and Saturn IB (S-IB stage). The engines were arranged in a double pattern. Four engines, located inboard, were fixed in a square pattern around the stage axis, while the remaining four engines were located outboard in a larger square pattern and each outer engine was gimbaled. Each H-1 engine, fueled with liquid oxygen (LOX) and kerosene (RP-1), had a thrust of 188,000 pound each for a combined thrust of over 1,500,000 pounds. The H-1 engine was developed under the direction of Marshall Space Flight Center (MSFC).

Saturn Apollo Program

NASA Image and Video Library

1960-01-01

A Cluster of eight H-1 engines were used to thrust the first stage of Saturn I (S-I stage) and Saturn IB (S-IB stage). The engines were arranged in a double pattern. Four engines, located inboard, were fixed in a square pattern around the stage axis, while the remaining four engines were located outboard in a larger square pattern and each outer engine was gimbaled. The H-1 engine, fueled with liquid oxygen (LOX) and kerosene (RP-1), had a thrust of 188,000 pound each for a combined thrust of over 1,500,000 pounds. Each H-1 engine was developed under the direction of Marshall Space Flight Center (MSFC).

DoD High Performance Computing Modernization Program FY16 Annual Report

DTIC Science & Technology

2018-05-02

vortex shedding from rotor blade tips using adaptive mesh refinement gives Helios the unique capability to assess the interaction of these vortices...with the fuselage and nearby rotor blades . Helios provides all the benefits for rotary-winged aircraft that Kestrel does for fixed-wing aircraft...rotor blade upgrade of the CH-47F Chinook helicopter to achieve up to an estimated 2,000 pounds increase in hover thrust (~10%) with limited

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.

Lateral thrust of anterior cruciate ligament-insufficient knees and posterior cruciate ligament-insufficient knees.

PubMed

Yoshimura, Ichiro; Naito, Masatoshi; Zhang, Jingfan

2002-01-01

Leaving anterior cruciate ligament (ACL) insufficiency and posterior cruciate ligament (PCL) insufficiency untreated frequently leads to osteoarthritis (OA). The purpose of this study was to evaluate dynamically the lateral thrust of ACL-insufficient knees and PCL-insufficient knees, and from the findings investigate the relationship between cruciate ligament insufficiency and OA occurrence. An acceleration sensor was attached to the affected and control anterior tibial tubercles, acting in medial-lateral and perpendicular directions. The lateral thrust immediately after heel strike was measured continuously by a telemeter under stabilised walking conditions. When compared to the contralateral healthy knee, the peak value of lateral acceleration immediately after heel strike was significantly larger in the ACL-insufficient knee; and lateral thrust was increased, but not significantly, in the PCL-insufficient knee. Given that lateral thrust of the knee during walking increases due to ACL or PCL injury, it may be a principal contributor to OA progression.

Power-by-Wire Development and Demonstration for Subsonic Civil Transport

NASA Technical Reports Server (NTRS)

1996-01-01

During the last decade, three significant studies by the Lockheed Martin Corporation, the NASA Lewis Research Center, and McDonnell Douglas Corporation have clearly shown operational, weight, and cost advantages for commercial subsonic transport aircraft that use all-electric or more-electric technologies in the secondary electric power systems. Even though these studies were completed on different aircraft, used different criteria, and applied a variety of technologies, all three have shown large benefits to the aircraft industry and to the nation's competitive position. The Power-by-Wire (PBW) program is part of the highly reliable Fly-By-Light/Power-By-Wire (FBL/PBW) Technology Program, whose goal is to develop the technology base for confident application of integrated FBL/PBW systems for transport aircraft. This program is part of the NASA aeronautics strategic thrust in subsonic aircraft/national airspace (Thrust 1) to "develop selected high-leverage technologies and explore new means to ensure the competitiveness of U.S. subsonic aircraft and to enhance the safety and productivity of the national aviation system" (The Aeronautics Strategic Plan). Specifically, this program is an initiative under Thrust 1, Key Objective 2, to "develop, in cooperation with U.S. industry, selected high-payoff technologies that can enable significant improvements in aircraft efficiency and cost."

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.

Rotational reflectance of dispersed vitrinite from the Arkoma basin

USGS Publications Warehouse

Houseknecht, D.W.; Weesner, C.M.B.

1997-01-01

Rotational reflectance of dispersed vitrinite provides superior documentation of thermal maturity and a capability for interpreting relative timing between thermal and kinematic events in Arkoma Basin strata characterized by vitrinite reflectances up to 5%. Rotational reflectance (R(rot)) is a more precise and less ambiguous index of thermal maturity than maximum (R'(max)), minimum (R(min)), and random (R(ran)) reflectance. Vitrinite reflectance anisotropy becomes sufficiently large to be measurable (using a microscope equipped with an automated rotating polarizer) at ???2% R(rot) and increases following a power function with increasing thermal maturity. Rotational reflectance data can be used to infer the shape of the vitrinite reflectance indicating surface (i.e. indicatrix) and, in turn, to enhance interpretations of the timing between thermal maxima and compressional tectonic events. Data from three wells in the Arkoma Basin Ouachita frontal thrust belt are used as examples. The absence of offsets in measured R(rot) across thrust faults combined with a predominance of uniaxial vitrinite in the thrust faulted part of the section suggest thermal maximum postdated thrust faulting in the western Ouachita frontal thrust belt of Oklahoma. In contrast, the general absence of offsets in measured R(rot) across thrust faults combined with a predominance of biaxial vitrinite in the thrust faulted part of the section suggest that the thermal maximum was coeval with thrust faulting in the eastern Ouachita frontal thrust belt of Arkansas. The presence of biaxial vitrinite in an allochthonous section and uniaxial vitrinite in an underlying, autochthonous section suggests that the thermal maximum was coeval with listric thrust faulting in the central Arkoma Basin of Oklahoma, and that rotational reflectance data can be used as a strain indicator to detect subtle decollement zones.

Replacing critical and strategic refractory metal elements in nickel-base superalloys. [NASA's COSAM program

NASA Technical Reports Server (NTRS)

Stephens, J. R.; Dreshfield, R. L.; Nathal, M. V.

1983-01-01

Because of the import status and essential nature of their use, cobalt, chromium, tantalum, and niobium were identified as strategic and critical in the aerospace industry. NASA's Conservation of Strategic Aerospace Materials (COSAM) program aims to reduce the need for strategic materials used in gas turbine engines. Technological thrusts in two major areas are under way to meet the primary objective of conserving the use of strategic materials in nickelbase superalloys. These thrusts consist of strategic element substitution and alternative material identification. The program emphasizes cooperative research teams involving NASA Lewis Research Center, universities, and industry. The adoption of refractory metals in nickel-base superalloys is summarized including their roles in mechanical strengthening and environmental resistance; current research activities under way in the COSAM Program are presented as well as research findings to date.

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.

14 CFR Appendix I to Part 25 - Installation of an Automatic Takeoff Thrust Control System (ATTCS)

Code of Federal Regulations, 2010 CFR

2010-01-01

...) This appendix specifies additional requirements for installation of an engine power control system that... crew to increase thrust or power. I25.2Definitions. (a) Automatic Takeoff Thrust Control System (ATTCS... mechanical and electrical, that sense engine failure, transmit signals, actuate fuel controls or power levers...

Influence of the thrust bearing on the natural frequencies of a 72-MW hydropower rotor

NASA Astrophysics Data System (ADS)

Cupillard, S.; Aidanpää, J.-O.

2016-11-01

The thrust bearing is an essential element of a hydropower machine. Not only does it carry the total axial load but it also introduces stiffness and damping properties in the system. The focus of this study is on the influence of the thrust bearing on the lateral vibrations of the shaft of a 72-MW propeller turbine. The thrust bearing has a non-conventional design with a large radius and two rows of thrust pads. A numerical model is developed to estimate natural frequencies. Numerical results are analyzed and related to experimental measurements of a runaway test. The results show the need to include the thrust bearing in the model. In fact, the vibration modes are substantially increased towards higher frequencies with the added properties from the thrust bearing. The second mode of vibration has been identified in the experimental measurements. Its frequency and mode shape compare well with numerical results.

Low thrust vehicle concept study

NASA Technical Reports Server (NTRS)

1980-01-01

Low thrust chemical (hydrogen-oxygen) propulsion systems configured specifically for low acceleration orbit transfer of large space systems were defined. Results indicate that it is cost effective and least risk to combine the OTV and stowed spacecraft in a single 65 K Shuttle. The study shows that the engine for an optimized low thrust stage (1) does not require very low thrust; (2) 1-3 K thrust range appears optimum; (3) thrust transient is not a concern; (4) throttling probably not worthwhile; and (5) multiple thrusters complicate OTV/LSS design and aggravate LSS loads. Regarding the optimum vehicle for low acceleration missions, the single shuttle launch (LSS and expendable OTV) is most cost effective and least risky. Multiple shuttles increase diameter 20%. The space based radar structure short OTV (which maximizes space available for packaged LSS) favors use of torus tank. Propellant tank pressures/vapor residuals are little affected by engine thrust level or number of burns.

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.

Dynamic Information Collection and Fusion

DTIC Science & Technology

2015-12-02

AFRL-AFOSR-VA-TR-2016-0069 DYNAMIC INFORMATION COLLECTION AND FUSION Venugopal Veeravalli UNIVERSITY OF ILLINOIS CHAMPAIGN Final Report 12/02/2015...TITLE AND SUBTITLE Dynamic Information Collection and Fusion 5a. CONTRACT NUMBER FA9550-10-1-0458 5b. GRANT NUMBER AF FA9550-10-1-0458 5c. PROGRAM...information collection, fusion , and inference from diverse modalities Our research has been organized under three inter-related thrusts. The first thrust

Saturn Apollo Program

NASA Image and Video Library

1969-01-01

In one of the initial assembly steps for the Saturn IB launch vehicle's S-IB (first) stage, workers at the Michoud Assembly Facility (MAF) near New Orleans, Louisiana, complete the lower shroud assembly. Developed by the Marshall Space Flight Center and built by the Chrysler Corporation at Michoud Assembly Facility (MAF), the S-IB utilized the eight H-1 engines and each produced 200,000 pounds of thrust, a combined thrust of 1,600,000 pounds.

Status of Low Thrust Work at JSC

NASA Technical Reports Server (NTRS)

Condon, Gerald L.

2004-01-01

High performance low thrust (solar electric, nuclear electric, variable specific impulse magnetoplasma rocket) propulsion offers a significant benefit to NASA missions beyond low Earth orbit. As NASA (e.g., Prometheus Project) endeavors to develop these propulsion systems and associated power supplies, it becomes necessary to develop a refined trajectory design capability that will allow engineers to develop future robotic and human mission designs that take advantage of this new technology. This ongoing work addresses development of a trajectory design and optimization tool for assessing low thrust (and other types) trajectories. This work targets to advance the state of the art, enable future NASA missions, enable science drivers, and enhance education. This presentation provides a summary of the low thrust-related JSC activities under the ISP program and specifically, provides a look at a new release of a multi-gravity, multispacecraft trajectory optimization tool (Copernicus) along with analysis performed using this tool over the past year.

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.

Thrust chamber material technology program

NASA Technical Reports Server (NTRS)

Andrus, J. S.; Bordeau, R. G.

1989-01-01

This report covers work performed at Pratt & Whitney on development of copper-based materials for long-life, reusable, regeneratively cooled rocket engine thrust chambers. The program approached the goal of enhanced cyclic life through the application of rapid solidification to alloy development, to introduce fine dispersions to strengthen and stabilize the alloys at elevated temperatures. After screening of alloy systems, copper-based alloys containing Cr, Co, Hf, Ag, Ti, and Zr were processed by rapid-solidification atomization in bulk quantities. Those bulk alloys showing the most promise were characterized by tensile testing, thermal conductivity testing, and elevated-temperature, low-cycle fatigue (LFC) testing. Characterization indicated that Cu- 1.1 percent Hf exhibited the greatest potential as an improved-life thrust chamber material, exhibiting LCF life about four times that of NASA-Z. Other alloys (Cu- 0.6 percent Zr, and Cu- 0.6 percent Zr- 1.0 percent Cr) exhibited promise for use in this application, but needed more development work to balance properties.

Propulsion technology needs for advanced space transportation systems. [orbit maneuvering engine (space shuttle), space shuttle boosters

NASA Technical Reports Server (NTRS)

Gregory, J. W.

1975-01-01

Plans are formulated for chemical propulsion technology programs to meet the needs of advanced space transportation systems from 1980 to the year 2000. The many possible vehicle applications are reviewed and cataloged to isolate the common threads of primary propulsion technology that satisfies near term requirements in the first decade and at the same time establish the technology groundwork for various potential far term applications in the second decade. Thrust classes of primary propulsion engines that are apparent include: (1) 5,000 to 30,000 pounds thrust for upper stages and space maneuvering; and (2) large booster engines of over 250,000 pounds thrust. Major classes of propulsion systems and the important subdivisions of each class are identified. The relative importance of each class is discussed in terms of the number of potential applications, the likelihood of that application materializing, and the criticality of the technology needed. Specific technology programs are described and scheduled to fulfill the anticipated primary propulsion technology requirements.

The aerodynamic characteristics of large angled cones with retrorockets

NASA Technical Reports Server (NTRS)

Jarvinen, P. O.; Adams, R. H.

1970-01-01

Analytical and experimental phases of the subject investigation are described. The analytical program for the single jet determines the terminal shock location, the jet boundary, the interface profile, the bow shock profile, the shear layer growth and the dead air region pressure. The experimental program described was conducted over the range from free stream Mach 0.4 to 2.0 at angles-of-attack up to 18 deg and at thrusting coefficients up to C sub T = T/q sub infinity A sub m = 30. Variables investigated included aeroshell angle, number of nozzles, engine thrust, size of nozzles, nozzle throttling and gas composition. The influence of these variables on the aeroshell stability, drag, and loads was determined by integrating pressure measurements on the aeroshell. The total system forces consist of components due to pure thrust and components due to pressure on the aeroshell arising from the jet-free stream interaction. Shadowgraphs provided flow field geometries which proved to be within 10% of those predicted analytically.

Passive Thrust Oscillation Mitigation for the CEV Crew Pallet System

NASA Technical Reports Server (NTRS)

Sammons, Matthew; Powell, Cory; Pellicciotti, Joseph; Buehrle, Ralph; Johnson, Keith

2012-01-01

The Crew Exploration Vehicle (CEV) was intended to be the next-generation human spacecraft for the Constellation Program. The CEV Isolator Strut mechanism was designed to mitigate loads imparted to the CEV crew caused by the Thrust Oscillation (TO) phenomenon of the proposed Ares I Launch Vehicle (LV). The Isolator Strut was also designed to be compatible with Launch Abort (LA) contingencies and landing scenarios. Prototype struts were designed, built, and tested in component, sub-system, and system-level testing. The design of the strut, the results of the tests, and the conclusions and lessons learned from the program will be explored in this paper.

Generation 1.5 High Speed Civil Transport (HSCT) Exhaust Nozzle Program

NASA Technical Reports Server (NTRS)

Thayer, E. B.; Gamble, E. J.; Guthrie, A. R.; Kehret, D. F.; Barber, T. J.; Hendricks, G. J.; Nagaraja, K. S.; Minardi, J. E.

2004-01-01

The objective of this program was to conduct an experimental and analytical evaluation of low noise exhaust nozzles suitable for future High-Speed Civil Transport (HSCT) aircraft. The experimental portion of the program involved parametric subscale performance model tests of mixer/ejector nozzles in the takeoff mode, and high-speed tests of mixer/ejectors converted to two-dimensional convergent-divergent (2-D/C-D), plug, and single expansion ramp nozzles (SERN) in the cruise mode. Mixer/ejector results show measured static thrust coefficients at secondary flow entrainment levels of 70 percent of primary flow. Results of the high-speed performance tests showed that relatively long, straight-wall, C-D nozzles could meet supersonic cruise thrust coefficient goal of 0.982; but the plug, ramp, and shorter C-D nozzles required isentropic contours to reach the same level of performance. The computational fluid dynamic (CFD) study accurately predicted mixer/ejector pressure distributions and shock locations. Heat transfer studies showed that a combination of insulation and convective cooling was more effective than film cooling for nonafterburning, low-noise nozzles. The thrust augmentation study indicated potential benefits for use of ejector nozzles in the subsonic cruise mode if the ejector inlet contains a sonic throat plane.

Presentations - Delaney, P.R. and others, 2008 | Alaska Division of

Science.gov Websites

, northeastern Alaska (poster): AAPG Abstracts with Programs, San Antonio, Texas Authors: Delaney, P.R., Loveland thrust belt, Kavik River area, northeastern Alaska (poster): AAPG Abstracts with Programs, San Antonio

Technology transfer

NASA Technical Reports Server (NTRS)

Penaranda, Frank E.

1992-01-01

The topics are presented in viewgraph form and include the following: international comparison of R&D expenditures in 1989; NASA Technology Transfer Program; NASA Technology Utilization Program thrusts for FY 1992 and FY 1993; National Technology Transfer Network; and NTTC roles.

Spinal manipulation force and duration affect vertebral movement and neuromuscular responses.

PubMed

Colloca, Christopher J; Keller, Tony S; Harrison, Deed E; Moore, Robert J; Gunzburg, Robert; Harrison, Donald D

2006-03-01

Previous study in human subjects has documented biomechanical and neurophysiological responses to impulsive spinal manipulative thrusts, but very little is known about the neuromechanical effects of varying thrust force-time profiles. Ten adolescent Merino sheep were anesthetized and posteroanterior mechanical thrusts were applied to the L3 spinous process using a computer-controlled, mechanical testing apparatus. Three variable pulse durations (10, 100, 200 ms, force = 80 N) and three variable force amplitudes (20, 40, 60 N, pulse duration = 100 ms) were examined for their effect on lumbar motion response (L3 displacement, L1, L2 acceleration) and normalized multifidus electromyographic response (L3, L4) using a repeated measures analysis of variance. Increasing L3 posteroanterior force amplitude resulted in a fourfold linear increase in L3 posteroanterior vertebral displacement (p < 0.001) and adjacent segment (L1, L2) posteroanterior acceleration response (p < 0.001). L3 displacement was linearly correlated (p < 0.001) to the acceleration response over the 20-80 N force range (100 ms). At constant force, 10 ms thrusts resulted in nearly fivefold lower L3 displacements and significantly increased segmental (L2) acceleration responses compared to the 100 ms (19%, p = 0.005) and 200 ms (16%, p = 0.023) thrusts. Normalized electromyographic responses increased linearly with increasing force amplitude at higher amplitudes and were appreciably affected by mechanical excitation pulse duration. Changes in the biomechanical and neuromuscular response of the ovine lumbar spine were observed in response to changes in the force-time characteristics of the spinal manipulative thrusts and may be an underlying mechanism in related clinical outcomes.

Thrust Vectoring on the NASA F-18 High Alpha Research Vehicle

NASA Technical Reports Server (NTRS)

Bowers, Albion H.; Pahle, Joseph W.

1996-01-01

Investigations into a multiaxis thrust-vectoring system have been conducted on an F-18 configuration. These investigations include ground-based scale-model tests, ground-based full-scale testing, and flight testing. This thrust-vectoring system has been tested on the NASA F-18 High Alpha Research Vehicle (HARV). The system provides thrust vectoring in pitch and yaw axes. Ground-based subscale test data have been gathered as background to the flight phase of the program. Tests investigated aerodynamic interaction and vane control effectiveness. The ground-based full-scale data were gathered from static engine runs with image analysis to determine relative thrust-vectoring effectiveness. Flight tests have been conducted at the NASA Dryden Flight Research Center. Parameter identification input techniques have been developed. Individual vanes were not directly controlled because of a mixer-predictor function built into the flight control laws. Combined effects of the vanes have been measured in flight and compared to combined effects of the vanes as predicted by the cold-jet test data. Very good agreement has been found in the linearized effectiveness derivatives.

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.

TMS delivered for A-3 Test Stand

NASA Image and Video Library

2010-03-17

A state-of-the-art thrust measurement system for the A-3 Test Stand under construction at NASA's John C. Stennis Space Center was delivered March 17. Once completed, the A-3 stand (seen in background) will allow simulated high-altitude testing on the next generation of rocket engines for America's space program. Work on the stand began in 2007, with activation scheduled for 2012. The stand is the first major test structure to be built at Stennis since the 1960s. The recently delivered TMS was fabricated by Thrust Measurement Systems in Illinois. It is an advanced calibration system capable of measuring vertical and horizontal thrust loads with an accuracy within 0.15 percent at 225,000 pounds.

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).

Electric propulsion technology

NASA Technical Reports Server (NTRS)

Finke, R. C.

1980-01-01

The advanced electric propulsion program is directed towards lowering the specific impulse and increasing the thrust per unit of ion thruster systems. In addition, electrothermal and electromagnetic propulsion technologies are being developed to attempt to fill the gap between the conventional ion thruster and chemical rocket systems. Most of these new concepts are exagenous and are represented by rail accelerators, ablative Teflon thrusters, MPD arcs, Free Radicals, etc. Endogenous systems such as metallic hydrogen offer great promise and are also being pursued.

Design, fabrication and test of the RL10 derivative II chamber/primary nozzle

NASA Technical Reports Server (NTRS)

Marable, R. W.

1989-01-01

The design, fabrication and test of the RL10-II chamber/primary nozzle was accomplished as part of the RL10 Product Improvement Program (PIP). The overall goal of the RL10 PIP was to gain the knowledge and experience necessary to develop new cryogenic upper stage engines to fulfill future NASA requirements. The goal would be reached by producing an RL10 engine designed to be reusable, operate at several thrust levels, and have increased performance. The goals for the chamber/primary nozzle task were: (1) to design a reusable assembly capable of operation at increased mixture ratio and low thrust; (2) to fabricate three assemblies using new or updated techniques where possible; and (3) to test one assembly to verify the design and construction. The design and fabrication phases produced an assembly having improved features such as single piece reinforcing band segments (i.e., Mae West segments) and relocated tube exit braze joints (i.e., hooked tube exit). In addition, a computer program was developed to design the chamber tubes to meet both performance and heat transfer requirements. The test phase showed the specific impulse of the test bed engine system to be as predicted. These results, along with the heat transfer data obtained, sufficiently proved the overall design of the RL10-II recontoured and shortened chamber/primary nozzle assembly.

Evaluation of propellent tank insulation concepts for low-thrust chemical propulsion systems: Executive summary

NASA Technical Reports Server (NTRS)

Kramer, T.; Brogren, E.; Siegel, B.

1984-01-01

Cryogenic propellant tank insulations or liquid oxygen/liquid hydrogen low-thrust 2224N (500 lbf) propulsion systems (LTPS) were assessed. The insulation studied consisted of combinations of N2-purged foam and multilayer insulation (MLI) as well as He-purged MLI-only. Heat leak and payload performance predictions were made for three shuttle-launched LTPS designed for shuttle bay packaged payload densities of 56 kg cu/m (3.5 lbm/cu ft), 40 kg/cu m (2.5 lbm/cu ft) and 24 kg/cu m (1.5 lbm/cu ft). Foam/MLI insulations were found to increase LTPS payload delivery capability when compared with He-purged MLI-only. An additional benefit of foam/MLI was reduced operational complexity because orbiter cargo bay N2 purge gas could be used for MLI purging. Maximum payload mass benefit occurred when an enhanced convection, rather than natural convection, heat transfer was specified for the insulation purge enclosure. The enhanced convection environment allowed minimum insulation thickness to be used for the foam/MLI interface temperature selected to correspond to the moisture dew point in the N2 purge gas. Experimental verification of foam/MLI benefits was recommended. A conservative program cost estimate for testing a MLI-foam insulated tank was 2.1 million dollars. This cost could be reduced significantly without increasing program risk.

Transpiration cooled throat for hydrocarbon rocket engines

NASA Technical Reports Server (NTRS)

May, Lee R.; Burkhardt, Wendel M.

1991-01-01

The objective for the Transpiration Cooled Throat for Hydrocarbon Rocket Engines Program was to characterize the use of hydrocarbon fuels as transpiration coolants for rocket nozzle throats. The hydrocarbon fuels investigated in this program were RP-1 and methane. To adequately characterize the above transpiration coolants, a program was planned which would (1) predict engine system performance and life enhancements due to transpiration cooling of the throat region using analytical models, anchored with available data; (2) a versatile transpiration cooled subscale rocket thrust chamber was designed and fabricated; (3) the subscale thrust chamber was tested over a limited range of conditions, e.g., coolant type, chamber pressure, transpiration cooled length, and coolant flow rate; and (4) detailed data analyses were conducted to determine the relationship between the key performance and life enhancement variables.

Analysis of cryogenic propellant behavior in microgravity and low thrust environments

NASA Technical Reports Server (NTRS)

Fisher, Mark F.; Schmidt, George R.; Martin, James J.

1991-01-01

Predictions of a CFD program calculating a fluid-free surface shape and motion as a function of imposed acceleration are validated against the drop-tower test data collected to support design and performance assessments of the Saturn S-IVB stage liquid-hydrogen tank. The drop-tower facility, experimental package, and experiment procedures are outlined, and the program is described. It is noted that the validation analysis confirms the program's suitability for predicting low-g fluid slosh behavior, and that a similar analysis could examine the effect of incorporating baffles and screens to impede initiation of any unwanted side loads due to slosh. It is concluded that in actual vehicle applications, the engine thrust tailoff profile should be included in computer simulations if the precise interface versus time definition is needed.

Program documentation. Program description and user information for the hydraulics/auxiliary power unit (HYDRA) computer program. [for the space shuttle orbiter

NASA Technical Reports Server (NTRS)

Redwine, W. J.

1979-01-01

A timeline containing altitude, control surface deflection rates and angles, hinge moment loads, thrust vector control gimbal rates, and main throttle settings is used to derive the model. The timeline is constructed from the output of one or more trajectory simulation programs.

The Lewis thrust fault and related structures in the Disturbed Belt, northwestern Montana

USGS Publications Warehouse

Mudge, Melville Rhodes; Earhart, Robert L.

1980-01-01

The classical Lewis thrust fault in Glacier National Park has now been mapped 125 km south of the park to Steamboat Mountain, where the trace dies out in folded middle Paleozoic rocks. The known length of the fault is 452 km, extending northward from Steamboat Mountain to a point 225 km into Canada, where the fault also dies out in Paleozoic rocks. At the south end, the surface expression of the Lewis thrust begins in a shear zone in folded Mississippian rocks. To the north, the thrust progressively cuts downsection into Proterozoic Y (Belt) rocks near Glacier National Park. Displacement on the Lewis plate increases northward from approximately 3 km on an easterly trending hinge line at the West Fork of the Sun River to a postulated 65 km at the southern edge of the park, where the stratigraphic throw is about 6,500 m. Present data indicate the thrust formed during very late Paleocene to very early Eocene time. The Lewis thrust and related structures, the Hoadley thrust and the Continental Divide syncline, probably formed concurrently under the same stress field. The northern limit of the trace of the Hoadley thrust is within the lower portion of the Lewis plate, about 28 km north of where the Lewis thrust develops, and the Hoadley extends for at least 125 km to the south. Displacement of the Hoadley increases southward from about 1 km at the hinge line to an inferred 70 km near its known southern extent. If our inference is correct, the Hoadley is nearly the southern mirror image of the Lewis to the north. The Continental Divide syncline, a doubly plunging, broad, northerly trending open fold that is about 120 km long, is a major fold within the Lewis plate.

Estimation of wing nonlinear aerodynamic characteristics at supersonic speeds

NASA Technical Reports Server (NTRS)

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

1980-01-01

A computational system for estimation of nonlinear aerodynamic characteristics of wings at supersonic speeds was developed and was incorporated in a computer program. This corrected linearized theory method accounts for nonlinearities in the variation of basic pressure loadings with local surface slopes, predicts the degree of attainment of theoretical leading edge thrust, and provides an estimate of detached leading edge vortex loadings that result when the theoretical thrust forces are not fully realized.

Saturn Apollo Program

NASA Image and Video Library

1967-01-01

Workmen secure a J-2 engine onto the S-IVB (second) stage thrust structure. As part of Marshall Space Center's "building block" approach to the Saturn development, the S-IVB was utilized in the Saturn IBC launch vehicle as a second stage and the Saturn V launch vehicle as a third stage. The booster, built for NASA by McDornell Douglas Corporation, was powered by a single J-2 engine, initially capable of 200,000 pounds of thrust.

Saturn Apollo Program

NASA Image and Video Library

1969-01-01

In the clustering procedure, an initial assembly step for the Saturn IB launch vehicle's S-IB (first) stage, workers at the Michoud Assembly Facility (MAF) near New Orleans, Louisiana, position the central liquid-oxygen tank. Developed by the Marshall Space Flight Center and built by the Chrysler Corporation at Michoud Assembly Facility (MAF), the S-IB utilized eight H-1 engines and each produced 200,000 pounds of thrust, a combined thrust of 1,600,000 pounds.

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.

Automated Assistance for Designing Active Magnetic Bearings

NASA Technical Reports Server (NTRS)

Imlach, Joseph

2008-01-01

MagBear12 is a computer code that assists in the design of radial, heteropolar active magnetic bearings (AMBs). MagBear12 was developed to help in designing the system described in "Advanced Active-Magnetic-Bearing Thrust-Measurement System". Beyond this initial application, MagBear12 is expected to be useful for designing AMBs for a variety of rotating machinery. This program incorporates design rules and governing equations that are also implemented in other, proprietary design software used by AMB manufacturers. In addition, this program incorporates an advanced unpublished fringing-magnetic-field model that increases accuracy beyond that offered by the other AMB-design software.

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.

Computational Electronics and Electromagnetics

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

DeFord, J.F.

The Computational Electronics and Electromagnetics thrust area is a focal point for computer modeling activities in electronics and electromagnetics in the Electronics Engineering Department of Lawrence Livermore National Laboratory (LLNL). Traditionally, they have focused their efforts in technical areas of importance to existing and developing LLNL programs, and this continues to form the basis for much of their research. A relatively new and increasingly important emphasis for the thrust area is the formation of partnerships with industry and the application of their simulation technology and expertise to the solution of problems faced by industry. The activities of the thrust areamore » fall into three broad categories: (1) the development of theoretical and computational models of electronic and electromagnetic phenomena, (2) the development of useful and robust software tools based on these models, and (3) the application of these tools to programmatic and industrial problems. In FY-92, they worked on projects in all of the areas outlined above. The object of their work on numerical electromagnetic algorithms continues to be the improvement of time-domain algorithms for electromagnetic simulation on unstructured conforming grids. The thrust area is also investigating various technologies for conforming-grid mesh generation to simplify the application of their advanced field solvers to design problems involving complicated geometries. They are developing a major code suite based on the three-dimensional (3-D), conforming-grid, time-domain code DSI3D. They continue to maintain and distribute the 3-D, finite-difference time-domain (FDTD) code TSAR, which is installed at several dozen university, government, and industry sites.« less

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.

Modeling and Analysis of the Static Characteristics and Dynamic Responses of Herringbone-grooved Thrust Bearings

NASA Astrophysics Data System (ADS)

Yu, Yunluo; Pu, Guang; Jiang, Kyle

2017-12-01

This paper describes a theoretical investigation of static and dynamic characteristics of herringbone-grooved air thrust bearings. Firstly, Finite Difference Method (FDM) and Finite Volume Method (FVM) are used in combination to solve the non-linear Reynolds equation and to find the pressure distribution of the film and the total loading capacity of the bearing. The influence of design parameters on air film gap characteristics, including the air film thickness, depth of the groove and rotating speed, are analyzed based on the FDM model. The simulation results show that hydrostatic thrust bearings can achieve a better load capacity with less air consumption than herringbone grooved thrust bearings at low compressibility number; herringbone grooved thrust bearings can achieve a higher load capacity but with more air consumption than hydrostatic thrust bearing at high compressibility number; herringbone grooved thrust bearings would lose stability at high rotating speeds, and the stability increases with the depth of the grooves.

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.

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.

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.

Project implementation plan : variable dynamic testbed vehicle

DOT National Transportation Integrated Search

1997-02-01

This document is the project implementation plan for the Variable Dynamic Testbed Vehicle (VDTV) program, sponsored by the Jet Propulsion Laboratory for the Office of Crash Avoidance Research (OCAR) programs in support of Thrust One of the National H...

Preliminary Analysis of Low-Thrust Gravity Assist Trajectories by An Inverse Method and a Global Optimization Technique.

NASA Astrophysics Data System (ADS)

de Pascale, P.; Vasile, M.; Casotto, S.

The design of interplanetary trajectories requires the solution of an optimization problem, which has been traditionally solved by resorting to various local optimization techniques. All such approaches, apart from the specific method employed (direct or indirect), require an initial guess, which deeply influences the convergence to the optimal solution. The recent developments in low-thrust propulsion have widened the perspectives of exploration of the Solar System, while they have at the same time increased the difficulty related to the trajectory design process. Continuous thrust transfers, typically characterized by multiple spiraling arcs, have a broad number of design parameters and thanks to the flexibility offered by such engines, they typically turn out to be characterized by a multi-modal domain, with a consequent larger number of optimal solutions. Thus the definition of the first guesses is even more challenging, particularly for a broad search over the design parameters, and it requires an extensive investigation of the domain in order to locate the largest number of optimal candidate solutions and possibly the global optimal one. In this paper a tool for the preliminary definition of interplanetary transfers with coast-thrust arcs and multiple swing-bys is presented. Such goal is achieved combining a novel methodology for the description of low-thrust arcs, with a global optimization algorithm based on a hybridization of an evolutionary step and a deterministic step. Low thrust arcs are described in a 3D model in order to account the beneficial effects of low-thrust propulsion for a change of inclination, resorting to a new methodology based on an inverse method. The two-point boundary values problem (TPBVP) associated with a thrust arc is solved by imposing a proper parameterized evolution of the orbital parameters, by which, the acceleration required to follow the given trajectory with respect to the constraints set is obtained simply through algebraic computation. By this method a low-thrust transfer satisfying the boundary conditions on position and velocity can be quickly assessed, with low computational effort since no numerical propagation is required. The hybrid global optimization algorithm is made of a double step. Through the evolutionary search a large number of optima, and eventually the global one, are located, while the deterministic step consists of a branching process that exhaustively partitions the domain in order to have an extensive characterization of such a complex space of solutions. Furthermore, the approach implements a novel direct constraint-handling technique allowing the treatment of mixed-integer nonlinear programming problems (MINLP) typical of multiple swingby trajectories. A low-thrust transfer to Mars is studied as a test bed for the low-thrust model, thus presenting the main characteristics of the different shapes proposed and the features of the possible sub-arcs segmentations between two planets with respect to different objective functions: minimum time and minimum fuel consumption transfers. Other various test cases are also shown and further optimized, proving the effective capability of the proposed tool.

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.

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.

Space shuttle maneuvering engine reusable thrust chamber program. Task 11: Low epsilon stability test report data dump

NASA Technical Reports Server (NTRS)

Pauckert, R. P.

1974-01-01

The stability characteristics of the like-doublet injector were defined over the range of OME chamber pressures and mixture ratios. This was accomplished by bomb testing the injector and cavity configurations in solid wall thrust chamber hardware typical of a flight contour with fuel heated to regenerative chamber outlet temperatures. It was found that stability in the 2600-2800 Hz region depends upon injector hydraulics and on chamber acoustics.

Novel, Post-Stall, Thrust-Vectored F-15 RPVs: Laboratory and Flight Tests

DTIC Science & Technology

1990-04-24

Flight Tests Program Manager : Douglas Bowers 1ST-Year Report Principal Investigator: Benjamin 6al-Or April 24, 1990 DTIC.LECTE AUG201990 i/ E...constructed. The geometry, dimensions and preliminary wind-tunnel test data for such a design are provided In Appendix A. If funded, such a 3rd...Preliminary Calibration Flight Test Data Obtained from the Onboard Computer ........ 33 Talless, PST-RaNPAS, Roll-Yaw-Pitch, Thrust-Vectored, PST F-15 (Cf. ADp

Comparative kinematics of the forelimb during swimming in red-eared slider (Trachemys scripta) and spiny softshell (Apalone spinifera) turtles.

PubMed

Pace, C M; Blob, R W; Westneat, M W

2001-10-01

Softshell turtles (Family Trionychidae) possess extensive webbing between the digits of the manus, suggesting that the forelimb may serve as an effective thrust generator during aquatic locomotion. However, the hindlimb has previously been viewed as the dominant propulsive organ in swimming freshwater turtles. To evaluate the potential role of the forelimb in thrust production during swimming in freshwater turtles, we compared the forelimb morphology and three-dimensional forelimb kinematics of a highly aquatic trionychid turtle, the spiny softshell Apalone spinifera, and a morphologically generalized emydid turtle, the red-eared slider Trachemys scripta. Spiny softshells possess nearly twice as much forelimb surface area as sliders for generating drag-based thrust. In addition, although both species use drag-based propulsion, several aspects of forelimb kinematics differ significantly between these species. During the thrust phase of the forelimb cycle, spiny softshells hold the elbow and wrist joints significantly straighter than sliders, thereby further increasing the surface area of the limb that can move water posteriorly and increasing the velocity of the distal portion of the forelimb. These aspects of swimming kinematics in softshells should increase forelimb thrust production and suggest that the forelimbs make more substantial contributions to forward thrust in softshell turtles than in sliders. Spiny softshells also restrict forelimb movements to a much narrower dorsoventral and anteroposterior range than sliders throughout the stroke, thereby helping to minimize limb movements potentially extraneous to forward thrust production. These comparisons demonstrate considerable diversity in the forelimb kinematics of turtles that swim using rowing motions of the limbs and suggest that the evolution of turtle forelimb mechanics produced a variety of contrasting solutions for aquatic specialization.

Advanced Subsonic Technology (AST) Separate-Flow High-Bypass Ratio Nozzle Noise Reduction Program Test Report

NASA Technical Reports Server (NTRS)

Low, John K. C.; Schweiger, Paul S.; Premo, John W.; Barber, Thomas J.; Saiyed, Naseem (Technical Monitor)

2000-01-01

NASA s model-scale nozzle noise tests show that it is possible to achieve a 3 EPNdB jet noise reduction with inwardfacing chevrons and flipper-tabs installed on the primary nozzle and fan nozzle chevrons. These chevrons and tabs are simple devices and are easy to be incorporated into existing short duct separate-flow nonmixed nozzle exhaust systems. However, these devices are expected to cause some small amount of thrust loss relative to the axisymmetric baseline nozzle system. Thus, it is important to have these devices further tested in a calibrated nozzle performance test facility to quantify the thrust performances of these devices. The choice of chevrons or tabs for jet noise suppression would most likely be based on the results of thrust loss performance tests to be conducted by Aero System Engineering (ASE) Inc. It is anticipated that the most promising concepts identified from this program will be validated in full scale engine tests at both Pratt & Whitney and Allied-Signal, under funding from NASA s Engine Validation of Noise Reduction Concepts (EVNRC) programs. This will bring the technology readiness level to the point where the jet noise suppression concepts could be incorporated with high confidence into either new or existing turbofan engines having short-duct, separate-flow nacelles.

Rocketdyne - J-2 Saturn V 2nd and 3rd Stage Engine. Chapter 2, Appendix D

NASA Technical Reports Server (NTRS)

Coffman, Paul

2009-01-01

The J-2 engine was unique in many respects. Technology was not nearly as well-developed in oxygen/hydrogen engines at the start of the J-2 project. As a result, it experienced a number of "teething" problems. It was used in two stages on the Saturn V vehicle in the Apollo Program, as well as on the later Skylab and Apollo/Soyuz programs. In the Apollo Program, it was used on the S-II stage, which was the second stage of the Saturn V vehicle. There were five J-2 engines at the back end of the S-II Stage. In the S-IV-B stage, it was a single engine, but that single engine had to restart. The Apollo mission called for the entire vehicle to reach orbital velocity in low Earth orbit after the first firing of the Saturn-IV-B stage and, subsequently, to fire a second time to go on to the moon. The engine had to be man-rated (worthy of transporting humans). It had to have a high thrust rate and performance associated with oxygen/hydrogen engines, although there were some compromises there. It had to gimbal for thrust vector control. It was an open-cycle gas generator engine delivering up to 230,000 pounds of thrust.

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.

Optimization of transfer trajectories to the Apophis asteroid for spacecraft with high and low thrust

NASA Astrophysics Data System (ADS)

Ivashkin, V. V.; Krylov, I. V.

2014-03-01

The problem of optimization of a spacecraft transfer to the Apophis asteroid is investigated. The scheme of transfer under analysis includes a geocentric stage of boosting the spacecraft with high thrust, a heliocentric stage of control by a low thrust engine, and a stage of deceleration with injection to an orbit of the asteroid's satellite. In doing this, the problem of optimal control is solved for cases of ideal and piecewise-constant low thrust, and the optimal magnitude and direction of spacecraft's hyperbolic velocity "at infinity" during departure from the Earth are determined. The spacecraft trajectories are found based on a specially developed comprehensive method of optimization. This method combines the method of dynamic programming at the first stage of analysis and the Pontryagin maximum principle at the concluding stage, together with the parameter continuation method. The estimates are obtained for the spacecraft's final mass and for the payload mass that can be delivered to the asteroid using the Soyuz-Fregat carrier launcher.

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.

Drone noise

NASA Astrophysics Data System (ADS)

Tinney, Charles; Sirohi, Jayant; University of Texas at Austin Team

2017-11-01

A basic understanding of the noise produced by single and multirotor drones operating at static thrust conditions is presented. This work acts as an extension to previous efforts conducted at The University of Texas at Austin (Tinney et al. 2017, AHS Forum 73). Propeller diameters ranging from 8 inch to 12 inch are examined for configurations comprising an isolated rotor, a quadcopter configuration and a hexacopter configuration, and with a constant drone pitch of 2.25. An azimuthal array of half-inch microphones, placed between 2 and 3 hub-center diameters from the drone center, are used to assess the acoustic near-field. Thrust levels, acquired using a six degree-of-freedom load cell, are then used to correlate acoustic noise levels to aerodynamic performance for each drone configuration. The findings reveal a nearly logarithmic increase in noise with increasing thrust. However, for the same thrust condition, considerable noise reduction is achieved by increasing the number of propeller blades thereby reducing the blade passage frequency and both the thickness and loading noise sources that accompany it.

Noise, Turbulence, and Thrust of Subsonic Free Jets from Lobed Nozzles

NASA Technical Reports Server (NTRS)

Zaman, K. B. M. Q.; Wang, F. Y.

2002-01-01

A study of noise benefit, vis-a-vis thrust penalty, and its correlation to turbulence intensities was conducted for free jets issuing from lobed nozzles. Four convergent nozzles with constant exit area were used in the experiments. Three of these were of rectangular lobed configuration having six, ten and fourteen lobes; the fourth was a circular nozzle. Increasing the number of lobes resulted in a progressive reduction in the turbulence intensities as well as in the overall radiated noise. The noise reduction was pronounced at the low frequency end of the spectrum. However, there was an increase in the high frequency noise that rendered the overall benefit less attractive when compared on a scaled-up A-weighted basis. A reduction in noise was accompanied by a commensurate reduction in the turbulent kinetic energy in the flow field. As expected, increasing the number of lobes involved progressive reduction in the thrust coefficient. Among the cases studied, the six-lobed nozzle had the optimum reduction in turbulence and noise with the least thrust penalty.

A strategy for developing future academic leaders for South Africa in a resource-constrained environment.

PubMed

Lalloo, Umesh G; Bobat, Raziya A; Pillay, Sandy; Wassenaar, Douglas

2014-08-01

A key challenge in addressing the shortage of health care workers in resource-constrained environments is ensuring that there is optimal academic capacity for their training. South Africa's University of KwaZulu-Natal has placed academic and research capacity building at the heart of its program with the Medical Education Partnership Initiative in a program called ENhancing Training and REsearch capacity and Expertise (ENTREE). The program aims to increase the quantity, quality, and retention of health care graduates. It is premised on the basis that research capacity development will lead to an increase in teachers who will be essential to improving the quality and quantity of health care workers needed to meet South Africa's health challenges. This is being achieved through four components of the program: (1) infusion of the undergraduate program with research modules; (2) attraction of academically talented students in the middle of their undergraduate program into a parallel track that has research capacity as its major thrust; (3) attraction of qualified health care personnel into a supported PhD program; and (4) providing strong research ethics training and mentorship. A significant proportion of the program is being executed in rural training sites, to increase the probability that trainees will return to the sites as mentors.

Fault-related fluid flow, Beech Mountain thrust sheet, Blue Ridge Province, Tennessee-North Carolina

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

Waggoner, W.K.; Mora, C.I.

1992-01-01

The latest proterozoic Beech Granite is contained within the Beech Mountain thrust sheet (BMTS), part of a middle-late Paleozoic thrust complex located between Mountain City and Grandfather Mountain windows in the western Blue Ridge of TN-NC. At the base of the BMTS, Beech Granite is juxtaposed against lower Paleozoic carbonate and elastics of the Rome Fm. along the Stone Mountain thrust on the southeaster margin of the Mountain City window. At the top of the BMTS, Beech Granite occurs adjacent to Precambrian mafic rocks of the Pumpkin Patch thrust sheet (PPTS). The Beech Granite is foliated throughout the BMTS withmore » mylonitization and localized cataclasis occurring within thrust zones along the upper and lower margins of the BMTS. Although the degree of mylonitization and cataclasis increases towards the thrusts, blocks of relatively undeformed granite also occur within these fault zones. Mylonites and thrusts are recognized as conduits for fluid movement, but the origin of the fluids and magnitude and effects of fluid migration are not well constrained. This study was undertaken to characterize fluid-rock interaction within the Beech Granite and BMTS. Extensive mobility of some elements/compounds within the thrust zones, and the isotopic and mineralogical differences between the thrust zones and interior of the BMTS indicate that fluid flow was focused within the thrust zones. The wide range of elevated temperatures (400--710 C) indicated by qz-fsp fractionations suggest isotopic disequilibrium. Using a more likely temperature range of 300--400 C for Alleghanian deformation, calculated fluid compositions indicate interactions with a mixture of meteoric-hydrothermal and metamorphic water with delta O-18 = 2.6--7.5[per thousand] for the upper thrust zone and 1.3 to 6.2[per thousand] for the lower thrust zone. These ranges are similar to isotopic data reported for other Blue Ridge thrusts and may represent later periods of meteoric water influx.« less

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.

HILTOP supplement: Heliocentric interplanetary low thrust trajectory optimization program, supplement 1

NASA Technical Reports Server (NTRS)

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

1974-01-01

Modifications and improvements are described that were made to the HILTOP electric propulsion trajectory optimization computer program during calendar years 1973 and 1974. New program features include the simulation of power degradation, housekeeping power, launch asymptote declination optimization, and powered and unpowered ballistic multiple swingby missions with an optional deep space burn.

The Yellow Pages for Rural Development in New Mexico.

ERIC Educational Resources Information Center

Pino, Ricardo, Comp.

Designed to inform the public of rural development programs and services in the State of New Mexico, this directory presents over 100 program summaries, providing mailing addresses and telephone numbers and naming the agency, the specific program thrust, the Officer in Charge, and other contacts. Both an alphabetical and functional index are…

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

Control research in the NASA high-alpha technology program

NASA Technical Reports Server (NTRS)

Gilbert, William P.; Nguyen, Luat T.; Gera, Joseph

1990-01-01

NASA is conducting a focused technology program, known as the High-Angle-of-Attack Technology Program, to accelerate the development of flight-validated technology applicable to the design of fighters with superior stall and post-stall characteristics and agility. A carefully integrated effort is underway combining wind tunnel testing, analytical predictions, piloted simulation, and full-scale flight research. A modified F-18 aircraft has been extensively instrumented for use as the NASA High-Angle-of-Attack Research Vehicle used for flight verification of new methods and concepts. This program stresses the importance of providing improved aircraft control capabilities both by powered control (such as thrust-vectoring) and by innovative aerodynamic control concepts. The program is accomplishing extensive coordinated ground and flight testing to assess and improve available experimental and analytical methods and to develop new concepts for enhanced aerodynamics and for effective control, guidance, and cockpit displays essential for effective pilot utilization of the increased agility provided.

Varus Thrust and Knee Frontal Plane Dynamic Motion in Persons with Knee Osteoarthritis

PubMed Central

Chang, Alison H.; Chmiel, Joan S.; Moisio, Kirsten C.; Almagor, Orit; Zhang, Yunhui; Cahue, September; Sharma, Leena

2013-01-01

Objective Varus thrust visualized during walking is associated with a greater medial knee and an increased risk of medial knee osteoarthritis (OA) progression. Little is known about varus thrust presence determined by visual observation relates to quantitative gait kinematic We hypothesized that varus thrust presence is associated with greater knee frontal plane dynamic movement during the stance phase of gait. Methods Participants had knee OA in at least one knee. Trained examiners assessed participants for varus thrust presence during ambulation. Frontal plane knee motion during ambulation captured using external passive reflective markers and an 8-camera motion analysis system. To examine the cross-sectional relationship between varus thrust and frontal plane knee motion, used multivariable regression models with the quantitative motion measures as dependent variables and varus thrust (present/absent) as predictor; models were adjusted for age, gender, BMI, gait speed, and knee static alignment. Results 236 persons [mean BMI: 28.5 kg/m2 (SD 5.5), mean age: 64.9 years (SD 10.4), 75.8% women] contributing 440 knees comprised the study sample. 82 knees (18.6%) had definite varus thrust. Knees with varus thrust had greater peak varus angle and greater peak varus angular velocity during stance than knees without varus thrust (mean differences 0.90° and 6.65°/sec, respectively). These patterns remained significant after adjusting for age, gender, BMI, gait speed, and knee static alignment. Conclusion Visualized varus thrust during walking was associated with a greater peak knee varus angular velocity and a greater peak knee varus angle during stance phase of gait. PMID:23948980

Numerical Modelling and Analysis of Hydrostatic Thrust Air Bearings for High Loading Capacities and Low Air Consumption

NASA Astrophysics Data System (ADS)

Yu, Yunluo; Pu, Guang; Jiang, Kyle

2017-12-01

The paper presents a numerical simulation study on hydrostatic thrust air bearings to assess the load capacity, compressed air consumptions, and the dynamic response. Finite Difference Method (FDM) and Finite Volume Method (FVM) are combined to solve the non-linear Reynolds equation to find the pressure distribution of the air bearing gas film and the total loading capacity of the bearing. The influence of design parameters on air film gap characteristics, including the air film thickness, supplied pressure, depth of the groove and external load, are investigated based on the proposed FDM model. The simulation results show that the thrust air bearings with a groove have a higher load capacity and air consumption than without a groove, and the load capacity and air consumption both increase with the depth of the groove. Bearings without the groove are better damped than those with the grooves, and the stability of thrust bearing decreases when the groove depth increases. The stability of the thrust bearings is also affected by their loading.

Stress triggering in thrust and subduction earthquakes and stress interaction between the southern San Andreas and nearby thrust and strike-slip faults

USGS Publications Warehouse

Lin, J.; Stein, R.S.

2004-01-01

We argue that key features of thrust earthquake triggering, inhibition, and clustering can be explained by Coulomb stress changes, which we illustrate by a suite of representative models and by detailed examples. Whereas slip on surface-cutting thrust faults drops the stress in most of the adjacent crust, slip on blind thrust faults increases the stress on some nearby zones, particularly above the source fault. Blind thrusts can thus trigger slip on secondary faults at shallow depth and typically produce broadly distributed aftershocks. Short thrust ruptures are particularly efficient at triggering earthquakes of similar size on adjacent thrust faults. We calculate that during a progressive thrust sequence in central California the 1983 Mw = 6.7 Coalinga earthquake brought the subsequent 1983 Mw = 6.0 Nunez and 1985 Mw = 6.0 Kettleman Hills ruptures 10 bars and 1 bar closer to Coulomb failure. The idealized stress change calculations also reconcile the distribution of seismicity accompanying large subduction events, in agreement with findings of prior investigations. Subduction zone ruptures are calculated to promote normal faulting events in the outer rise and to promote thrust-faulting events on the periphery of the seismic rupture and its downdip extension. These features are evident in aftershocks of the 1957 Mw = 9.1 Aleutian and other large subduction earthquakes. We further examine stress changes on the rupture surface imparted by the 1960 Mw = 9.5 and 1995 Mw = 8.1 Chile earthquakes, for which detailed slip models are available. Calculated Coulomb stress increases of 2-20 bars correspond closely to sites of aftershocks and postseismic slip, whereas aftershocks are absent where the stress drops by more than 10 bars. We also argue that slip on major strike-slip systems modulates the stress acting on nearby thrust and strike-slip faults. We calculate that the 1857 Mw = 7.9 Fort Tejon earthquake on the San Andreas fault and subsequent interseismic slip brought the Coalinga fault ???1 bar closer to failure but inhibited failure elsewhere on the Coast Ranges thrust faults. The 1857 earthquake also promoted failure on the White Wolf reverse fault by 8 bars, which ruptured in the 1952 Mw = 7.3 Kern County shock but inhibited slip on the left-lateral Garlock fault, which has not ruptured since 1857. We thus contend that stress transfer exerts a control on the seismicity of thrust faults across a broad spectrum of spatial and temporal scales. Copyright 2004 by the American Geophysical Union.

Effect of applied magnetic nozzle on an MPD Thruster

NASA Astrophysics Data System (ADS)

Ando, Akira; Izawa, Yuki; Okawa, Kohei; Hashima, Yoko; Watanabe, Hiroshi; Tanaka, Nozomi

2012-10-01

Electric propulsion systems are suitable for long-term mission in space due to its higher specific impulse. An Magneto-Plasma-Dynamic Thruster (MPDT) is one of the promising thrusters of high power electric propulsion systems. It has been reported that the thrust performance of an MPDT can be improved by applying an axial magnetic field on it. In order to investigate the effect of applied field on an MPDT, we have investigated plume plasma parameters and thrust performance in an applied field MPDT. Different types of divergent magnetic nozzle were applied to an MPDT, and thrust was measured using a pendulum type thrust target. Experiments were performed with hydrogen, helium, and argon as propellant gas. Thrust increased with a discharge current up to 6kA and applied magnetic field up to 0.4T. Maximum thrust of 7N was obtained when the peak position of the applied magnetic field was set upstream of the muzzle of the MPDT. The highest thrust performance was obtained with hydrogen gas with divergent magnetic nozzle applied to the MPDT.

Water, energy, and biogeochemical budgets investigation at Panola Mountain research watershed, Stockbridge, Georgia; a research plan

USGS Publications Warehouse

Huntington, T.G.; Hooper, R.P.; Peters, N.E.; Bullen, T.D.; Kendall, Carol

1993-01-01

The Panola Mountain Research Watershed (PMRW), located in the Panola Mountain State Conservation Park near Stockbridge, Georgia has been selected as a core research watershed under the Water, Energy and Biogeochemical Budgets (WEBB) research initiative of the U.S. Geological Survey (USGS) Global Climate Change Program. This research plan describes ongoing and planned research activities at PMRW from 1984 to 1994. Since 1984, PMRW has been studied as a geochemical process research site under the U.S. Acid Precipitation Thrust Program. Research conducted under this Thrust Program focused on the estimation of dry atmospheric deposition, short-term temporal variability of streamwater chemistry, sulfate adsorption characteristics of the soils, groundwater chemistry, throughfall chemistry, and streamwater quality. The Acid Precipitation Thrust Program continues (1993) to support data collection and a water-quality laboratory. Proposed research to be supported by the WEBB program is organized in 3 interrelated categories: streamflow generation and water-quality evolution, weathering and geochemical evolution, and regulation of soil-water chemistry. Proposed research on streamflow generation and water-quality evolution will focus on subsurface water movement, its influence in streamflow generation, and the associated chemical changes of the water that take place along its flowpath. Proposed research on weathering and geochemical evolution will identify the sources of cations observed in the streamwater at Panola Mountain and quantify the changes in cation source during storms. Proposed research on regulation of soil-water chemistry will focus on the poorly understood processes that regulate soil-water and groundwater chemistry. (USGS)

ESEA Title I Evaluation Report: Programs for Educationally Deprived Children. September 1973-August 1974.

ERIC Educational Resources Information Center

Turner, W. E..; And Others

The major thrust of 1965 Elementary Secondary Education Act Title I, or Project SPEEDY (Special Programs to Enhance the Education of Disadvantaged Youth), has been in the area of corrective reading and other programs designed to promote language development. While the emphasis of federal programs has been to concentrate funds on fewer recipients,…

A Strategy for Developing Future Academic Leaders for South Africa in a Resource-Constrained Environment

PubMed Central

Lalloo, Umesh G.; Bobat, Raziya A.; Pillay, Sandy; Wassenaar, Douglas

2014-01-01

A key challenge in addressing the shortage of health care workers in resource-constrained environments is ensuring that there is optimal academic capacity for their training. South Africa’s University of KwaZulu-Natal has placed academic and research capacity building at the heart of its program with the Medical Education Partnership Initiative (MEPI) in a program called ENhancing Training, REsearch Capacity, and Expertise (ENTREE). The program is premised on the basis that research capacity development will lead to an increase in teachers who will be essential to improving the quality and quantity of health care workers needed to meet South Africa’s health challenges. This is being achieved through four components of the program: (1) infusion of the undergraduate program with research modules; (2) attraction of academically talented students in the middle of their undergraduate program into a parallel track that has research capacity as its major thrust; (3) attraction of qualified health care personnel into a supported PhD program; and (4) providing strong research ethics training and mentorship. A significant proportion of the program is being executed in rural training sites, to increase the probability that trainees will return to the sites as mentors. PMID:25072580

Saturn Apollo Program

NASA Image and Video Library

1964-11-01

The Saturn I S-IV stage (second stage) assembly for the SA-9 mission underwent the weight and balance test in the hanger building at Cape Canaveral. The S-IV stage had six RL-10 engines which used liquid hydrogen and liquid oxygen as its propellants arranged in a circle. Each RL-10 engine produced a thrust of 15,000 pounds, a total combined thrust of 90,000 pounds. The SA-9 mission was the first Saturn with operational payload Pegasus I, meteoroid detection satellite, and launched on February 16, 1965.

Saturn Apollo Program

NASA Image and Video Library

1969-01-01

In the clustering procedure, an initial assembly step for the Saturn IB launch vehicle's S-IB (first) stage, workers at the Michoud Assembly Facility (MAF) near New Orleans, Louisiana, place the first of eight outboard fuel tanks atop the central liquid-oxygen tank. Developed by the Marshall Space Flight Center and built by the Chrysler Corporation at Michoud Assembly Facility (MAF), the S-IB utilized eight H-1 engines and each produced 200,000 pounds of thrust, a combined thrust of 1,600,000 pounds.

A-1 Test Stand work

NASA Technical Reports Server (NTRS)

2010-01-01

A structural steel beam to support the new thrust measurement system on the A-1 Test Stand at NASA's John C. Stennis Space Center is lifted to waiting employees for installation. The beam is part of the thrust takeout structure needed to support the new measurement system. Four such beams have been installed at the stand in preparation for installation of the system in upcoming weeks. Operators are preparing the stand for testing the next generation of rocket engines for the U.S. space program.

X-31 Unloading Returning from Paris Air Show

NASA Technical Reports Server (NTRS)

1995-01-01

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

STE thrust chamber technology: Main injector technology program and nozzle Advanced Development Program (ADP)

NASA Technical Reports Server (NTRS)

1993-01-01

The purpose of the STME Main Injector Program was to enhance the technology base for the large-scale main injector-combustor system of oxygen-hydrogen booster engines in the areas of combustion efficiency, chamber heating rates, and combustion stability. The initial task of the Main Injector Program, focused on analysis and theoretical predictions using existing models, was complemented by the design, fabrication, and test at MSFC of a subscale calorimetric, 40,000-pound thrust class, axisymmetric thrust chamber operating at approximately 2,250 psi and a 7:1 expansion ratio. Test results were used to further define combustion stability bounds, combustion efficiency, and heating rates using a large injector scale similar to the Pratt & Whitney (P&W) STME main injector design configuration including the tangential entry swirl coaxial injection elements. The subscale combustion data was used to verify and refine analytical modeling simulation and extend the database range to guide the design of the large-scale system main injector. The subscale injector design incorporated fuel and oxidizer flow area control features which could be varied; this allowed testing of several design points so that the STME conditions could be bracketed. The subscale injector design also incorporated high-reliability and low-cost fabrication techniques such as a one-piece electrical discharged machined (EDMed) interpropellant plate. Both subscale and large-scale injectors incorporated outer row injector elements with scarfed tip features to allow evaluation of reduced heating rates to the combustion chamber.

Eddy Viscosity for Variable Density Coflowing Streams,

DTIC Science & Technology

EDDY CURRENTS, *JET MIXING FLOW, *VISCOSITY, *AIR FLOW, MATHEMATICAL MODELS, INCOMPRESSIBLE FLOW, AXISYMMETRIC FLOW, MATHEMATICAL PREDICTION, THRUST AUGMENTATION , EJECTORS , COMPUTER PROGRAMMING, SECONDARY FLOW, DENSITY, MODIFICATION.

Non-contact thrust stand calibration method for repetitively pulsed electric thrusters.

PubMed

Wong, Andrea R; Toftul, Alexandra; Polzin, Kurt A; Pearson, J Boise

2012-02-01

A thrust stand calibration technique for use in testing repetitively pulsed electric thrusters for in-space propulsion has been developed and tested using a modified hanging pendulum thrust stand. In the implementation of this technique, current pulses are applied to a solenoid to produce a pulsed magnetic field that acts against a permanent magnet mounted to the thrust stand pendulum arm. The force on the magnet is applied in this non-contact manner, with the entire pulsed force transferred to the pendulum arm through a piezoelectric force transducer to provide a time-accurate force measurement. Modeling of the pendulum arm dynamics reveals that after an initial transient in thrust stand motion the quasi-steady average deflection of the thrust stand arm away from the unforced or "zero" position can be related to the average applied force through a simple linear Hooke's law relationship. Modeling demonstrates that this technique is universally applicable except when the pulsing period is increased to the point where it approaches the period of natural thrust stand motion. Calibration data were obtained using a modified hanging pendulum thrust stand previously used for steady-state thrust measurements. Data were obtained for varying impulse bit at constant pulse frequency and for varying pulse frequency. The two data sets exhibit excellent quantitative agreement with each other. The overall error on the linear regression fit used to determine the calibration coefficient was roughly 1%.

On-Board Chemical Propulsion

NASA Technical Reports Server (NTRS)

Schneider, Steven J.

1997-01-01

NASA Lewis Research Center's On-Board Propulsion program (OBP) is developing low-thrust chemical propulsion technologies for both satellite and vehicle reaction control applications. There is a vigorous international competition to develop new, highperformance bipropellant engines. High-leverage bipropellant systems are critical to both commercial competitiveness in the international communications market and to cost-effective mission design in government sectors. To significantly improve bipropellant engine performance, we must increase the thermal margin of the chamber materials. Iridium-coated rhenium (Ir/Re) engines, developed and demonstrated under OBP programs, can operate at temperatures well above the constraints of state-of-practice systems, providing a sufficient margin to maximize performance with the hypergolic propellants used in most satellite propulsion systems.

NASA aviation safety program aircraft engine health management data mining tools roadmap

DOT National Transportation Integrated Search

2000-04-01

Aircraft Engine Health Management Data Mining Tools is a project led by NASA Glenn Research Center in support of the NASA Aviation Safety Program's Aviation System Monitoring and Modeling Thrust. The objective of the Glenn-led effort is to develop en...

Fortran 4 program for two-impulse rendezvous analysis

NASA Technical Reports Server (NTRS)

Barling, W. H., Jr.; Brothers, W. J.; Darling, W. H., Jr.

1967-01-01

Program determines if rendezvous in near space is possible, and performs an analysis to determine the approximate required values of the magnitude and direction of two thrust applications of the upper stage of a rocket firing. The analysis is performed by using ordinary Keplerian mechanics.

Comparison of theoretical and experimental thrust performance of a 1030:1 area ratio rocket nozzle at a chamber pressure of 2413 kN/sq m (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 performance-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 that has been incorporated into the December 1984 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. A detailed description of the test conditions and TDK input parameters is given. The reuslts indicate 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.

Boundary layer simulator improvement

NASA Technical Reports Server (NTRS)

Praharaj, Sarat C.; Schmitz, Craig P.; Nouri, Joseph A.

1989-01-01

Boundary Layer Integral Matrix Procedure (BLIMPJ) has been identified by the propulsion community as the rigorous boundary layer program in connection with the existing JANNAF reference programs. The improvements made to BLIMPJ and described herein have potential applications in the design of the future Orbit Transfer Vehicle engines. The turbulence model is validated to include the effects of wall roughness and a way is devised to treat multiple smooth-rough surfaces. A prediction of relaminarization regions is examined as is the combined effects of wall cooling and surface roughness on relaminarization. A turbulence model to represent the effects of constant condensed phase loading is given. A procedure is described for thrust decrement calculation in thick boundary layers by coupling the T-D Kinetics Program and BLIMPJ and a way is provided for thrust loss optimization. Potential experimental studies in rocket nozzles are identified along with the required instrumentation to provide accurate measurements in support of the presented new analytical models.

Full scale technology demonstration of a modern counterrotating unducted fan engine concept. Engine test

NASA Technical Reports Server (NTRS)

1987-01-01

The Unducted Fan (UDF) engine is an innovative aircraft engine concept based on an ungeared, counterrotating, unducted, ultra-high-bypass turbofan configuration. This engine is being developed to provide a high thrust-to-weight ratio power plant with exceptional fuel efficiency for subsonic aircraft application. This report covers the successful ground testing of this engine. A test program exceeding 100-hr duration was completed, in which all the major goals were achieved. The following accomplishments were demonstrated: (1) full thrust (25,000 lb); (2) full counterrotating rotor speeds (1393+ rpm); (3) low specific fuel consumption (less than 0.24 lb/hr/lb); (4) new composite fan design; (5) counterrotation of structures, turbines, and fan blades; (6) control system; (7) actuation system; and (8) reverse thrust.

The effect of exhaust plume/afterbody interaction on installed Scramjet performance

NASA Technical Reports Server (NTRS)

Edwards, Thomas Alan

1988-01-01

Newly emerging aerospace technology points to the feasibility of sustained hypersonic flight. Designing a propulsion system capable of generating the necessary thrust is now the major obstacle. First-generation vehicles will be driven by air-breathing scramjet (supersonic combustion ramjet) engines. Because of engine size limitations, the exhaust gas leaving the nozzle will be highly underexpanded. Consequently, a significant amount of thrust and lift can be extracted by allowing the exhaust gases to expand along the underbody of the vehicle. Predicting how these forces influence overall vehicle thrust, lift, and moment is essential to a successful design. This work represents an important first step toward that objective. The UWIN code, an upwind, implicit Navier-Stokes computer program, has been applied to hypersonic exhaust plume/afterbody flow fields. The capability to solve entire vehicle geometries at hypersonic speeds, including an interacting exhaust plume, has been demonstrated for the first time. Comparison of the numerical results with available experimental data shows good agreement in all cases investigated. For moderately underexpanded jets, afterbody forces were found to vary linearly with the nozzle exit pressure, and increasing the exit pressure produced additional nose-down pitching moment. Coupling a species continuity equation to the UWIN code enabled calculations indicating that exhaust gases with low isentropic exponents (gamma) contribute larger afterbody forces than high-gamma exhaust gases. Moderately underexpanded jets, which remain attached to unswept afterbodies, underwent streamwise separation on upswept afterbodies. Highly underexpanded jets produced altogether different flow patterns, however. The highly underexpanded jet creates a strong plume shock, and the interaction of this shock with the afterbody was found to produce complicated patterns of crossflow separation. Finally, the effect of thrust vectoring on vehicle balance has been shown to alter dramatically the vehicle pitching moment.

Negative Thrust and Torque Characteristics of an Adjustable-Pitch Metal Propeller

NASA Technical Reports Server (NTRS)

Hartman, Edwin P

1934-01-01

This report presents the results of a series of negative thrust and torque measurements made with a 4 foot diameter model of a conventional aluminum-alloy propeller. The tests were made in the 20-foot propeller-research tunnel of the National Advisory Committee for Aeronautics. The results show that the negative thrust is considerably affected by the shape and size of the body behind the propeller, that the maximum negative thrust increases with decrease in blade-angle setting, and that the drag of a locked propeller may be greatly reduced by feathering it into the wind. Several examples of possible applications of the data are given.

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.

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.

Thermal barrier coatings (TBC's) for high heat flux thrust chambers

NASA Astrophysics Data System (ADS)

Bradley, Christopher M.

The last 30 years materials engineers have been under continual pressure to develop materials with a greater temperature potential or to produce configurations that can be effectively cooled or otherwise protected at elevated temperature conditions. Turbines and thrust chambers produce some of the harshest service conditions for materials which lead to the challenges engineers face in order to increase the efficiencies of current technologies due to the energy crisis that the world is facing. The key tasks for the future of gas turbines are to increase overall efficiencies to meet energy demands of a growing world population and reduce the harmful emissions to protect the environment. Airfoils or blades tend to be the limiting factor when it comes to the performance of the turbine because of their complex design making them difficult to cool as well as limitations of their thermal properties. Key tasks for space transportation it to lower costs while increasing operational efficiency and reliability of our space launchers. The important factor to take into consideration is the rocket nozzle design. The design of the rocket nozzle or thrust chamber has to take into account many constraints including external loads, heat transfer, transients, and the fluid dynamics of expanded hot gases. Turbine engines can have increased efficiencies if the inlet temperature for combustion is higher, increased compressor capacity and lighter weight materials. In order to push for higher temperatures, engineers need to come up with a way to compensate for increased temperatures because material systems that are being used are either at or near their useful properties limit. Before thermal barrier coatings were applied to hot-section components, material alloy systems were able to withstand the service conditions necessary. But, with the increased demand for performance, higher temperatures and pressures have become too much for those alloy systems. Controlled chemistry of hot-section components has become critical, but at the same time the service conditions have put our best alloy systems to their limits. As a result, implementation of cooling holes and thermal barrier coatings are new advances in hot-section technologies now looked at for modifications to reach higher temperature applications. Current thermal barrier coatings used in today's turbine applications is known as 8%yttria-stabilized zirconia (YSZ) and there are no coatings for current thrust chambers. Current research is looking at the applicability of 8%yttria-stabilized hafnia (YSH) for turbine applications and the implementation of 8%YSZ onto thrust chambers. This study intends to determine if the use of thermal barrier coatings are applicable for high heat flux thrust chambers using industrial YSZ will be advantageous for improvements in efficiency, thrust and longer service life by allowing the thrust chambers to be used more than once.

Making Programs Self-Sustaining at a Small Historic House Museum

ERIC Educational Resources Information Center

Alleyne, Shirley Brown

2010-01-01

Museum educators at small museums are sometimes thrust, solo, into building budgets from scratch. In this case study an emerging professional discovers that she has limited funding for an existing program that she was hired to run. She finds that she has to rebuild the budget to discover the program's true costs, and that sustainability depends…

Kinematics and Hydrodynamics of Burst-and-Coast Strategy in Carangiform Swimming

NASA Astrophysics Data System (ADS)

Han, Pan; Dong, Haibo; di Santo, Valentina; Lauder, George

2017-11-01

In this work, burst-and-coast swimming hydrodynamics of a trout is studied using a combined experimental and computational approach. The associated kinematics is reconstructed from the output of a high-speed photogrammetry system. The hydrodynamic performance and wake structures are then investigated using an in-house immersed-boundary-method based flow solver and compared with those found in steady undulatory swimming. Results have shown that the carangiform swimmer uses a completely different trust producing strategy when conducts burst-and-coast swimming. Comparing to steady swimming, the trunk curvature of the fish has increased twofold during the burst phase. As a result, it contributes about 15% of total trust during the swimming. Results have also shown that the thrust produced by the caudal fin has increased by tenfold during burst swimming due to larger flapping amplitude and pitching angle. Vortex dynamics analysis has shown that unlike the steady swimming, vortex rings formed during burst swimming result in a stronger downstream jet, which suggests a new thrust enhancement mechanism in carangiform swimming. This work was funded by the Office of Naval Research under Program Director Dr B. Brizzolara, MURI Grant Number N00014-14-1-0533.

ASRM Multi-Port Igniter Flow Field Analysis

NASA Technical Reports Server (NTRS)

Kania, Lee; Dumas, Catherine; Doran, Denise

1993-01-01

The Advanced Solid Rocket Motor (ASRM) program was initiated by NASA in response to the need for a new generation rocket motor capable of providing increased thrust levels over the existing Redesigned Solid Rocket Motor (RSRM) and thus augment the lifting capacity of the space shuttle orbiter. To achieve these higher thrust levels and improve motor reliability, advanced motor design concepts were employed. In the head end of the motor, for instance, the propellent cast has been changed from the conventional annular configuration to a 'multi-slot' configuration in order to increase the burn surface area and guarantee rapid motor ignition. In addition, the igniter itself has been redesigned and currently features 12 exhaust ports in order to channel hot igniter combustion gases into the circumferential propellent slots. Due to the close proximity of the igniter ports to the propellent surfaces, new concerns over possible propellent deformation and erosive burning have arisen. The following documents the effort undertaken using computational fluid dynamics to perform a flow field analysis in the top end of the ASRM motor to determine flow field properties necessary to permit a subsequent propellent fin deformation analysis due to pressure loading and an assessment of the extent of erosive burning.

Improvement of Thrust Characteristics of Helicon Plasma Thruster using Local Gas Fueling Method

NASA Astrophysics Data System (ADS)

Kuwahara, Daisuke; Amma, Kosuke; Ishigami, Yuichi; Igarashi, Akihiko; Nishimoto, Shinichi; Shinohara, Shunjiro; Miyazawa, Junichi

2017-10-01

A helicon plasma thruster is proposed as a long-lifetime electric thruster which has non-direct contact electrodes. Here, a neutral particle, e.g., H2, Ar, and Xe works, as a fuel gas. In most cases, these gases are supplied into a discharge tube by the use of a simple nozzle. Therefore, the neutral particle fills a discharge tube homogenous. However, there are two problems in this configuration. First, there is a limitation of an electron density increase, due to a neutral particle depletion in the central region of the high-density helicon plasma. This limitation reduces the thrust performance directly. Second, the high-density plasma causes an erosion of an inner discharge tube wall. For the future MW class thruster, this problem will become serious because the particle and heat fluxes of the plasma will increase drastically. To solve above-mentioned problems, we have proposed local fueling methods for the high-density helicon plasma. In this presentation, we will show the methods and experimental results using a fueling tube, inserted in a plasma directly. This work is supported by JSPS KAKENHI Grant Number 16K17843 and NIFS Collaboration Research program (NIFSKBAF016).

ASRM multi-port igniter flow field analysis

NASA Astrophysics Data System (ADS)

Kania, Lee; Dumas, Catherine; Doran, Denise

1993-07-01

The Advanced Solid Rocket Motor (ASRM) program was initiated by NASA in response to the need for a new generation rocket motor capable of providing increased thrust levels over the existing Redesigned Solid Rocket Motor (RSRM) and thus augment the lifting capacity of the space shuttle orbiter. To achieve these higher thrust levels and improve motor reliability, advanced motor design concepts were employed. In the head end of the motor, for instance, the propellent cast has been changed from the conventional annular configuration to a 'multi-slot' configuration in order to increase the burn surface area and guarantee rapid motor ignition. In addition, the igniter itself has been redesigned and currently features 12 exhaust ports in order to channel hot igniter combustion gases into the circumferential propellent slots. Due to the close proximity of the igniter ports to the propellent surfaces, new concerns over possible propellent deformation and erosive burning have arisen. The following documents the effort undertaken using computational fluid dynamics to perform a flow field analysis in the top end of the ASRM motor to determine flow field properties necessary to permit a subsequent propellent fin deformation analysis due to pressure loading and an assessment of the extent of erosive burning.

Varus thrust and knee frontal plane dynamic motion in persons with knee osteoarthritis.

PubMed

Chang, A H; Chmiel, J S; Moisio, K C; Almagor, O; Zhang, Y; Cahue, S; Sharma, L

2013-11-01

Varus thrust visualized during walking is associated with a greater medial knee load and an increased risk of medial knee osteoarthritis (OA) progression. Little is known about how varus thrust presence determined by visual observation relates to quantitative gait kinematic data. We hypothesized that varus thrust presence is associated with greater knee frontal plane dynamic movement during the stance phase of gait. Participants had knee OA in at least one knee. Trained examiners assessed participants for varus thrust presence during ambulation. Frontal plane knee motion during ambulation was captured using external passive reflective markers and an 8-camera motion analysis system. To examine the cross-sectional relationship between varus thrust and frontal plane knee motion, we used multivariable regression models with the quantitative motion measures as dependent variables and varus thrust (present/absent) as predictor; models were adjusted for age, gender, body mass index (BMI), gait speed, and knee static alignment. 236 persons [mean BMI: 28.5 kg/m(2) (standard deviation (SD) 5.5), mean age: 64.9 years (SD 10.4), 75.8% women] contributing 440 knees comprised the study sample. 82 knees (18.6%) had definite varus thrust. Knees with varus thrust had greater peak varus angle and greater peak varus angular velocity during stance than knees without varus thrust (mean differences 0.90° and 6.65°/s, respectively). These patterns remained significant after adjusting for age, gender, BMI, gait speed, and knee static alignment. Visualized varus thrust during walking was associated with a greater peak knee varus angular velocity and a greater peak knee varus angle during stance phase of gait. Copyright © 2013 Osteoarthritis Research Society International. Published by Elsevier Ltd. All rights reserved.

Research Technology

NASA Image and Video Library

2002-03-11

Engineers at the Marshall Space Flight Center (MSFC) have begun a series of engine tests on a new breed of space propulsion: a Reaction Control Engine developed for the Space Launch Initiative (SLI). The engine, developed by TRW Space and Electronics of Redondo Beach, California, is an auxiliary propulsion engine designed to maneuver vehicles in orbit. It is used for docking, reentry, attitude control, and fine-pointing while the vehicle is in orbit. The engine uses nontoxic chemicals as propellants, a feature that creates a safer environment for ground operators, lowers cost, and increases efficiency with less maintenance and quicker turnaround time between missions. Testing includes 30 hot-firings. This photograph shows the first engine test performed at MSFC that includes SLI technology. Another unique feature of the Reaction Control Engine is that it operates at dual thrust modes, combining two engine functions into one engine. The engine operates at both 25 and 1,000 pounds of force, reducing overall propulsion weight and allowing vehicles to easily maneuver in space. The low-level thrust of 25 pounds of force allows the vehicle to fine-point maneuver and dock while the high-level thrust of 1,000 pounds of force is used for reentry, orbit transfer, and coarse positioning. SLI is a NASA-wide research and development program, managed by the MSFC, designed to improve safety, reliability, and cost effectiveness of space travel for second generation reusable launch vehicles.

Reaction Control Engine for Space Launch Initiative

NASA Technical Reports Server (NTRS)

2002-01-01

Engineers at the Marshall Space Flight Center (MSFC) have begun a series of engine tests on a new breed of space propulsion: a Reaction Control Engine developed for the Space Launch Initiative (SLI). The engine, developed by TRW Space and Electronics of Redondo Beach, California, is an auxiliary propulsion engine designed to maneuver vehicles in orbit. It is used for docking, reentry, attitude control, and fine-pointing while the vehicle is in orbit. The engine uses nontoxic chemicals as propellants, a feature that creates a safer environment for ground operators, lowers cost, and increases efficiency with less maintenance and quicker turnaround time between missions. Testing includes 30 hot-firings. This photograph shows the first engine test performed at MSFC that includes SLI technology. Another unique feature of the Reaction Control Engine is that it operates at dual thrust modes, combining two engine functions into one engine. The engine operates at both 25 and 1,000 pounds of force, reducing overall propulsion weight and allowing vehicles to easily maneuver in space. The low-level thrust of 25 pounds of force allows the vehicle to fine-point maneuver and dock while the high-level thrust of 1,000 pounds of force is used for reentry, orbit transfer, and coarse positioning. SLI is a NASA-wide research and development program, managed by the MSFC, designed to improve safety, reliability, and cost effectiveness of space travel for second generation reusable launch vehicles.

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

Busbey, A.B.

Seismic Processing Workshop, a program by Parallel Geosciences of Austin, TX, is discussed in this column. The program is a high-speed, interactive seismic processing and computer analysis system for the Apple Macintosh II family of computers. Also reviewed in this column are three products from Wilkerson Associates of Champaign, IL. SubSide is an interactive program for basin subsidence analysis; MacFault and MacThrustRamp are programs for modeling faults.

Computer program for post-flight evaluation of a launch vehicle upper-stage on-off reaction control system

NASA Technical Reports Server (NTRS)

Knauber, R. N.

1982-01-01

This report describes a FORTRAN IV coded computer program for post-flight evaluation of a launch vehicle upper stage on-off reaction control system. Aerodynamic and thrust misalignment disturbances are computed as well as the total disturbing moments in pitch, yaw, and roll. Effective thrust misalignment angle time histories of the rocket booster motor are calculated. Disturbing moments are integrated and used to estimate the required control system total inpulse. Effective control system specific inpulse is computed for the boost and coast phases using measured control fuel useage. This method has been used for more than fifteen years for analyzing the NASA Scout launch vehicle second and third-stage reaction control system performance. The computer program is set up in FORTRAN IV for a CDC CYBER 175 system. With slight modification it can be used on other machines having a FORTRAN compiler. The program has optional CALCOMP plotting output. With this option the program requires 19K words of memory and has 786 cards. Running time on a CDC CYBER 175 system is less than three (3) seconds for a typical problem.

Varus thrust in women with early medial knee osteoarthritis and its relation with the external knee adduction moment.

PubMed

Mahmoudian, Armaghan; van Dieen, Jaap H; Bruijn, Sjoerd M; Baert, Isabel Ac; Faber, Gert S; Luyten, Frank P; Verschueren, Sabine Mp

2016-11-01

Varus thrust, defined as an abrupt increase of the knee varus angle during weight-bearing in gait, has been shown to be present in patients with moderate to severe knee osteoarthritis and is considered to be one of the risk factors for progression of symptomatic medial knee osteoarthritis. We evaluated the presence and magnitude of varus thrust and its relation with the Knee Adduction Moment in women with early medial knee osteoarthritis, and compared it to that in a group of controls and in a group of subjects with established medial knee osteoarthritis. Twenty-seven women with early medial knee osteoarthritis, 20 women with established medial knee osteoarthritis and 24 asymptomatic controls were evaluated. Varus thrust was estimated as an increase of the knee varus angle during the weight-bearing phase of gait at self-selected speed, assessed by 3D motion analysis. Varus thrust was significantly higher in both early and established osteoarthritis groups compared to the control group (P<0.001), but not different between osteoarthritis groups. While the knee adduction moments were higher than controls only in the established osteoarthritis group, the magnitude of varus thrust was significantly correlated with the second peak knee adduction moment. Higher varus thrust was found both in early and established stages of knee osteoarthritis, suggesting that problems with dynamic stabilization of the knee are present early in the development of knee osteoarthritis. This highlights the necessity of considering dynamic alignment in rehabilitation already in the early stages of the disease. Copyright © 2016 Elsevier Ltd. All rights reserved.

The Role of Modern Control Theory in the Design of Controls for Aircraft Turbine Engines

NASA Technical Reports Server (NTRS)

Zeller, J.; Lehtinen, B.; Merrill, W.

1982-01-01

Accomplishments in applying Modern Control Theory to the design of controls for advanced aircraft turbine engines were reviewed. The results of successful research programs are discussed. Ongoing programs as well as planned or recommended future thrusts are also discussed.

A Program of Behavioral Objectives for the Students.

ERIC Educational Resources Information Center

McDowell, Thomas, Comp.; Sharbaugh, Marilyn, Comp.

Provided is a program of behavioral objectives for use with speech handicapped preschool and elementary level children in the areas of articulation disorders, language disorders, stuttering, tongue thrust, and voice. Communication disorders are defined as substandard behaviors which interfere with the student's ability to perform acceptably for…

Non-Contact Thrust Stand Calibration Method for Repetitively-Pulsed Electric Thrusters

NASA Technical Reports Server (NTRS)

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

2011-01-01

A thrust stand calibration technique for use in testing repetitively-pulsed electric thrusters for in-space propulsion has been developed and tested using a modified hanging pendulum thrust stand. In the implementation of this technique, current pulses are applied to a solenoidal coil to produce a pulsed magnetic field that acts against the magnetic field produced by a permanent magnet mounted to the thrust stand pendulum arm. The force on the magnet is applied in this non-contact manner, with the entire pulsed force transferred to the pendulum arm through a piezoelectric force transducer to provide a time-accurate force measurement. Modeling of the pendulum arm dynamics reveals that after an initial transient in thrust stand motion the quasisteady average deflection of the thrust stand arm away from the unforced or zero position can be related to the average applied force through a simple linear Hooke s law relationship. Modeling demonstrates that this technique is universally applicable except when the pulsing period is increased to the point where it approaches the period of natural thrust stand motion. Calibration data were obtained using a modified hanging pendulum thrust stand previously used for steady-state thrust measurements. Data were obtained for varying impulse bit at constant pulse frequency and for varying pulse frequency. The two data sets exhibit excellent quantitative agreement with each other as the constant relating average deflection and average thrust match within the errors on the linear regression curve fit of the data. Quantitatively, the error on the calibration coefficient is roughly 1% of the coefficient value.

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.

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.

Directed Energy Deflection Laboratory Measurements of Asteroids and Space Debris

NASA Astrophysics Data System (ADS)

Brashears, T.; Lubin, P. M.

2016-12-01

We report on laboratory studies of the effectiveness of directed energy planetary and space defense as a part of the DE-STAR (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" or a space debris 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 60 µN/Wabsorbed as a reasonable lower limit to the thrust per optical watt absorbed.

CCARES: A computer algorithm for the reliability analysis of laminated CMC components

NASA Technical Reports Server (NTRS)

Duffy, Stephen F.; Gyekenyesi, John P.

1993-01-01

Structural components produced from laminated CMC (ceramic matrix composite) materials are being considered for a broad range of aerospace applications that include various structural components for the national aerospace plane, the space shuttle main engine, and advanced gas turbines. Specifically, these applications include segmented engine liners, small missile engine turbine rotors, and exhaust nozzles. Use of these materials allows for improvements in fuel efficiency due to increased engine temperatures and pressures, which in turn generate more power and thrust. Furthermore, this class of materials offers significant potential for raising the thrust-to-weight ratio of gas turbine engines by tailoring directions of high specific reliability. The emerging composite systems, particularly those with silicon nitride or silicon carbide matrix, can compete with metals in many demanding applications. Laminated CMC prototypes have already demonstrated functional capabilities at temperatures approaching 1400 C, which is well beyond the operational limits of most metallic materials. Laminated CMC material systems have several mechanical characteristics which must be carefully considered in the design process. Test bed software programs are needed that incorporate stochastic design concepts that are user friendly, computationally efficient, and have flexible architectures that readily incorporate changes in design philosophy. The CCARES (Composite Ceramics Analysis and Reliability Evaluation of Structures) program is representative of an effort to fill this need. CCARES is a public domain computer algorithm, coupled to a general purpose finite element program, which predicts the fast fracture reliability of a structural component under multiaxial loading conditions.

Wind tunnel test of model target thrust reversers for the Pratt and Whitney aircraft JT8D-100 series engines installed on a 727-200 airplane

NASA Technical Reports Server (NTRS)

Hambly, D.

1974-01-01

The results of a low speed wind tunnel test of 0.046 scale model target thrust reversers installed on a 727-200 model airplane are presented. The full airplane model was mounted on a force balance, except for the nacelles and thrust reversers, which were independently mounted and isolated from it. The installation had the capability of simulating the inlet airflows and of supplying the correct proportions of primary and secondary air to the nozzles. The objectives of the test were to assess the compatibility of the thrust reversers target door design with the engine and airplane. The following measurements were made: hot gas ingestion at the nacelle inlets; model lift, drag, and pitching moment; hot gas impingement on the airplane structure; and qualitative assessment of the rudder effectiveness. The major parameters controlling hot gas ingestion were found to be thrust reverser orientation, engine power setting, and the lip height of the bottom thrust reverser doors on the side nacelles. The thrust reversers tended to increase the model lift, decrease the drag, and decrease the pitching moment.

A Review of Evidence for High Life Coefficients on Propeller and Rotor Blades Under Static Thrust Conditions with Some New Experimental Results

NASA Technical Reports Server (NTRS)

Talbot, Peter D.; Meyer, Mark; Branum, Lonnie; Burks, John S. (Technical Monitor)

1994-01-01

Interest has increased recently in the thrust-producing capability of rotors at very high collective pitch angles. An early reference noted this behaviour in rotors and offered alternative models for section lift characteristics to explain it. The same phenomenon was coincidentally noted and used in a propeller code, resulting in very good correlation with static thrust data. The proposed paper will present experimental data demonstrating the pronounced persistence of thrust for propellers at increasing collective pitch angles. Comparisons with blade element/momentum theory will be made. These results are expected to point to the need to define (ultimately to explain) aerodynamic lift and drag behaviour in a rotating environment. Experimental measurements made by the U.S. Army Aeroflightdynamics Directorate at the Ames Research Center have shown that locally measured normal force coefficients along the span of a highly twisted rotor blade continue to increase at high values of collective pitch. In some cases these coefficients exceed expected values for the same type of airfoil tested under two dimensional conditions. To date no one to the authors' knowledge has defined the variation of C(n) with pitch for very high angles (to 45 deg) in a rotating environment and for a blade of reasonably high aspect ratio; however, total propeller thrust measurements support the idea that stalling does not occur in the same way as on a wing. This paper will present experimental data in the form of surface pressure distributions as well as flow visualization (microtufts) to explore the aerodynamic behavior of the rotating airfoil at high values of blade incidence. This paper also reviews experimental evidence and infers some high lift coefficient behavior from it. Comparisons between predicted thrust, utilizing modified airfoil characteristics and a blade element model, and measured thrust for both rotors and propellers that cover the extremes of collective pitch are shown and discussed.

The Experimental Study about the Effect of Operating Conditions on Multi-tube Pulse Detonation Engine Performance

NASA Astrophysics Data System (ADS)

Kim, Jung-Min; Han, Hyung-Seok; Choi, Jeong-Yeol

2018-04-01

This study examines a multi-tube pulse detonation engine (PDE) which has a type of constant volume combustion. We designed and made the multi-tube PDE and then conducted an experiment in various operating frequencies and equivalence ratios. First, experiments with operating frequencies of 40, 80, 120, 160, and 200 Hz resulted in an average thrust and specific impulse 23.14 N and 42.34 s. The next experiment resulted in the equivalence ratio varying from 0.81 to 1.38, which resulted in an average thrust and specific impulse 22.36 N and 40.11 s. The average detonation velocity was 8% lower than that calculated according to C-J theory. The incidence ratios of the detonation wave were stable with the exception of the operating frequency of 200 Hz. However, at 200 Hz, the incidence ratio was less than 50%. We assumed that a low fill fraction occurred for this problem. The thrust of the PDE increased with the operating frequency. However, the thrust increase was at a lower rate than in previous studies, because of a lost thrust output result from the slow response time of the load cell amplifier.

A flow visualization study of single-arm sculling movement emulating cephalopod thrust generation

NASA Astrophysics Data System (ADS)

Kazakidi, Asimina; Gnanamanickam, Ebenezer P.; Tsakiris, Dimitris P.; Ekaterinaris, John A.

2014-11-01

In addition to jet propulsion, octopuses use arm-swimming motion as an effective means of generating bursts of thrust, for hunting, defense, or escape. The individual role of their arms, acting as thrust generators during this motion, is still under investigation, in view of an increasing robotic interest for alternative modes of propulsion, inspired by the octopus. Computational studies have revealed that thrust generation is associated with complex vortical flow patterns in the wake of the moving arm, however further experimental validation is required. Using the hydrogen bubble technique, we studied the flow disturbance around a single octopus-like robotic arm, undergoing two-stroke sculling movements in quiescent fluid. Although simplified, sculling profiles have been found to adequately capture the fundamental kinematics of the octopus arm-swimming behavior. In fact, variation of the sculling parameters alters considerably the generation of forward thrust. Flow visualization revealed the generation of complex vortical structures around both rigid and compliant arms. Increased disturbance was evident near the tip, particularly at the transitional phase between recovery and power strokes. These results are in good qualitative agreement with computational and robotic studies. Work funded by the ESF-GSRT HYDRO-ROB Project PE7(281).

JT8D-100 turbofan engine, phase 1. [noise reduction

NASA Technical Reports Server (NTRS)

1974-01-01

The JT8D turbofan engine, widely used in short and medium range transport aircraft, contributes substantially to airport community noise. The jet noise is predominant in the JT8D engine and may be reduced in a modified engine, without loss of thrust, by increasing the airflow to reduce jet velocity. A configuration study evaluated the effects of fan airflow, fan pressure ratio, and bypass ratio on noise, thrust, and fuel comsumption. The cycle selected for the modified engine was based upon an increased diameter, single-stage fan and two additional core engine compressor stages, which replace the existing two-stage fan. Modifications were also made to the low pressure turbine to provide the increased torque required by the larger diameter fan. The resultant JT8D-100 engine models have the following characteristics at take-off thrust, compared to the current JT8D engine: Airflow and bypass ratio are increased, and fan pressure ratio and engine speed are reduced. The resultant engine is also longer, larger in diameter, and heavier than the JT8D base model, but these latter changes are compensated by the increased thrust and decreased fuel comsumption of the modified engine, thus providing the capability for maintaining the performance of the current JT8D-powered aircraft.

Performance variation due to stiffness in a tuna-inspired flexible foil model.

PubMed

Rosic, Mariel-Luisa N; Thornycroft, Patrick J M; Feilich, Kara L; Lucas, Kelsey N; Lauder, George V

2017-01-17

Tuna are fast, economical swimmers in part due to their stiff, high aspect ratio caudal fins and streamlined bodies. Previous studies using passive caudal fin models have suggested that while high aspect ratio tail shapes such as a tuna's generally perform well, tail performance cannot be determined from shape alone. In this study, we analyzed the swimming performance of tuna-tail-shaped hydrofoils of a wide range of stiffnesses, heave amplitudes, and frequencies to determine how stiffness and kinematics affect multiple swimming performance parameters for a single foil shape. We then compared the foil models' kinematics with published data from a live swimming tuna to determine how well the hydrofoil models could mimic fish kinematics. Foil kinematics over a wide range of motion programs generally showed a minimum lateral displacement at the narrowest part of the foil, and, immediately anterior to that, a local area of large lateral body displacement. These two kinematic patterns may enhance thrust in foils of intermediate stiffness. Stiffness and kinematics exhibited subtle interacting effects on hydrodynamic efficiency, with no one stiffness maximizing both thrust and efficiency. Foils of intermediate stiffnesses typically had the greatest coefficients of thrust at the highest heave amplitudes and frequencies. The comparison of foil kinematics with tuna kinematics showed that tuna motion is better approximated by a zero angle of attack foil motion program than by programs that do not incorporate pitch. These results indicate that open questions in biomechanics may be well served by foil models, given appropriate choice of model characteristics and control programs. Accurate replication of biological movements will require refinement of motion control programs and physical models, including the creation of models of variable stiffness.

Lower extremity thrust and non-thrust joint mobilization for patellofemoral pain syndrome: a case report.

PubMed

Simpson, Brad G; Simon, Corey B

2014-05-01

A 40-year old female presented to physical therapy with a one-year history of insidious right anteromedial and anterolateral knee pain. Additionally, the patient had a history of multiple lateral ankle sprains bilaterally, the last sprain occurring on the right ankle 1 year prior to the onset of knee pain. The patient was evaluated and given a physical therapy diagnosis of patellofemoral pain syndrome (PFPS), with associated talocrural and tibiofemoral joint hypomobility limiting ankle dorsiflexion and knee extension, respectively. Treatment included a high-velocity low amplitude thrust manipulation to the talocrural joint, which helped restore normal ankle dorsiflexion range of motion. The patient also received tibiofemoral joint non-thrust manual therapy to regain normal knee extension mobility prior to implementing further functional progression exercises to her home program (HEP). This case report highlights the importance of a detailed evaluation of knee and ankle joint mobility in patients presenting with anterior knee pain. Further, manual physical therapy to the lower extremity was found to be successful in restoring normal movement patterns and pain-free function in a patient with chronic anterior knee pain.

A thermal control approach for a solar electric propulsion thrust subsystem

NASA Technical Reports Server (NTRS)

Maloy, J. E.; Oglebay, J. C.

1979-01-01

A thrust subsystem thermal control design is defined for a Solar Electric Propulsion System (SEPS) proposed for the comet Halley Flyby/comet Tempel 2 rendezvous mission. A 114 node analytic model, developed and coded on the systems improved numerical differencing analyzer program, was employed. A description of the resulting thrust subsystem thermal design is presented as well as a description of the analytic model and comparisons of the predicted temperature profiles for various SEPS thermal configurations that were generated using this model. It was concluded that: (1) a BIMOD engine system thermal design can be autonomous; (2) an independent thrust subsystem thermal design is feasible; (3) the interface module electronics temperatures can be controlled by a passive radiator and supplementary heaters; (4) maintaining heat pipes above the freezing point would require an additional 322 watts of supplementary heating power for the situation where no thrusters are operating; (5) insulation is required around the power processors, and between the interface module and the avionics module, as well as in those areas which may be subjected to solar heating; and (6) insulation behind the heat pipe radiators is not necessary.

Stem thrust prediction model for W-K-M double wedge parallel expanding gate valves

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

Eldiwany, B.; Alvarez, P.D.; Wolfe, K.

1996-12-01

An analytical model for determining the required valve stem thrust during opening and closing strokes of W-K-M parallel expanding gate valves was developed as part of the EPRI Motor-Operated Valve Performance Prediction Methodology (EPRI MOV PPM) Program. The model was validated against measured stem thrust data obtained from in-situ testing of three W-K-M valves. Model predictions show favorable, bounding agreement with the measured data for valves with Stellite 6 hardfacing on the disks and seat rings for water flow in the preferred flow direction (gate downstream). The maximum required thrust to open and to close the valve (excluding wedging andmore » unwedging forces) occurs at a slightly open position and not at the fully closed position. In the nonpreferred flow direction, the model shows that premature wedging can occur during {Delta}P closure strokes even when the coefficients of friction at different sliding surfaces are within the typical range. This paper summarizes the model description and comparison against test data.« less

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.

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 chemical orbit to orbit propulsion system propellant management study

NASA Technical Reports Server (NTRS)

Dergance, R. H.

1980-01-01

Propellant requirements, tankage configurations, preferred propellant management techniques, propulsion systems weights, and technology deficiencies for low thrust expendable propulsion systems are examined. A computer program was utilized which provided a complete propellant inventory (including boil-off for cryogenic cases), pressurant and propellant tank dimensions for a given ullage, pressurant requirements, insulation requirements, and miscellaneous masses. The output also includes the masses of all tanks; the mass of the insulation, engines and other components; total wet system and burnout mass; system mass fraction; total impulse and burn time.

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.

Advanced Space Transportation Program (ASTP)

NASA Image and Video Library

1997-08-07

This double exposure depicts Marshall Space Flight Center's (MSFC) Test Stand 116 hosting a 60K Bantam Fastrac thrust chamber assembly test. The lower right exposure shows the engine firing in the test stand while the center exposure reveals workers monitoring the test in the interior block house of the test facility. The thrust chamber assembly is only part of the Fastrac engine project to build a low-cost engine for the X-34, an alternate light-weight unmarned launch vehicle. Both the nozzle and the engine for Fastrac are being manufactured at MSFC.

General view of a Solid Rocket Motor Nozzle in the ...

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

General view of a Solid Rocket Motor Nozzle in the Solid Rocket Booster (SRB) Assembly and Refurbishment Facility at Kennedy Space Center, being prepared to be mated with the Aft Skirt. In this view you can see the attach brackets where the Thrust Vector Control System actuators connect to the nozzle which can swivel the nozzle up to 3.5 degrees to redirect the thrust to steer and maintain the Shuttle's programmed trajectory. - Space Transportation System, Solid Rocket Boosters, Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

Saturn Apollo Program

NASA Image and Video Library

1968-01-01

A complete F-1 engine assembly is shown in this photograph. Designed and developed by Rocketdye under the direction of the Marshall Space Flight Center, the engine measured 19-feet tall by 12.5 feet at the nozzle exit, and each engine produced a 1,500,000-pound thrust using liquid oxygen and kerosene as the propellant. A cluster of five F-1 engines was mounted on the Saturn V S-IC (first) stage and burned 15 tons of liquid oxygen and kerosene each second to produce 7,500,000 pounds of thrust.

Saturn Apollo Program

NASA Image and Video Library

2004-04-15

H-1 Engine major components with callouts (chart 1): The H-1 engine was used in a cluster of eight on the the first stage of Saturn I (S-I stage) and Saturn IB (S-IB stage). The engines were arranged in a double pattern: four engines, located inboard, were fixed in a square pattern around the stage axis, while the remaining four engines were located outboard in a larger square pattern and each outer engine was gimbaled. Each H-1 engine had a thrust of 188,000 pounds for a combined thrust of over 1,500,000 pounds.

Saturn Apollo Program

NASA Image and Video Library

2004-04-15

H-1 engine major components with callouts (chart 1). The H-1 engine was used in a cluster of eight on the the first stage of Saturn I (S-I stage) and Saturn IB (S-IB stage). The engines were arranged in a double pattern: four engines, located inboard, were fixed in a square pattern around the stage axis, while the remaining four engines were located outboard in a larger square pattern and each outer engine was gimbaled. Each H-1 engine had a thrust of 188,000 pounds for a combined thrust of over 1,500,000 pounds.

Saturn Apollo Program

NASA Image and Video Library

1968-03-01

The Saturn 1B first stage (S-IB) enters the NASA barge Point Barrow, in March 1968. The Marshall Space Flight Center (MSFC) utilized a number of water transportation craft to transport the Saturn stages to-and-from the manufacturing facilities and test sites, as well as delivery to the Kennedy Space Center for launch. Developed by the Marshall Space Flight Center and built by the Chrysler Corporation at Michoud Assembly Facility (MAF), the S-IB utilized the eight H-1 engines and each produced 200,000 pounds of thrust, a combined thrust of 1,600,000 pounds.

Sirius-5 experimental rocket

NASA Astrophysics Data System (ADS)

Kerstein, A.; Omersel, P.; Goljuf, L.; Zidaric, M.

1981-09-01

After giving a historical account of multistage rocket development in Yugoslavia, a status report is presented for the three-stage Sirius-5 program. The rocket is composed of: (1) a solid-propellant first stage, consisting of a cluster of eight standard motors yielding 220 kN thrust for 1.3 sec; (2) a mixed amines/inhibited red fuming nitric acid, bipropellant second stage generating 50 kN thrust; and (3) a third stage of the same design as the second but with only 62 kg of fuel, by contrast to 168 kg. Among the design principles adhered to are: minimization of the number of components, conservative design margins, and specifications for key subsystems based on demonstration programs. The primary use of this system is in amateur rocketry, being able to carry a 20 kg payload to 150 km.

LIDAR Remote Sensing Concepts

NASA Technical Reports Server (NTRS)

Spiers, Gary D.

1997-01-01

The primary goal of the NASA New Millennium Program (NMP) is to develop technology for use on future operational missions. The Program consists of two thrust areas, one oriented towards developing technologies for Deep Space Probes and one oriented towards developing technology for Earth Observing Probes. Each thrust area intends to fly several technology demonstrator space designated DS-X and EO-X respectively where X is the mission number. Each mission has an approximately $100 million cap on total mission cost. The EO-1 mission has been selected and is under development. The instrument discussed here was submitted by NASA MSFC as a potential candidate for the EO-2 or EO-3 missions due to launch in 2001 and late 2002 or early 2003 respectively. This report summarizes and follows the format of the material provided to NMP.

LIDAR Remote Sensing Concepts

NASA Technical Reports Server (NTRS)

Spiers, Gary D.

1997-01-01

The primary goal of the NASA New Millennium Program (NMP) is to develop technology for use on future operational missions. The Program consists of two thrust areas, one oriented towards developing technologies for Deep Space Probes and one oriented towards developing technology for Earth Observing Probes. Each thrust area intends to fly several technology demonstrator spacecraft designated DS-X and EO-X respectively where X is the mission number. Each mission has an approximately $100 million cap on total mission cost. The EO-1 mission has been selected and is under development. The instrument discussed here was submitted by NASA MSFC as a potential candidate for the EO-2 or EO-3 missions due to launch in 2001 and late 2002 or early 2003 respectively. This report summarizes and follows the format of the material provided to NMP.

OAST Space Theme Workshop. Volume 3: Working group summary. 4: Software (E-4). A. Summary. B. Technology needs (form 1). C. Priority assessment (form 2)

NASA Technical Reports Server (NTRS)

1976-01-01

Only a few efforts are currently underway to develop an adequate technology base for the various themes. Particular attention must be given to software commonality and evolutionary capability, to increased system integrity and autonomy; and to improved communications among the program users, the program developers, and the programs themselves. There is a need for quantum improvement in software development methods and increasing the awareness of software by all concerned. Major thrusts identified include: (1) data and systems management; (2) software technology for autonomous systems; (3) technology and methods for improving the software development process; (4) advances related to systems of software elements including their architecture, their attributes as systems, and their interfaces with users and other systems; and (5) applications of software including both the basic algorithms used in a number of applications and the software specific to a particular theme or discipline area. The impact of each theme on software is assessed.

Deformation of Fold-and-Thrust Belts above a Viscous Detachment: New Insights from Analogue Modelling Experiments

NASA Astrophysics Data System (ADS)

Nogueira, Carlos R.; Marques, Fernando O.

2015-04-01

Theoretical and experimental studies on fold-and-thrusts belts (FTB) have shown that, under Coulomb conditions, deformation of brittle thrust wedges above a dry frictional basal contact is characterized by dominant frontward vergent thrusts (forethrusts) with thrust spacing and taper angle being directly influenced by the basal strength (increase in basal strength leading to narrower thrust spacing and higher taper angles); whereas thrust wedges deformed above a weak viscous detachment, such as salt, show a more symmetric thrust style (no prevailing vergence of thrusting) with wider thrust spacing and shallower wedges. However, different deformation patterns can be found on this last group of thrust wedges both in nature and experimentally. Therefore we focused on the strength (friction) of the wedge basal contact, the basal detachment. We used a parallelepiped box with four fixed walls and one mobile that worked as a vertical piston drove by a computer controlled stepping motor. Fine dry sand was used as the analogue of brittle rocks and silicone putty (PDMS) with Newtonian behaviour as analogue of the weak viscous detachment. To investigate the strength of basal contact on thrust wedge deformation, two configurations were used: 1) a horizontal sand pack with a dry frictional basal contact; and 2) a horizontal sand pack above a horizontal PDMS layer, acting as a basal weak viscous contact. Results of the experiments show that: the model with a dry frictional basal detachment support the predictions for the Coulomb wedges, showing a narrow wedge with dominant frontward vergence of thrusting, close spacing between FTs and high taper angle. The model with a weak viscous frictional basal detachment show that: 1) forethrusts (FT) are dominant showing clearly an imbricate asymmetric geometry, with wider spaced thrusts than the dry frictional basal model; 2) after FT initiation, the movement on the thrust can last up to 15% model shortening, leading to great amount of displacement along the FT; 3) intermittent reactivation of FTs also occurs despite the steepening of the FT plane and existence of new FT ahead, creating a high critical taper angle; 4) injection of PDMS from the basal weak layer into the FTs planes also favours to the long living of FTs and to the high critical taper angle; 5) vertical sand thickening in the hanging block also added to the taper angle.

Comparisons Between Stability Prediction and Measurements for the Reusable Solid Rocket Motor

NASA Technical Reports Server (NTRS)

Fischbach, Sean R.; Kenny, R. Jeremy

2010-01-01

The Space Transportation System has used the solid rocket boosters for lift-off and ascent propulsion over the history of the program. Part of the structural loads assessment of the assembled vehicle is the contribution due to solid rocket booster thrust oscillations. These thrust oscillations are a consequence of internal motor pressure oscillations active during operation. Understanding of these pressure oscillations is key to predicting the subsequent thrust oscillations and vehicle loading. The pressure oscillation characteristics of the Reusable Solid Rocket Motor (RSRM) design are reviewed in this work. Dynamic pressure data from the static test and flight history are shown, with emphasis on amplitude, frequency, and timing of the oscillations. Physical mechanisms that cause these oscillations are described by comparing data observations to predictions made by the Solid Stability Prediction (SSP) code.

Optimization of self-acting step thrust bearings for load capacity and stiffness.

NASA Technical Reports Server (NTRS)

Hamrock, B. J.

1972-01-01

Linearized analysis of a finite-width rectangular step thrust bearing. Dimensionless load capacity and stiffness are expressed in terms of a Fourier cosine series. The dimensionless load capacity and stiffness were found to be a function of the dimensionless bearing number, the pad length-to-width ratio, the film thickness ratio, the step location parameter, and the feed groove parameter. The equations obtained in the analysis were verified. The assumptions imposed were substantiated by comparing the results with an existing exact solution for the infinite width bearing. A digital computer program was developed which determines optimal bearing configuration for maximum load capacity or stiffness. Simple design curves are presented. Results are shown for both compressible and incompressible lubrication. Through a parameter transformation the results are directly usable in designing optimal step sector thrust bearings.

Examination of various turbulence models for application in liquid rocket thrust chambers

NASA Technical Reports Server (NTRS)

Hung, R. J.

1991-01-01

There is a large variety of turbulence models available. These models include direct numerical simulation, large eddy simulation, Reynolds stress/flux model, zero equation model, one equation model, two equation k-epsilon model, multiple-scale model, etc. Each turbulence model contains different physical assumptions and requirements. The natures of turbulence are randomness, irregularity, diffusivity and dissipation. The capabilities of the turbulence models, including physical strength, weakness, limitations, as well as numerical and computational considerations, are reviewed. Recommendations are made for the potential application of a turbulence model in thrust chamber and performance prediction programs. The full Reynolds stress model is recommended. In a workshop, specifically called for the assessment of turbulence models for applications in liquid rocket thrust chambers, most of the experts present were also in favor of the recommendation of the Reynolds stress model.

In-flight thrust determination on a real-time basis

NASA Technical Reports Server (NTRS)

Ray, R. J.; Carpenter, T.; Sandlin, T.

1984-01-01

A real time computer program was implemented on a F-15 jet fighter to monitor in-flight engine performance of a Digital Electronic Engine Controlled (DEES) F-100 engine. The application of two gas generator methods to calculate in-flight thrust real time is described. A comparison was made between the actual results and those predicted by an engine model simulation. The percent difference between the two methods was compared to the predicted uncertainty based on instrumentation and model uncertainty and agreed closely with the results found during altitude facility testing. Data was obtained from acceleration runs of various altitudes at maximum power settings with and without afterburner. Real time in-flight thrust measurement was a major advancement to flight test productivity and was accomplished with no loss in accuracy over previous post flight methods.

Space tug geosynchronous mission simulation

NASA Technical Reports Server (NTRS)

Lang, T. J.

1973-01-01

Near-optimal three dimensional trajectories from a low earth park orbit inclined at 28.5 deg to a synchronous-equatorial mission orbit were developed for both the storable (thrust = 28,912 N (6,500 lbs), I sub sp = 339 sec) and cryogenic (thrust = 44,480 N (10,000 lbs), I sub sp = 470 sec) space tug using the iterative cost function minimization technique contained within the modularized vehicle simulation (MVS) program. The finite burn times, due to low thrust-to-weight ratios, and the associated gravity losses are accounted for in the trajectory simulation and optimization. The use of an ascent phasing orbit to achieve burnout in synchronous orbit at any longitude is investigated. The ascent phasing orbit is found to offer the additional advantage of significantly reducing the overall delta velocity by splitting the low altitude burn into two parts and thereby reducing gravity losses.

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.

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.

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.

Effect of applied force and blade speed on histopathology of bone during resection by sagittal saw.

PubMed

James, Thomas P; Chang, Gerard; Micucci, Steven; Sagar, Amrit; Smith, Eric L; Cassidy, Charles

2014-03-01

A sagittal saw is commonly used for resection of bone during joint replacement surgery. During sawing, heat is generated that can lead to an increase in temperature at the resected surface. The aim of this study was to determine the effect of applied thrust force and blade speed on generating heat. The effect of these factors and their interactions on cutting temperature and bone health were investigated with a full factorial Design of Experiments approach for two levels of thrust force, 15 N and 30 N, and for two levels of blade oscillation rate, 12,000 and 18,000 cycles per minute (cpm). In addition, a preliminary study was conducted to eliminate blade wear as a confounding factor. A custom sawing fixture was used to crosscut samples of fresh bovine cortical bone while temperature in the bone was measured by thermocouple (n=40), followed by measurements of the depth of thermal necrosis by histopathological analysis (n=200). An analysis of variance was used to determine the significance of the factor effects on necrotic depth as evidenced by empty lacunae. Both thrust force and blade speed demonstrated a statistically significant effect on the depth of osteonecrosis (p<0.05), while the interaction of thrust force with blade speed was not significant (p=0.22). The minimum necrotic depth observed was 0.50mm, corresponding to a higher level of force and blade speed (30 N, 18,000 cpm). Under these conditions, a maximum temperature of 93°C was measured at 0.3mm from the kerf. With a decrease in both thrust force and blade speed (15N, 12,000 cpm), the temperature in the bone increased to 109°C, corresponding to a nearly 50% increase in depth of the necrotic zone to 0.74 mm. A predictive equation for necrotic depth in terms of thrust force and blade speed was determined through regression analysis and validated by experiment. The histology results imply that an increase in applied thrust force is more effective in reducing the depth of thermal damage to surrounding bone than an increase in blade speed. Copyright © 2013 IPEM. Published by Elsevier Ltd. All rights reserved.

Quantifying the air quality-CO2 tradeoff potential for airports

NASA Astrophysics Data System (ADS)

Ashok, Akshay; Dedoussi, Irene C.; Yim, Steve H. L.; Balakrishnan, Hamsa; Barrett, Steven R. H.

2014-12-01

Aircraft movements on the airport surface are responsible for CO2 emissions that contribute to climate change and other emissions that affect air quality and human health. While the potential for optimizing aircraft surface movements to minimize CO2 emissions has been assessed, the implications of CO2 emissions minimization for air quality have not been quantified. In this paper, we identify conditions in which there is a tradeoff between CO2 emissions and population exposure to O3 and secondary PM2.5 - i.e. where decreasing fuel burn (which is directly proportional to CO2 emissions) results in increased exposure. Fuel burn and emissions are estimated as a function of thrust setting for five common gas turbine engines at 34 US airports. Regional air quality impacts, which are dominated by ozone and secondary PM2.5, are computed as a function of airport location and time using the adjoint of the GEOS-Chem chemistry-transport model. Tradeoffs between CO2 emissions and population exposure to PM2.5 and O3 occur between 2-18% and 5-60% of the year, respectively, depending on airport location, engine type, and thrust setting. The total duration of tradeoff conditions is 5-12 times longer at maximum thrust operations (typical for takeoff) relative to 4% thrust operations (typical for taxiing). Per kilogram of additional fuel burn at constant thrust setting during tradeoff conditions, reductions in population exposure to PM2.5 and O3 are 6-13% and 32-1060% of the annual average (positive) population exposure per kilogram fuel burn, where the ranges encompass the medians over the 34 airports. For fuel burn increases due to thrust increases (i.e. for constant operating time), reductions in both PM2.5 and O3 exposure are 1.5-6.4 times larger in magnitude than those due to increasing fuel burn at constant thrust (i.e. increasing operating time). Airports with relatively high population exposure reduction potentials - which occur due to a combination of high duration and magnitude of tradeoff conditions - are identified. Our results are the first to quantify the extent of the tradeoff between CO2 emissions and air quality impacts at airports. This raises the possibility of reducing the air quality impacts of airports beyond minimizing fuel burn and/or optimizing for minimum net environmental impact.

Stress transfer among en echelon and opposing thrusts and tear faults: Triggering caused by the 2003 Mw = 6.9 Zemmouri, Algeria, earthquake

USGS Publications Warehouse

Lin, J.; Stein, R.S.; Meghraoui, M.; Toda, S.; Ayadi, A.; Dorbath, C.; Belabbes, S.

2011-01-01

The essential features of stress interaction among earthquakes on en echelon thrusts and tear faults were investigated, first through idealized examples and then by study of thrust faulting in Algeria. We calculated coseismic stress changes caused by the 2003 Mw = 6.9 Zemmouri earthquake, finding that a large majority of the Zemmouri afterslip sites were brought several bars closer to Coulomb failure by the coseismic stresses, while the majority of aftershock nodal planes were brought closer to failure by an average of ~2 bars. Further, we calculated that the shallow portions of the adjacent Thenia tear fault, which sustained ~0.25 m slip, were brought >2 bars closer to failure. We calculated that the Coulomb stress increased by 1.5 bars on the deeper portions of the adjacent Boumerdes thrust, which lies just 10–20 km from the city of Algiers; both the Boumerdes and Thenia faults were illuminated by aftershocks. Over the next 6 years, the entire south dipping thrust system extending 80 km to the southwest experienced an increased rate of seismicity. The stress also increased by 0.4 bar on the east Sahel thrust fault west of the Zemmouri rupture. Algiers suffered large damaging earthquakes in A.D. 1365 and 1716 and is today home to 3 million people. If these shocks occurred on the east Sahel fault and if it has a ~2 mm/yr tectonic loading rate, then enough loading has accumulated to produce a Mw = 6.6–6.9 shock today. Thus, these potentially lethal faults need better understanding of their slip rate and earthquake history.

Pulsed Ejector Thrust Amplification Tested and Modeled

NASA Technical Reports Server (NTRS)

Wilson, Jack

2004-01-01

There is currently much interest in pulsed detonation engines for aeronautical propulsion. This, in turn, has sparked renewed interest in pulsed ejectors to increase the thrust of such engines, since previous, though limited, research had indicated that pulsed ejectors could double the thrust in a short device. An experiment has been run at the NASA Glenn Research Center, using a shrouded Hartmann-Sprenger tube as a source of pulsed flow, to measure the thrust augmentation of a statistically designed set of ejectors. A Hartmann- Sprenger tube directs the flow from a supersonic nozzle (Mach 2 in the present experiment) into a closed tube. Under appropriate conditions, an oscillation is set up in which the jet flow alternately fills the tube and then spills around flow emerging from the tube. The tube length determines the frequency of oscillation. By shrouding the tube, the flow was directed out of the shroud as an axial stream. The set of ejectors comprised three different ejector lengths, three ejector diameters, and three nose radii. The thrust of the jet alone, and then of the jet plus ejector, was measured using a thrust plate. The arrangement is shown in this photograph. Thrust augmentation is defined as the thrust of the jet with an ejector divided by the thrust of the jet alone. The experiments exhibited an optimum ejector diameter and length for maximizing the thrust augmentation, but little dependence on nose radius. Different frequencies were produced by changing the length of the Hartmann-Sprenger tube, and the experiment was run at a total of four frequencies. Additional measurements showed that the major feature of the pulsed jet was a starting vortex ring. The size of the vortex ring depended on the frequency, as did the optimum ejector diameter.

Intrinsic And Extrinsic Controls On Unsteady Deformation Rates, Northern Apennine Mountains, Italy

NASA Astrophysics Data System (ADS)

Anastasio, D. J.; Gunderson, K. L.; Pazzaglia, F. J.; Kodama, K. P.

2017-12-01

The slip rates of faults in the Northern Apennine Mountains were unsteady at 104-105 year timescales during the Neogene and Quaternary. Fault slip rates were recovered from growth strata and uplifted fluvial terraces associated with the Salsomaggiore, Quatto Castella, and Castevetro fault-related folds, sampled along the Stirone, Enza, and Panaro Rivers, respectively. The forelimb stratigraphy of each anticline was dated using rock magnetic-based cyclostratigraphy, which varies with Milankovitch periodicity, multispecies biostratigraphy, magnetostratigraphy, OSL luminescence dating, TCN burial dating, and radiocarbon dating of uplifted and folded fluvial terraces. Fault slip magnitudes were constrained with trishear forward models. We observed decoupled deformation and sediment accumulation rates at each structure. From 3.5Ma deformation of a thick and thin-skinned thrusts was temporally variable and controlled by intrinsic rock processes, whereas, the more regional Pede-Apenninic thrust fault, a thick-skinned thrust underlying the mountain front, was likely activated because of extrinsic forcing from foreland basin sedimentation rate accelerations since 1.4Ma. We found that reconstructed slip rate variability increased as the time resolution increased. The reconstructed slip history of the thin-skinned thrust faults was characterized relatively long, slow fold growth and associated fault slip, punctuated by shorter, more rapid periods limb rotation, and slip on the underlying thrust fault timed asynchronously. Thrust fault slip rates slip rates were ≤ 0.1 to 6 mm/yr at these intermediate timescales. The variability of slip rates on the thrusts is likely related to strain partitioning neighboring faults within the orogenic wedge. The studied structures slowed down at 1Ma when there was a switch to slower synchronous fault slip coincident with orogenic wedge thickening due to the emplacement of the out of sequence Pene-Apenninic thrust fault that was emplaced at 1.4±0.7 mm/yr. Both tectonic control and climate controlled variability on syntectonic sedimentation was observed in the growth sections.

Neogene compressional deformation and possible thrust faulting in southwest Dominican Republic

NASA Technical Reports Server (NTRS)

Golombek, M. P.; Goreau, P.; Dixon, T. H.

1985-01-01

Analysis of regional and high resolution remote sensing data coupled with detailed field investigations indicates Neogene compressional deformation in the southwest Dominican Republic. Airborne synthetic aperture radar data and high resolution near infrared photography show folds in Tertiary sediments and possible thrust fault scarps implying NE to SW compression in the region. Large road cuts through the scarps allow study of otherwise poorly accessible, heavily vegetated karst terrain. Deformation increases toward scrap fronts where small bedding-plane thrust faults become more numerous. Analysis of mesoscopic faults with slickensides indicates compression oriented between N to S and E to W. The lowermost scarp has highly sheared fault breccia and undeformed frontal talus breccias implying it is the basal thrust into which the higher thrust faults sole. Thus, the scarps probably formed in a regional NE to SW compressional stress regime and are the toes of thrust sheets. Previous workers have suggested that these scarps are ancient shorelines. However, the gross morphology of the scarps differs substantially from well known erosional terraces on the north coast.

Performance of a green propellant thruster with discharge plasma

NASA Astrophysics Data System (ADS)

Shindo, Takahiro; Wada, Asato; Maeda, Hiroshi; Watanabe, Hiroki; Takegahara, Haruki

2017-02-01

A discharge plasma was applied to initiate the combustion of a hydroxylammonium nitrate-based propellant as a substitute for the catalysts that are typically employed. The resulting thrust and thrust-to-power ratio during short interval firing tests as well as the chamber pressure with a single pulse discharge were evaluated. A 1.5-s firing test generated a maximum thrust of 322 mN along with a thrust-to-power ratio of 0.95 mN/W. During the single-pulse discharge trials, pulsed discharge capacitor energies of 5.4, 10.8, and 16.4 J were assessed, and the maximum chamber pressure was found to increase as the energy was raised. The maximum chamber pressures varied widely between experimental trials, and a 16.4-J energy value resulted in the highest chamber pressure of over 1 MPaG. The time spans between the pulsed discharge and the peak chamber pressure were in the range of 1-2 ms, representing a chamber pressure increase rate much higher than those obtained with standard catalysts.

Effect of tibial plateau leveling on stability of the canine cranial cruciate-deficient stifle joint: an in vitro study.

PubMed

Reif, Ullrich; Hulse, Donald A; Hauptman, Joe G

2002-01-01

To evaluate the effect of tibial plateau leveling on joint motion in canine stifle joints in which the cranial cruciate ligament (CCL) had been severed. In vitro cadaver study. Six canine cadaver hind legs. Radiographs of the stifle joints were made to evaluate the tibial plateau angle with respect to the long axis of the tibia. The specimens were mounted in a custom-made testing device to measure cranio-caudal translation of the tibia with respect to the femur. An axial load was applied to the tibia, and its position was recorded in the normal stifle, after transection of the CCL, and after tibial plateau leveling. Further, the amount of caudal tibial thrust was measured in the tibial plateau leveled specimen while series of eight linearly increasing axial tibial loads were applied. Transection of the CCL resulted in cranial tibial translation when axial tibial load was applied. After tibial plateau leveling, axial loading resulted in caudal translation of the tibia. Increasing axial tibial load caused a linear increase in caudal tibial thrust in all tibial plateau-leveled specimens. After tibial plateau leveling, axial tibial load generates caudal tibial thrust, which increases if additional axial load is applied. Tibial plateau leveling osteotomy may prevent cranial translation during weight bearing in dogs with CCL rupture by converting axial load into caudal tibial thrust. The amount of caudal tibial thrust seems to be proportional to the amount of weight bearing. Copyright 2002 by The American College of Veterinary Surgeons

Simplified power processing for ion-thruster subsystems

NASA Technical Reports Server (NTRS)

Wessel, F. J.; Hancock, D. J.

1983-01-01

Compared to chemical propulsion, ion propulsion offers distinct payload-mass increases for many future low-thrust earth-orbital and deep-space missions. Despite this advantage, the high initial cost and complexity of ion-propulsion subsystems reduce their attractiveness for most present and near-term spacecraft missions. Investigations have, therefore, been conducted with the objective to attempt to simplify the power-processing unit (PPU), which is the single most complex and expensive component in the thruster subsystem. The present investigation is concerned with a program to simplify the design of the PPU employed in a 8-cm mercury-ion-thruster subsystem. In this program a dramatic simplification in the design of the PPU could be achieved, while retaining essential thruster control and subsystem operational flexibility.

Varus alignment leads to increased forces in the anterior cruciate ligament.

PubMed

van de Pol, Gerrit Jan; Arnold, Markus P; Verdonschot, Nico; van Kampen, Albert

2009-03-01

Varus thrust of the knee is a dynamic increase of an often preexisting varus angle and it is suspected to be a major reason for failure of anterior cruciate ligament reconstructions. However, it is not known if a direct relationship exists between varus thrust and forces in the anterior cruciate ligament. Forces in the anterior cruciate ligament increase with increasing varus alignment, and consequently an anterior cruciate ligament deficiency in a varus-aligned leg leads to more lateral tibiofemoral joint opening. Controlled laboratory study. Six human cadaver legs were axially loaded with 3 different weightbearing lines--a neutral weightbearing line, a weightbearing line that passes through the middle of the medial tibial plateau (50% varus), and a line passing the edge of the medial tibial plateau (100% varus)--that were used to create a varus moment. The resulting lateral tibiofemoral joint opening and corresponding anterior cruciate ligament tension were measured. The tests were repeated with and without the anterior cruciate ligament in place. In the neutral aligned legs, there was no apparent lateral joint opening, and no anterior cruciate ligament tension change was noted. The lateral joint opening increased when the weightbearing line increased from 0% to 50% to 100%. The lateral joint opening was significantly higher in 10 degrees of knee flexion compared with knee extension. In the 100% varus weightbearing line, the anterior cruciate ligament tension was significantly higher (53.9 N) compared with neutral (31 N) or the 50% weightbearing line (37.9 N). A thrust could only be observed in the 100% weightbearing line tests. In the absence of an anterior cruciate ligament, there was more lateral joint opening, although this was only significant in the 100% weightbearing line. There is a direct relationship between varus alignment and anterior cruciate ligament tension. In the absence of an anterior cruciate ligament, the amount of lateral opening tends to increase. With increasing lateral opening, a thrust can sometimes be experimentally observed. A varus alignment in an anterior cruciate ligament-deficient knee does not necessarily lead to a varus thrust and therefore does not always need operative varus alignment correction. However, in an unstable anterior cruciate ligament-deficient knee with a varus thrust, it might be safer to perform a high valgus tibial osteotomy to minimize the risk of an anterior cruciate ligament reconstruction failure.

NASA/FAA Tailplane Icing Program: Flight Test Report

NASA Technical Reports Server (NTRS)

Ratvasky, Thomas P.; VanZante, Judith Foss; Sim, Alex

2000-01-01

This report presents results from research flights that explored the characteristics of an ice-contaminated tailplane using various simulated ice shapes attached to the leading edge of the horizontal tailplane. A clean leading edge provided the baseline case, then three ice shapes were flown in order of increasing severity. Flight tests included both steady state and dynamic maneuvers. The steady state points were 1G wings level and steady heading sideslips. The primary dynamic maneuvers were pushovers to various G-levels; elevator doublets; and thrust transitions. These maneuvers were conducted for a full range of flap positions and aircraft angle of attack where possible. The analysis of this data set has clearly demonstrated the detrimental effects of ice contamination on aircraft stability and controllability. Paths to tailplane stall were revealed through parameter isolation and transition studies. These paths are (1) increasing ice shape severity, (2) increasing flap deflection, (3) high or low speeds, depending on whether the aircraft is in a steady state (high speed) or pushover maneuver (low speed), and (4) increasing thrust. The flight research effort was very comprehensive, but did not examine effects of tailplane design and location, or other aircraft geometry configuration effects. However, this effort provided the role of some of the parameters in promoting tailplane stall. The lessons learned will provide guidance to regulatory agencies, aircraft manufacturers, and operators on ice-contaminated tailplane stall in the effort to increase aviation safety and reduce the fatal accident rate.

A theoretical analysis of the effect of thrust-related turbulence distortion on helicopter rotor low-frequency broadband noise

NASA Technical Reports Server (NTRS)

Williams, M.; Harris, W. L.

1984-01-01

The purpose of the analysis is to determine if inflow turbulence distortion may be a cause of experimentally observed changes in sound pressure levels when the rotor mean loading is varied. The effect of helicopter rotor mean aerodynamics on inflow turbulence is studied within the framework of the turbulence rapid distortion theory developed by Pearson (1959) and Deissler (1961). The distorted inflow turbulence is related to the resultant noise by conventional broadband noise theory. A comparison of the distortion model with experimental data shows that the theoretical model is unable to totally explain observed increases in model rotor sound pressures with increased rotor mean thrust. Comparison of full scale rotor data with the theoretical model shows that a shear-type distortion may explain decreasing sound pressure levels with increasing thrust.

An approach to evaluate capacitance, capacitive reactance and resistance of pivoted pads of a thrust bearing

NASA Astrophysics Data System (ADS)

Prashad, Har

1992-07-01

A theoretical approach is developed for determining the capacitance and active resistance between the interacting surfaces of pivoted pads and thrust collar, under different conditions of operation. It is shown that resistance and capacitive reactance of a thrust bearing decrease with the number of pads times the values of these parameters for an individual pad, and that capacitance increases with the number of pads times the capacitance of an individual pad. The analysis presented has a potential to diagnose the behavior of pivoted pad thrust bearings with the angle of tilt and the ratio of film thickness at the leading to trailing edge, by determining the variation of capacitance, resistance, and capacitive reactance.

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.

14 CFR 417.207 - Trajectory analysis.

Code of Federal Regulations, 2013 CFR

2013-01-01

... any stage that has the potential to impact the Earth and does not burn to propellant depletion before a programmed thrust termination. (3) For launch vehicles flown with a flight safety system, a...

14 CFR 417.207 - Trajectory analysis.

Code of Federal Regulations, 2012 CFR

2012-01-01

... any stage that has the potential to impact the Earth and does not burn to propellant depletion before a programmed thrust termination. (3) For launch vehicles flown with a flight safety system, a...

14 CFR 417.207 - Trajectory analysis.

Code of Federal Regulations, 2014 CFR

2014-01-01

... any stage that has the potential to impact the Earth and does not burn to propellant depletion before a programmed thrust termination. (3) For launch vehicles flown with a flight safety system, a...

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.

Thrust production and wake structure of a batoid-inspired oscillating fin

PubMed Central

CLARK, R. P.; SMITS, A. J.

2009-01-01

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. PMID:19746188

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.

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.

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.

Multi-Objective Optimization of Moving-magnet Linear Oscillatory Motor Using Response Surface Methodology with Quantum-Behaved PSO Operator

NASA Astrophysics Data System (ADS)

Lei, Meizhen; Wang, Liqiang

2018-01-01

To reduce the difficulty of manufacturing and increase the magnetic thrust density, a moving-magnet linear oscillatory motor (MMLOM) without inner-stators was Proposed. To get the optimal design of maximum electromagnetic thrust with minimal permanent magnetic material, firstly, the 3D finite element analysis (FEA) model of the MMLOM was built and verified by comparison with prototype experiment result. Then the influence of design parameters of permanent magnet (PM) on the electromagnetic thrust was systematically analyzed by the 3D FEA to get the design parameters. Secondly, response surface methodology (RSM) was employed to build the response surface model of the new MMLOM, which can obtain an analytical model of the PM volume and thrust. Then a multi-objective optimization methods for design parameters of PM, using response surface methodology (RSM) with a quantum-behaved PSO (QPSO) operator, was proposed. Then the way to choose the best design parameters of PM among the multi-objective optimization solution sets was proposed. Then the 3D FEA of the optimal design candidates was compared. The comparison results showed that the proposed method can obtain the best combination of the geometric parameters of reducing the PM volume and increasing the thrust.

Evaluation of Candidate Millimeter Wave Sensors for Synthetic Vision

NASA Technical Reports Server (NTRS)

Alexander, Neal T.; Hudson, Brian H.; Echard, Jim D.

1994-01-01

The goal of the Synthetic Vision Technology Demonstration Program was to demonstrate and document the capabilities of current technologies to achieve safe aircraft landing, take off, and ground operation in very low visibility conditions. Two of the major thrusts of the program were (1) sensor evaluation in measured weather conditions on a tower overlooking an unused airfield and (2) flight testing of sensor and pilot performance via a prototype system. The presentation first briefly addresses the overall technology thrusts and goals of the program and provides a summary of MMW sensor tower-test and flight-test data collection efforts. Data analysis and calibration procedures for both the tower tests and flight tests are presented. The remainder of the presentation addresses the MMW sensor flight-test evaluation results, including the processing approach for determination of various performance metrics (e.g., contrast, sharpness, and variability). The variation of the very important contrast metric in adverse weather conditions is described. Design trade-off considerations for Synthetic Vision MMW sensors are presented.

Comparison of two procedures for predicting rocket engine nozzle performance

NASA Technical Reports Server (NTRS)

Davidian, Kenneth J.

1987-01-01

Two nozzle performance prediction procedures which are based on the standardized JANNAF methodology are presented and compared for four rocket engine nozzles. The first procedure required operator intercedence to transfer data between the individual performance programs. The second procedure is more automated in that all necessary programs are collected into a single computer code, thereby eliminating the need for data reformatting. Results from both procedures show similar trends but quantitative differences. Agreement was best in the predictions of specific impulse and local skin friction coefficient. Other compared quantities include characteristic velocity, thrust coefficient, thrust decrement, boundary layer displacement thickness, momentum thickness, and heat loss rate to the wall. Effects of wall temperature profile used as an input to the programs was investigated by running three wall temperature profiles. It was found that this change greatly affected the boundary layer displacement thickness and heat loss to the wall. The other quantities, however, were not drastically affected by the wall temperature profile change.

Liquid fluorine/hydrazine rhenium thruster update

NASA Technical Reports Server (NTRS)

Appel, M. A.; Kaplan, R. B.; Tuffias, R. H.

1983-01-01

The status of a fluorine/hydrazine thruster development program is discussed. A solid rhenium metal sea-level thrust chamber was successfully fabricated and tested for a total run duration of 1075 s with 17 starts. Rhenium fabrication methods are discussed. A test program was conducted to evaluate performance and chamber cooling. Acceptable performance was reached and cooling was adequate. A flight-type injector was fabricated that achieved an average extrapolated performance value of 3608 N-s/kg (368 lbf-s/lbm). Altitude thrust chambers were fabricated. One chamber incorporates a rhenium combustor and nozzle with an area ratio of 15:1, and a columbium nozzle extension with area ratios from 15:1 to 60:1. The other chamber was fabricated completely with a carbon/carbon composite. Because of the attributes of rhenium for use in high-temperature applications, a program to provide the materials and processes technology needed to reliably fabricate and/or repair vapor-deposited rhenium parts of relatively large size and complex shape is recommended.

Benefits and costs of low thrust propulsion systems

NASA Technical Reports Server (NTRS)

Robertson, R. I.; Rose, L. J.; Maloy, J. E.

1983-01-01

The results of costs/benefits analyses of three chemical propulsion systems that are candidates for transferring high density, low volume STS payloads from LEO to GEO are reported. Separate algorithms were developed for benefits and costs of primary propulsion systems (PPS) as functions of the required thrust levels. The life cycle costs of each system were computed based on the developmental, production, and deployment costs. A weighted criteria rating approach was taken for the benefits, with each benefit assigned a value commensurate to its relative worth to the overall system. Support costs were included in the costs modeling. Reference missions from NASA, commercial, and DoD catalog payloads were examined. The program was concluded reliable and flexible for evaluating benefits and costs of launch and orbit transfer for any catalog mission, with the most beneficial PPS being a dedicated low thrust configuration using the RL-10 system.

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.

Numerical Simulation of One-and Two-Phase Flows in Propulsion Systems

NASA Technical Reports Server (NTRS)

Blankson, Isaiah M. (Technical Monitor); Gilinsky, Mikhail

2002-01-01

In this report, we present some results of problems investigated during joint research between the Hampton University Fluid Mechanics and Acoustics Laboratory (FM&AL), NASA Glenn Research Center (GRC) and the Hyper-X Program of the NASA Langley Research Center (LaRC). The main areas of current scientific interest of the FM&AL include an investigation of the proposed and patented advanced methods for aircraft engine thrust and noise benefits. These methods are based on nontraditional 3D corrugated and composite nozzle, inlet, propeller and screw designs such as the Bluebell and Telescope nozzles, Mobius-shaped screws, etc. These are the main subject of our other projects, of which one is the NASA MURED's FAR Award. Working jointly with this project team, our team also analyzes additional methods for exhaust jet noise reduction. These methods are without essential thrust loss and even with thrust augmentation.

The Benefits of Nuclear Thermal Propulsion (NTP) in an Evolvable Mars Campaign

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Mccurdy, David R.

2014-01-01

NTR: High thrust high specific impulse (2 x LOXLH2chemical) engine uses high power density fission reactor with enriched uranium fuel as thermal power source. Reactor heat is removed using H2propellant which is then exhausted to produce thrust. Conventional chemical engine LH2tanks, turbopumps, regenerative nozzles and radiation-cooled shirt extensions used --NTR is next evolutionary step in high performance liquid rocket engines During the Rover program, a common fuel element tie tube design was developed and used in the design of the 50 klbf Kiwi-B4E (1964), 75 klbf Phoebus-1B (1967), 250 klbf Phoebus-2A (June 1968), then back down to the 25 klbf Pewee engine (Nov-Dec 1968) NASA and DOE are using this same approach: design, build, ground then flight test a small engine using a common fuel element that is scalable to a larger 25 klbf thrust engine needed for human missions

Breadboard RL10-2B low-thrust operating mode (second iteration) test report

NASA Technical Reports Server (NTRS)

Kanic, Paul G.; Kaldor, Raymond B.; Watkins, Pia M.

1988-01-01

Cryogenic rocket engines requiring a cooling process to thermally condition the engine to operating temperature can be made more efficient if cooling propellants can be burned. Tank head idle and pumped idle modes can be used to burn propellants employed for cooling, thereby providing useful thrust. Such idle modes required the use of a heat exchanger to vaporize oxygen prior to injection into the combustion chamber. During December 1988, Pratt and Whitney conducted a series of engine hot firing demonstrating the operation of two new, previously untested oxidizer heat exchanger designs. The program was a second iteration of previous low thrust testing conducted in 1984, during which a first-generation heat exchanger design was used. Although operation was demonstrated at tank head idle and pumped idle, the engine experienced instability when propellants could not be supplied to the heat exchanger at design conditions.

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.

HIDEC adaptive engine control system flight evaluation results

NASA Technical Reports Server (NTRS)

Yonke, W. A.; Landy, R. J.; Stewart, J. F.

1987-01-01

An integrated flight propulsion control mode, the Adaptive Engine Control System (ADECS), has been developed and flight tested on an F-15 aircraft as part of the NASA Highly Integrated Digital Electronic Control program. The ADECS system realizes additional engine thrust by increasing the engine pressure ratio (EPR) at intermediate and afterburning power, with the amount of EPR uptrim modulated using a predictor scheme for angle-of-attack and sideslip angle. Substantial improvement in aircraft and engine performance was demonstrated, with a 16 percent rate of climb increase, a 14 percent reduction in time to climb, and a 15 percent reduction in time to accelerate. Significant EPR uptrim capability was found with angles-of-attack up to 20 degrees.

Operant Control of Pathological Tongue Thrust in Spastic Cerebral Palsy.

ERIC Educational Resources Information Center

Thompson, George A., Jr.

1979-01-01

The behavior modification procedure, carried out at mealtime with a ten-year-old retarded boy who had spastic cerebral palsy, consisted of differential reinforcement and punishment, and resulted in substantial decreases in tongue thrust (reverse swallowing) and food expulsion, and a large increase in observed chewing. (Author/DLS)

Space Shuttle Projects

NASA Image and Video Library

2004-04-15

The Apollo program demonstrated that men could travel into space, perform useful tasks there, and return safely to Earth. But space had to be more accessible. This led to the development of the Space Shuttle. The Shuttle's major components are the orbiter spacecraft; the three main engines, with a combined thrust of more than 1.2 million pounds; the huge external tank (ET) that feeds the liquid hydrogen fuel and liquid oxygen oxidizer to the three main engines; and the two solid rocket boosters (SRBs), with their combined thrust of some 5.8 million pounds, that provide most of the power for the first two minutes of flight. Crucially involved with the Space Shuttle program virtually from its inception, the Marshall Space Flight Center (MSFC) played a leading role in the design, development, testing, and fabrication of many major Shuttle propulsion components.

Space Shuttle Drawing

NASA Technical Reports Server (NTRS)

2004-01-01

The Apollo program demonstrated that men could travel into space, perform useful tasks there, and return safely to Earth. But space had to be more accessible. This led to the development of the Space Shuttle. The Shuttle's major components are the orbiter spacecraft; the three main engines, with a combined thrust of more than 1.2 million pounds; the huge external tank (ET) that feeds the liquid hydrogen fuel and liquid oxygen oxidizer to the three main engines; and the two solid rocket boosters (SRBs), with their combined thrust of some 5.8 million pounds, that provide most of the power for the first two minutes of flight. Crucially involved with the Space Shuttle program virtually from its inception, the Marshall Space Flight Center (MSFC) played a leading role in the design, development, testing, and fabrication of many major Shuttle propulsion components.

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.

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

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.

Computational Investigations of Inboard Flow Separation and Mitigation Techniques on Multi-Megawatt Wind Turbines

NASA Astrophysics Data System (ADS)

Chow, Raymond

The aerodynamic characteristics of the NREL 5-MW rotor have been examined using a Reynolds-averaged Navier-Stokes method, OVERFLOW2. A comprehensive off-body grid independence study has been performed. A strong dependence on the size of the near-body wake grid has been found. Rapid diffusion of the wake appears to generate an overprediction of power and thrust. A large, continuous near-wake grid at minimum of two rotor diameters downstream of the rotor appears to be necessary for accurate predictions of near-body forces. The NREL 5-MW rotor demonstrates significant inboard flow separation up to 30% of span. This separation appears to be highly three-dimensional, with a significant amount of radial flow increasing the size of the separated region outboard. Both integrated aerodynamic coefficients and detailed wake structures for the baseline NREL 5-MW rotor are in excellent agreement with results by Riso at Uinfinity = 8 and 11 m/s. A simple, continuous full-chord fence was applied at the maximum chord location of the blade, within the region of separation. This non-optimized device reduced the boundary-layer cross-flow and resulting separation, and increased rotor power capture by 0.9% and 0.6% at U infinity = 8 and 11 m/s, respectively. Suction side only fences perform similarly in terms of power capture but reduce the increase in rotor thrust. Fence heights from 0.5% to 17.5% of the maximum chord all demonstrate some level of effectiveness, with fences (1-2.5%cmax) showing similar performance gains to taller fences with smaller penalties in thrust. Performance in terms of power capture is not very sensitive to spanwise location when placed within the separation region. Blunt trailing edge modifications to the inboard region of the blade showed a relatively significant effect on rotor power. Over a large range of trailing edge thicknesses from hTE = 10 to 25%c, power was found to increase by 1.4%. Thrust increased proportionally with the thicknesses examined, reaching a comparable increase of 1.4% by a trailing edge thickness of 15%c. Decreasing inboard twist only acted to increase thrust without increasing power capture any further at U infinity = 11 m/s. While increasing inboard blade twist decreased power, but decreased thrust at even a higher rate. Vortex generators were not successively configured to significantly improve power capture in this study. Two of the three configurations examined actually decreased power capture and increased the separation region. The results found in this study are not believed to be representative of a properly sized and located array of VGs. The presence of the nose cone and nacelle body at the hub of the rotor is found to have a minimal effect on the power and thrust of the overall rotor. The downstream wake structure however is changed by the nacelle, potentially useful for wake tailoring when turbines are closely spaced together.

Preliminary Characterization of the Altair Lunar Lander Slosh Dynamics and Some Implications for the Thrust Vector Control Design

NASA Technical Reports Server (NTRS)

Lee, Allan Y.; Strahan, Alan; Tanimoto, Rebekah; Casillas, Arturo

2010-01-01

This paper describes a conceptual design of the Thrust Vector Control (TVC) system and preliminary modeling of propellant slosh, for the Altair Lunar Lander. Altair is a vehicle element of the NASA Constellation Program aimed at returning humans to the moon. Guidance, Navigation, and Control (GN&C) is the measurement and control of spacecraft position, velocity, and attitude in support of mission objectives. One key GN&C function is the commanding of effectors that control attitude and impart delta V on the vehicle, utilizing both reaction control system (RCS) thrusters and throttling and TVC gimbaling of the vehicle main engine. Both the Altair descent and ascent modules carry fuel tanks. During thrusting maneuvers, the sloshing of liquid fuels in partially filled tanks can interact with the controlled system in such a way as to cause the overall system to be unstable. These fuel tanks must be properly placed, relative to the spacecraft's c.m., to avoid any unstable interactions. Following this will be a discussion of propellant slosh modeling work performed for the present vehicle configuration, including slosh frequency and participatory fluid mass predictions. Knowing the range of slosh mode frequencies over mission phases, the TVC bandwidth must be carefully selected so as not to excite the slosh modes at those frequencies. The likely need to increase the damping factor of slosh modes via baffles will also be discussed. To conclude, a discussion of operations procedures aimed at minimizing TVC-slosh interactions will be given.

Structural Health and Prognostics Management for Offshore Wind Plants; Final Report of Sandia R&D Activities.

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

Griffith, Daniel Todd

2015-04-01

This final report is a compilation of resear ch efforts - funded by the US Department of Energy Wind and Water Power Technolog ies Office over a four-year period from FY11 through FY14. The goals of this re search program were to develop and evaluate technical innovati ons with promise for maxi mizing revenues and reducing levelized cost of energy (LCOE) for offs hore wind plants - more specifically the goals of the Structural H ealth and Prognostics Management (SHPM) program were to reduce O&M costs and increase energy capture through use of SHPM-based technologies. A technology roadmap was devemore » loped at the start of the project to guide the research efforts. This roadmap identified and outlined six major research thrust areas each having five stages of ma turity. Research was conducted in each of these thrust areas, as documented throughout this report, although a major focus was on development of damage detection strategi es for the most frequent blade damage conditions and damage mitigation and life-exte nsion strategies via changes in turbine operations (smart loads management). Th e work summarized in this compilation report is the product of the work of many researchers. A summary of the major findings, status of the SHPM Technology Ro admap and recommendations for future work are also provided.« less

High sedimentation rates and thrust fault modulation: Insights from ocean drilling offshore the St. Elias Mountains, southern Alaska

NASA Astrophysics Data System (ADS)

Worthington, Lindsay L.; Daigle, Hugh; Clary, Wesley A.; Gulick, Sean P. S.; Montelli, Aleksandr

2018-02-01

The southern Alaskan margin offshore the St. Elias Mountains has experienced the highest recorded offshore sediment accumulation rates globally. Combined with high uplift rates, active convergence and extensive temperate glaciation, the margin provides a superb setting for evaluating competing influences of tectonic and surface processes on orogen development. We correlate results from Integrated Ocean Drilling Program (IODP) Expedition 341 Sites U1420 and U1421 with regional seismic data to determine the spatial and temporal evolution of the Pamplona Zone fold-thrust belt that forms the offshore St. Elias deformation front on the continental shelf. Our mapping shows that the pattern of active faulting changed from distributed across the shelf to localized away from the primary glacial depocenter over ∼300-780 kyrs, following an order-of-magnitude increase in sediment accumulation rates. Simple Coulomb stress calculations show that the suppression of faulting is partially controlled by the change in sediment accumulation rates which created a differential pore pressure regime between the underlying, faulted strata and the overlying, undeformed sediments.

Workshop on Jet Exhaust Noise Reduction for Tactical Aircraft - NASA Perspective

NASA Technical Reports Server (NTRS)

Huff, Dennis L.; Henderson, Brenda S.

2007-01-01

Jet noise from supersonic, high performance aircraft is a significant problem for takeoff and landing operations near air bases and aircraft carriers. As newer aircraft with higher thrust and performance are introduced, the noise tends to increase due to higher jet exhaust velocities. Jet noise has been a subject of research for over 55 years. Commercial subsonic aircraft benefit from changes to the engine cycle that reduce the exhaust velocities and result in significant noise reduction. Most of the research programs over the past few decades have concentrated on commercial aircraft. Progress has been made by introducing new engines with design features that reduce the noise. NASA has recently started a new program called "Fundamental Aeronautics" where three projects (subsonic fixed wing, subsonic rotary wing, and supersonics) address aircraft noise. For the supersonics project, a primary goal is to understand the underlying physics associated with jet noise so that improved noise prediction tools and noise reduction methods can be developed for a wide range of applications. Highlights from the supersonics project are presented including prediction methods for broadband shock noise, flow measurement methods, and noise reduction methods. Realistic expectations are presented based on past history that indicates significant jet noise reduction cannot be achieved without major changes to the engine cycle. NASA s past experience shows a few EPNdB (effective perceived noise level in decibels) can be achieved using low noise design features such as chevron nozzles. Minimal thrust loss can be expected with these nozzles (< 0.5%) and they may be retrofitted on existing engines. In the long term, it is desirable to use variable cycle engines that can be optimized for lower jet noise during takeoff operations and higher thrust for operational performance. It is also suggested that noise experts be included early in the design process for engine nozzle systems to participate in decisions that may impact the jet noise.

Perceived Noise Analysis for Offset Jets Applied to Commercial Supersonic Aircraft

NASA Technical Reports Server (NTRS)

Huff, Dennis L.; Henderson, Brenda S.; Berton, Jeffrey J.; Seidel, Jonathan A.

2016-01-01

A systems analysis was performed with experimental jet noise data, engine/aircraft performance codes and aircraft noise prediction codes to assess takeoff noise levels and mission range for conceptual supersonic commercial aircraft. A parametric study was done to identify viable engine cycles that meet NASA's N+2 goals for noise and performance. Model scale data from offset jets were used as input to the aircraft noise prediction code to determine the expected sound levels for the lateral certification point where jet noise dominates over all other noise sources. The noise predictions were used to determine the optimal orientation of the offset nozzles to minimize the noise at the lateral microphone location. An alternative takeoff procedure called "programmed lapse rate" was evaluated for noise reduction benefits. Results show there are two types of engines that provide acceptable mission range performance; one is a conventional mixed-flow turbofan and the other is a three-stream variable-cycle engine. Separate flow offset nozzles reduce the noise directed toward the thicker side of the outer flow stream, but have less benefit as the core nozzle pressure ratio is reduced. At the systems level for a three-engine N+2 aircraft with full throttle takeoff, there is a 1.4 EPNdB margin to Chapter 3 noise regulations predicted for the lateral certification point (assuming jet noise dominates). With a 10% reduction in thrust just after clearing the runway, the margin increases to 5.5 EPNdB. Margins to Chapter 4 and Chapter 14 levels will depend on the cumulative split between the three certification points, but it appears that low specific thrust engines with a 10% reduction in thrust (programmed lapse rate) can come close to meeting Chapter 14 noise levels. Further noise reduction is possible with engine oversizing and derated takeoff, but more detailed mission studies are needed to investigate the range impacts as well as the practical limits for safety and takeoff regulations.

Impact of emerging technologies on future combat aircraft agility

NASA Technical Reports Server (NTRS)

Nguyen, Luat T.; Gilert, William P.

1990-01-01

The foreseeable character of future within-visual-range air combat entails a degree of agility which calls for the integration of high-alpha aerodynamics, thrust vectoring, intimate pilot/vehicle interfaces, and advanced weapons/avionics suites, in prospective configurations. The primary technology-development programs currently contributing to these goals are presently discussed; they encompass the F-15 Short Takeoff and Landing/Maneuver Technology Demonstrator Program, the Enhanced Fighter Maneuverability Program, the High Angle-of-Attack Technology Program, and the X-29 Technology Demonstrator Program.

National Civil Applications Program: strategic plan vision for 2005

USGS Publications Warehouse

,

2004-01-01

The National Mapping Division (NMD) has developed this comprehensive strategic plan to chart the course of the National Civil Applications Program (NCAP) over the next 5 years. To meet the challenges of the future, the NCAP is changing its program emphases, methods of responding to customer needs, and business practices. The NCAP Strategic Plan identifies the new direction for the program through a series of strategic thrusts and goals for managers to use in formulating plans, establishing program emphases, and determining resource needs and allocations.

Orbit Transfer Rocket Engine Technology - 7.5K-LB Thrust Rocket Engine Preliminary Design

DTIC Science & Technology

1993-10-15

AND SPACE ADMINISTRATION October, 1993 r W NASA-Lewis Research Center Cleveland, Ohio 44135 94-08572 Contract Nc. NAS3-23773 Task B.7 and D.5 4I3’OA4 3 ...APPROACH 1 4.0 SUMMARY OF ACCOMPLISHMENTS 2 5.0 TECHNICAL DISCUSSIONS 3 6.0 PROGRAM WORK PLAN 5 6.1 Engine Analysis 5 6.2 Component Analysis 15 6.2.1...FIGURES Page Figure 1 Advanced Engine Studv Logic Diagram 4 Figure 2 Design Point Engine Pertormance at Full Thrust & MR = 6.0 7 Figure 3 Off-Design

Space shuttle orbit maneuvering engine, reusable thrust chamber program. Task 6: Data dump hot fuel element investigation

NASA Technical Reports Server (NTRS)

Nurick, W. H.

1974-01-01

An evaluation of reusable thrust chambers for the space shuttle orbit maneuvering engine was conducted. Tests were conducted using subscale injector hot-fire procedures for the injector configurations designed for a regenerative cooled engine. The effect of operating conditions and fuel temperature on combustion chamber performance was determined. Specific objectives of the evaluation were to examine the optimum like-doublet element geometry for operation at conditions consistent with a fuel regeneratively cooled engine (hot fuel, 200 to 250 F) and the sensitivity of the triplet injector element to hot fuels.

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.

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.

Saturn Apollo Program

NASA Image and Video Library

1964-12-01

At the Marshall Space Flight Center (MSFC), the fuel tank assembly for the Saturn V S-IC-T (static test stage) fuel tank assembly is mated to the liquid oxygen (LOX) tank in building 4705. This stage underwent numerous static firings at the newly-built S-IC Static Test Stand at the MSFC west test area. The S-IC (first) stage used five F-1 engines that produced a total thrust of 7,500,000 pounds as each engine produced 1,500,000 pounds of thrust. The S-IC stage lifted the Saturn V vehicle and Apollo spacecraft from the launch pad.

Simulating Operation of a Large Turbofan Engine

NASA Technical Reports Server (NTRS)

Litt, Jonathan S.; Frederick, Dean K.; DeCastro, Jonathan

2008-01-01

The Commercial Modular Aero- Propulsion System Simulation (C-MAPSS) is a computer program for simulating transient operation of a commercial turbofan engine that can generate as much as 90,000 lb (.0.4 MN) of thrust. It includes a power-management system that enables simulation of open- or closed-loop engine operation over a wide range of thrust levels throughout the full range of flight conditions. C-MAPSS provides the user with a set of tools for performing open- and closed-loop transient simulations and comparison of linear and non-linear models throughout its operating envelope, in an easy-to-use graphical environment.

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.

Solar Sail Propulsion for Interplanetary Cubesats

NASA Technical Reports Server (NTRS)

Johnson, Les; Sobey, Alex; Sykes, Kevin

2015-01-01

NASA is developing two small satellite missions as part of the Advanced Exploration Systems (AES) Program, both of which will use a solar sail to enable their scientific objectives. Solar sails use sunlight to propel vehicles through space by reflecting solar photons from a large, mirror-like sail made of a lightweight, highly reflective material. This continuous photon pressure provides propellantless thrust, allowing for very high (Delta)V maneuvers on long-duration, deep space exploration. Since reflected light produces thrust, solar sails require no onboard propellant. Solar sail technology is rapidly maturing for space propulsion applications within NASA and around the world.

Space shuttle orbit maneuvering engine reusable thrust chamber

NASA Technical Reports Server (NTRS)

1972-01-01

A data dump is presented containing space shuttle orbiter maneuvering engine performance, weight, envelope, and interface pressure requirements for candidate propellant combinations (NTO/MMH, NTO50-50, LOX/MMH, LOX/50-50, LOX/N2H4, LOX/C3H8, and LOX/RP-1) and cooling concepts (regenerative and dump/film). These data are presented parametrically for the thrust, chamber pressure, nozzle expansion ratio, and engine mixture ratio ranges of interest. Also included is information describing sensitivity to system changes; reliability, maintainability and safety; development programs and associated critical technology areas; engine cost comparisons during development and operation; and ecological effects.

Longhorn Business Jets

NASA Technical Reports Server (NTRS)

1980-01-01

Developed in NASA's Aircraft Energy Efficiency program and manufactured by Gates Learjet Corporation, the winglet is an aerodynamic innovation designed to reduce fuel consumption and improve airplane performance. Winglets are lifting surfaces designed to operate in the "vortex" or air whirlpool which occurs at an airplane's wingtip. Complex flow of air around wingtip creates drag which retards the plane's progress. Winglet reduces strength of vortex and thereby reduces strength of drag. Additionally, winglet generates its own lift, producing forward thrust in the manner of a boat's sail. Combination of reduced drag and additional thrust adds up to significant improvement in fuel efficiency.

Computer Tomography Analysis of Fastrac Composite Thrust Chamber Assemblies

NASA Technical Reports Server (NTRS)

Beshears, Ronald D.

2000-01-01

Computed tomography (CT) inspection has been integrated into the production process for NASA's Fastrac composite thrust chamber assemblies (TCAs). CT has been proven to be uniquely qualified to detect the known critical flaw for these nozzles, liner cracks that are adjacent to debonds between the liner and overwrap. CT is also being used as a process monitoring tool through analysis of low density indications in the nozzle overwraps. 3d reconstruction of CT images to produce models of flawed areas is being used to give program engineers better insight into the location and nature of nozzle flaws.

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.

Effect of Background Pressure on the Performance and Plume of the HiVHAc Hall Thruster

NASA Technical Reports Server (NTRS)

Huang, Wensheng; Kamhawi, Hani; Haag, Thomas

2013-01-01

During the Single String Integration Test of the NASA HiVHAc Hall thruster, a number of plasma diagnostics were implemented to study the effect of varying facility background pressure on thruster operation. These diagnostics include thrust stand, Faraday probe, ExB probe, and retarding potential analyzer. The test results indicated a rise in thrust and discharge current with background pressure. There was also a decrease in ion energy per charge, an increase in multiply-charged species production, a decrease in plume divergence, and a decrease in ion beam current with increasing background pressure. A simplified ingestion model was applied to determine the maximum acceptable background pressure for thrust measurement. The maximum acceptable ingestion percentage was found to be around 1%. Examination of the diagnostics results suggest the ionization and acceleration zones of the thruster were shifting upstream with increasing background pressure.

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.

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.

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.

Lower extremity thrust and non-thrust joint mobilization for patellofemoral pain syndrome: a case report

PubMed Central

Simpson, Brad G; Simon, Corey B

2014-01-01

A 40-year old female presented to physical therapy with a one-year history of insidious right anteromedial and anterolateral knee pain. Additionally, the patient had a history of multiple lateral ankle sprains bilaterally, the last sprain occurring on the right ankle 1 year prior to the onset of knee pain. The patient was evaluated and given a physical therapy diagnosis of patellofemoral pain syndrome (PFPS), with associated talocrural and tibiofemoral joint hypomobility limiting ankle dorsiflexion and knee extension, respectively. Treatment included a high-velocity low amplitude thrust manipulation to the talocrural joint, which helped restore normal ankle dorsiflexion range of motion. The patient also received tibiofemoral joint non-thrust manual therapy to regain normal knee extension mobility prior to implementing further functional progression exercises to her home program (HEP). This case report highlights the importance of a detailed evaluation of knee and ankle joint mobility in patients presenting with anterior knee pain. Further, manual physical therapy to the lower extremity was found to be successful in restoring normal movement patterns and pain-free function in a patient with chronic anterior knee pain. PMID:24976753

Thrust Augmentation Measurements for a Pulse Detonation Engine Driven Ejector

NASA Technical Reports Server (NTRS)

Pal, S.; Santoro, Robert J.; Shehadeh, R.; Saretto, S.; Lee, S.-Y.

2005-01-01

Thrust augmentation results of an ongoing study of pulse detonation engine driven ejectors are presented and discussed. The experiments were conducted using a pulse detonation engine (PDE) setup with various ejector configurations. The PDE used in these experiments utilizes ethylene (C2H4) as the fuel, and an equi-molar mixture of oxygen and nitrogen as the oxidizer at an equivalence ratio of one. High fidelity thrust measurements were made using an integrated spring damper system. The baseline thrust of the PDE engine was first measured and agrees with experimental and modeling results found in the literature. Thrust augmentation measurements were then made for constant diameter ejectors. The parameter space for the study included ejector length, PDE tube exit to ejector tube inlet overlap distance, and straight versus rounded ejector inlets. The relationship between the thrust augmentation results and various physical phenomena is described. To further understand the flow dynamics, shadow graph images of the exiting shock wave front from the PDE were also made. For the studied parameter space, the results showed a maximum augmentation of 40%. Further increase in augmentation is possible if the geometry of the ejector is tailored, a topic currently studied by numerous groups in the field.

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.

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.

Flat-ramp vs. convex-concave thrust geometries in a deformable hanging wall: new insights from analogue modeling experiments

NASA Astrophysics Data System (ADS)

Almeida, Pedro; Tomas, Ricardo; Rosas, Filipe; Duarte, Joao; Terrinha, Pedro

2015-04-01

Different modes of strain accommodation affecting a deformable hanging-wall in a flat-ramp-flat thrust system were previously addressed through several (sandbox) analog modeling studies, focusing on the influence of different variables, such as: a) thrust ramp dip angle and friction (Bonini et al, 2000); b) prescribed thickness of the hanging-wall (Koy and Maillot, 2007); and c) sin-thrust erosion (compensating for topographic thrust edification, e.g. Persson and Sokoutis, 2002). In the present work we reproduce the same experimental procedure to investigate the influence of two different parameters on hanging-wall deformation: 1) the geometry of the thrusting surface; and 2) the absence of a velocity discontinuity (VD) that is always present in previous similar analogue modeling studies. Considering the first variable we use two end member ramp geometries, flat-ramp-flat and convex-concave, to understand the control exerted by the abrupt ramp edges in the hanging-wall stress-strain distribution, comparing the obtain results with the situation in which such edge singularities are absent (convex-concave thrust ramp). Considering the second investigated parameter, our motivation was the recognition that the VD found in the different analogue modeling settings simply does not exist in nature, despite the fact that it has a major influence on strain accommodation in the deformable hanging-wall. We thus eliminate such apparatus artifact from our models and compare the obtained results with the previous ones. Our preliminary results suggest that both investigated variables play a non-negligible role on the structural style characterizing the hanging-wall deformation of convergent tectonic settings were such thrust-ramp systems were recognized. Acknowledgments This work was sponsored by the Fundação para a Ciência e a Tecnologia (FCT) through project MODELINK EXPL/GEO-GEO/0714/2013. Pedro Almeida wants to thank to FCT for the Ph.D. grant (SFRH/BD/52556/2014) under the Doctoral Program EarthSystems in IDL/UL. References Bonini, M., Sokoutis, D., Mulugeta, G., Katrivanos, E. (2000) - Modelling hanging wall accommodation above rigid thrust ramps. Journal of Structural Geology, 22, pp. 1165-1179. Persson, K. & Sokoutis, D (2002) - Analogue models of orogenic wedges controlled by erosion. Tectonophysics, 356, pp. 323- 336. Koy, H. & Bertrand, M. (2007) - Tectonic thickening of hanging-wall units over a ramp.Journal of Structural Geology, 29, pp. 924-932.

Strategic plan, 1985

NASA Technical Reports Server (NTRS)

1985-01-01

The Lewis Strategic Plan was updated for 1985 and beyond. Major programs for the space station, the advanced turboprop, the Advanced Communications Technology Satellite (ACTS), and the Altitude Wind Tunnel were begun or greatly expanded during 1984. In parallel, The Lewis aeropropulsion research and technology program was extensively evaluated and reviewed; a reduced and reoriented program emerged. The thrusts and implementation plans for these programs are described as they pertain to the individual directorates. Other key accomplishments and plans are summarized.

Program definition and assessment overview. [for thermal energy storage project management

NASA Technical Reports Server (NTRS)

Gordon, L. H.

1980-01-01

The implementation of a program level assessment of thermal energy storage technology thrusts for the near and far term to assure overall coherent energy storage program is considered. The identification and definition of potential thermal energy storage applications, definition of technology requirements, and appropriate market sectors are discussed along with the necessary coordination, planning, and preparation associated with program reviews, workshops, multi-year plans and annual operating plans for the major laboratory tasks.

Apollo guidance, navigation and control: Guidance system operations plan for manned CM earth orbital and lunar missions using Program COLOSSUS 3. Section 3: Digital autopilots (revision 14)

NASA Technical Reports Server (NTRS)

1972-01-01

Digital autopilots for the manned command module earth orbital and lunar missions using program COLOSSUS 3 are discussed. Subjects presented are: (1) reaction control system digital autopilot, (2) thrust vector control autopilot, (3) entry autopilot and mission control programs, (4) takeover of Saturn steering, and (5) coasting flight attitude maneuver routine.

NASA's Current Directions in the CETDP Micro-Technology Thrust Area

NASA Technical Reports Server (NTRS)

Stocky, J.

2000-01-01

NASA's program in micro-technologies seeks to develop the advanced technologies needed to reduce the mass of Earth-orbiting and deep-space spacecraft by several orders of magnitude over the next decade.

X-31 post-stall envelope expansion and tactical utility testing

NASA Technical Reports Server (NTRS)

Canter, Dave

1994-01-01

Technical and nontechnical lessons learned from the X-31 aircraft program are described in this viewgraph presentation. The tactical utility of high angle of attack flight and thrust vector control is discussed.

Advanced Concept

NASA Image and Video Library

2008-02-15

Testing of the Ascent Thrust Vector Control System in support of the Ares 1-X program at the Marshall Space Flight Center in Huntsville, Alabama. This image is extracted from a high definition video file and is the highest resolution available

Accelerating slip rates on the puente hills blind thrust fault system beneath metropolitan Los Angeles, California, USA

USGS Publications Warehouse

Bergen, Kristian J.; Shaw, John H.; Leon, Lorraine A.; Dolan, James F.; Pratt, Thomas L.; Ponti, Daniel J.; Morrow, Eric; Barrera, Wendy; Rhodes, Edward J.; Murari, Madhav K.; Owen, Lewis A.

2017-01-01

Slip rates represent the average displacement across a fault over time and are essential to estimating earthquake recurrence for proba-bilistic seismic hazard assessments. We demonstrate that the slip rate on the western segment of the Puente Hills blind thrust fault system, which is beneath downtown Los Angeles, California (USA), has accel-erated from ~0.22 mm/yr in the late Pleistocene to ~1.33 mm/yr in the Holocene. Our analysis is based on syntectonic strata derived from the Los Angeles River, which has continuously buried a fold scarp above the blind thrust. Slip on the fault beneath our field site began during the late-middle Pleistocene and progressively increased into the Holocene. This increase in rate implies that the magnitudes and/or the frequency of earthquakes on this fault segment have increased over time. This challenges the characteristic earthquake model and presents an evolving and potentially increasing seismic hazard to metropolitan Los Angeles.

NASA programs in advanced sensors and measurement technology for aeronautical applications

NASA Astrophysics Data System (ADS)

Conway, Bruce A.

NASA involvement in the development, implementation, and experimental use of advanced aeronautical sensors and measurement technologies is presently discussed within the framework of specific NASA research centers' activities. The technology thrusts are in the fields of high temperature strain gages and microphones, laser light-sheet flow visualization, LTA, LDV, and LDA, tunable laser-based aviation meteorology, and fiber-optic CARS measurements. IR thermography and close-range photogrammetry are undergoing substantial updating and application. It is expected that 'smart' sensors will be increasingly widely used, especially in conjunction with smart structures in aircraft and spacecraft.

OEO Programs for Migrant and Seasonal Farm Workers. Title IIIB EOA Projects.

ERIC Educational Resources Information Center

Office of Economic Opportunity, Washington, DC.

The major thrust of Title III-B (Economic Opportunity Act) programs is to prepare migrant and seasonal farm workers for upgraded jobs and to prepare entire farm worker families for alternatives to farm work if they so desire. These alternatives, supported through projects for adult heads of households, emphasize skill training in preparation for…

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.

Propeller thrust analysis using Prandtl's lifting line theory, a comparison between the experimental thrust and the thrust predicted by Prandtl's lifting line theory

NASA Astrophysics Data System (ADS)

Kesler, Steven R.

The lifting line theory was first developed by Prandtl and was used primarily on analysis of airplane wings. Though the theory is about one hundred years old, it is still used in the initial calculations to find the lift of a wing. The question that guided this thesis was, "How close does Prandtl's lifting line theory predict the thrust of a propeller?" In order to answer this question, an experiment was designed that measured the thrust of a propeller for different speeds. The measured thrust was compared to what the theory predicted. In order to do this experiment and analysis, a propeller needed to be used. A walnut wood ultralight propeller was chosen that had a 1.30 meter (51 inches) length from tip to tip. In this thesis, Prandtl's lifting line theory was modified to account for the different incoming velocity depending on the radial position of the airfoil. A modified equation was used to reflect these differences. A working code was developed based on this modified equation. A testing rig was built that allowed the propeller to be rotated at high speeds while measuring the thrust. During testing, the rotational speed of the propeller ranged from 13-43 rotations per second. The thrust from the propeller was measured at different speeds and ranged from 16-33 Newton's. The test data were then compared to the theoretical results obtained from the lifting line code. A plot in Chapter 5 (the results section) shows the theoretical vs. actual thrust for different rotational speeds. The theory over predicted the actual thrust of the propeller. Depending on the rotational speed, the error was: at low speeds 36%, at low to moderate speeds 84%, and at high speeds the error increased to 195%. Different reasons for these errors are discussed.

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.

The ISR Asymmetrical Capacitor Thruster: Experimental Results and Improved Designs

NASA Technical Reports Server (NTRS)

Canning, Francis X.; Cole, John; Campbell, Jonathan; Winet, Edwin

2004-01-01

A variety of Asymmetrical Capacitor Thrusters has been built and tested at the Institute for Scientific Research (ISR). The thrust produced for various voltages has been measured, along with the current flowing, both between the plates and to ground through the air (or other gas). VHF radiation due to Trichel pulses has been measured and correlated over short time scales to the current flowing through the capacitor. A series of designs were tested, which were increasingly efficient. Sharp features on the leading capacitor surface (e.g., a disk) were found to increase the thrust. Surprisingly, combining that with sharp wires on the trailing edge of the device produced the largest thrust. Tests were performed for both polarizations of the applied voltage, and for grounding one or the other capacitor plate. In general (but not always) it was found that the direction of the thrust depended on the asymmetry of the capacitor rather than on the polarization of the voltage. While no force was measured in a vacuum, some suggested design changes are given for operation in reduced pressures.

Yakataga fold-and-thrust belt: Structural geometry and tectonic implications of a small continental collision zone

NASA Astrophysics Data System (ADS)

Wallace, Wesley K.

Collision of the Yakutat terrane with southern Alaska created a collisional fold-and-thrust belt along the Pacific-North America plate boundary. This southerner fold-and-thrust belt formed within continental sedimentary rocks but with the seaward vergence and tectonic position typical of an accretionary wedge. Northward exposure of progressively older rocks reflects that the fold-and-thrust belt forms a southward-tapered orogenic wedge that increases northward in structural relief and depth of erosion. Narrow, sharp anticlines separate wider, flat-bottomed synclines. Relatively steep thrust faults commonly cut the forelimbs of anticlines. Fold shortening and fault displacement both generally increase northward, whereas fault dip generally decreases northward. The coal-bearing lower part of the sedimentary section serves as a detachment for both folds and thrust faults. The folded and faulted sedimentary section defines a regional south dip of about 8°. The structural relief combined with the low magnitude of shortening of the sedimentary section suggest that the underlying basement is structurally thickened. I propose a new interpretation in which this thickening was accommodated by a passive-roof duplex with basement horses that are separated from the overlying folded and thrust-faulted sedimentary cover by a roof thrust with a backthrust sense of motion. Basement horses are ˜7 km thick, based on the thickness between the inferred roof thrust and the top of the basement in offshore seismic reflection data. This thickness is consistent with the depth of the zone of seismicity onshore. The inferred zone of detachment and imbrication of basement corresponds with the area of surface exposure of the fold-and-thrust belt within the Yakutat terrane and with the Wrangell subduction zone and arc farther landward. By contrast, to the west, the crust of the Yakutat terrane has been carried down a subduction zone that extends far landward with a gentle dip, corresponding with a gap in arc magmatism, anomalous topography, and the rupture zone of the 1964 great southern Alaska earthquake. I suggest that, to the east, detachment and imbrication of basement combined with coupling in the fold-and-thrust belt allowed the delaminated dense mantle lithosphere to subduct with a steeper dip than to the west, where buoyant Yakutat terrane crust remains attached to the subducted lithosphere. According to this interpretation, the Wrangell subduction zone is lithosphere of the Yakutat terrane, not Pacific Ocean lithosphere subducted beneath the Yakutat terrane. The Pacific-North America plate boundary would be within the northern deformed part of the Yakutat terrane, not along the boundary between the undeformed southern part of the Yakutat terrane and oceanic crust of the Pacific Ocean. The plate boundary is an evolving zone of distributed deformation in which most of the convergent component has been accommodated within the fold-and-thrust belt south of the northern boundary of the Yakutat terrane, the Chugach-St. Elias thrust fault, and most of the right-lateral component likely has been accommodated on the Bagley Icefield fault just to the north.

X-31 in flight - Double Reversal

NASA Technical Reports Server (NTRS)

1995-01-01

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

X-31 in flight - Post Stall Maneuver

NASA Technical Reports Server (NTRS)

1995-01-01

Two X-31 Enhanced Fighter Maneuverability (EFM) demonstrators were flown at Rockwell International's Palmdale, Calif., facility and the NASA Dryden Flight Research Center, Edwards, Calif., to obtain data that may apply to the design of highly-maneuverable next-generation fighters. The program had its first flight on Oct. 11, 1990, in Palmdale; it ended in June 1995. The X-31 program demonstrated the value of thrust vectoring (directing engine exhaust flow) coupled with advanced flight control systems, to provide controlled flight during close-in air combat at very high angles of attack. The result of this increased maneuverability is a significant advantage over conventional fighters. 'Angle-of-attack' (alpha) is an engineering term to describe the angle of an aircraft's body and wings relative to its actual flight path. During maneuvers, pilots often fly at extreme angles of attack -- with the nose pitched up while the aircraft continues in its original direction. This can lead to loss of control and result in the loss of the aircraft, pilot or both. Three thrust vectoring paddles made of graphite epoxy mounted on the X-31's exhaust nozzle directed the exhaust flow to provide control in pitch (up and down) and yaw (right and left) to improve control. The paddles can sustain heat of up to 1,500 degrees centigrade for extended periods of time. In addition the X-31s were configured with movable forward canards and fixed aft strakes. The canards were small wing-like structures set on the wing line between the nose and the leading edge of the wing. The strakes were set on the same line between the trailing edge of the wing and the engine exhaust. Both supplyied additional control in tight maneuvering situations. The X-31 research program produced technical data at high angles of attack. This information is giving engineers and aircraft designers a better understanding of aerodynamics, effectiveness of flight controls and thrust vectoring, and airflow phenomena at high angles of attack. This is expected to lead to design methods providing better maneuverability in future high performance aircraft and make them safer to fly. An international test organization of about 110 people, managed by the Advanced Research Projects Agency (ARPA), conducted the flight operations at Dryden, to which flight research was moved in February 1992 at the request of the Advanced Research Projects Agency (ARPA). In addition to ARPA and NASA, the International Test Organization (ITO) included the U.S. Navy, the U.S. Air Force, Rockwell International, the Federal Republic of Germany, and Daimler-Benz Aerospace (formerly Messerschmitt-Bolkow-Blohm and Deutsche Aerospace). NASA was responsible for flight research operations, aircraft maintenance, and research engineering once the program moved to Dryden. The No. 1 X-31 aircraft was lost in an accident Jan. 19, 1995. The pilot, Karl Heinz-Lang, of the Federal Republic of Germany, ejected safely before the aircraft crashed in an unpopulated desert area just north of Edwards. The X-31 program logged an X-plane record of 580 flights during the program, including 555 research missions and 21 in Europe for the 1995 Paris Air Show. A total of 14 pilots representing all agencies of the ITO flew the aircraft. This 34-second movie clip shows the aircraft as it slides backwards, thrust vectoring the tail over the top, turning the stall into a loop in which the aircraft then reverses it's heading and resumes level flight.

X-31 in flight - Herbst Turn

NASA Technical Reports Server (NTRS)

1995-01-01

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

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.

Green Propellant Test Capabilities of the Altitude Combustion Stand at the NASA Glenn Research Center

NASA Technical Reports Server (NTRS)

Kubiak, Jonathan M.; Arnett, Lori A.

2016-01-01

The NASA Glenn Research Center (GRC) is committed to providing simulated altitude rocket test capabilities to NASA programs, other government agencies, private industry partners, and academic partners. A primary facility to support those needs is the Altitude Combustion Stand (ACS). ACS provides the capability to test combustion components at a simulated altitude up to 100,000 ft. (approx.0.2 psia/10 Torr) through a nitrogen-driven ejector system. The facility is equipped with an axial thrust stand, gaseous and cryogenic liquid propellant feed systems, data acquisition system with up to 1000 Hz recording, and automated facility control system. Propellant capabilities include gaseous and liquid hydrogen, gaseous and liquid oxygen, and liquid methane. A water-cooled diffuser, exhaust spray cooling chamber, and multi-stage ejector systems can enable run times up to 180 seconds to 16 minutes. The system can accommodate engines up to 2000-lbf thrust, liquid propellant supply pressures up to 1800 psia, and test at the component level. Engines can also be fired at sea level if needed. The NASA GRC is in the process of modifying ACS capabilities to enable the testing of green propellant (GP) thrusters and components. Green propellants are actively being explored throughout government and industry as a non-toxic replacement to hydrazine monopropellants for applications such as reaction control systems or small spacecraft main propulsion systems. These propellants offer increased performance and cost savings over hydrazine. The modification of ACS is intended to enable testing of a wide range of green propellant engines for research and qualification-like testing applications. Once complete, ACS will have the capability to test green propellant engines up to 880 N in thrust, thermally condition the green propellants, provide test durations up to 60 minutes depending on thrust class, provide high speed control and data acquisition, as well as provide advanced imaging and diagnostics such as infrared (IR) imaging.

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.

Thrust Chamber Material Technology Program

DTIC Science & Technology

1989-03-01

about 5 percent lower than those of NASA.Z. Alloy 1035 was essentially equivalent to NASA-Z. However, Alloy 1032, containing titanium, exhibited such...Characterization Methods .................................... 33 5 Evaluation Methods ......................................................... 37 III...115 5 Discussion ..................................................................... 120 V ALLOY CHARACTERIZATION

Energy Efficient Engine exhaust mixer model technology report addendum; phase 3 test program

NASA Technical Reports Server (NTRS)

Larkin, M. J.; Blatt, J. R.

1984-01-01

The Phase 3 exhaust mixer test program was conducted to explore the trends established during previous Phases 1 and 2. Combinations of mixer design parameters were tested. Phase 3 testing showed that the best performance achievable within tailpipe length and diameter constraints is 2.55 percent better than an optimized separate flow base line. A reduced penetration design achieved about the same overall performance level at a substantially lower level of excess pressure loss but with a small reduction in mixing. To improve reliability of the data, the hot and cold flow thrust coefficient analysis used in Phases 1 and 2 was augmented by calculating percent mixing from traverse data. Relative change in percent mixing between configurations was determined from thrust and flow coefficient increments. The calculation procedure developed was found to be a useful tool in assessing mixer performance. Detailed flow field data were obtained to facilitate calibration of computer codes.

Lessons Learned with Metallized Gelled Propellants

NASA Technical Reports Server (NTRS)

1996-01-01

During testing of metallized gelled propellants in a rocket engine, many changes had to be made to the normal test program for traditional liquid propellants. The lessons learned during the testing and the solutions for many of the new operational conditions posed with gelled fuels will help future programs run more smoothly. The major factors that influenced the success of the testing were propellant settling, piston-cylinder tank operation, control of self pressurization, capture of metal oxide particles, and a gelled-fuel protective layer. In these ongoing rocket combustion experiments at the NASA Lewis Research Center, metallized, gelled liquid propellants are used in a small modular engine that produces 30 to 40 lb of thrust. Traditional liquid RP-1 and gelled RP-1 with 0-, 5-, and 55-wt% loadings of aluminum are used with gaseous oxygen as the oxidizer. The figure compares the thrust chamber efficiencies of different engines.

Investigation of a hydrostatic azimuth thrust bearing for a large steerable antenna

NASA Technical Reports Server (NTRS)

Rumbarger, J.; Castelli, V.; Rippel, H.

1972-01-01

The problems inherent in the design and construction of a hydrostatic azimuth thrust bearing for a tracking antenna of very large size were studied. For a load of 48,000,000 lbs., it is concluded that the hydrostatic bearing concept is feasible, provided that a particular multiple pad arrangement, high oil viscosity, and a particular load spreading arrangement are used. Presently available computer programs and techniques are deemed to be adequate for a good portion of the design job but new integrated programs will have to be developed in the area of the computation of the deflections of the supporting bearing structure. Experimental studies might also be indicated to ascertain the life characteristics of grouting under cyclic loading, and the optimization of hydraulic circuits and pipe sizes to insure the long life operation of pumps with high viscosity oil while avoiding cavitation.

X-31 Loaded in C-5 Cargo Bay

NASA Technical Reports Server (NTRS)

1995-01-01

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

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.

Office of Science and Technology&International Year EndReport - 2005

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

Bodvarsson, G.S.

2005-10-27

Source Term, Materials Performance, Radionuclide Getters, Natural Barriers, and Advanced Technologies, a brief introduction in each section describes the overall organization and goals of each program area. All of these areas have great potential for improving our understanding of the safety performance of the proposed Yucca Mountain repository, as processes within these areas are generally very conservatively represented in the Total System Performance Assessment. In addition, some of the technology thrust areas in particular may enhance system efficiency and reduce risk to workers. Thus, rather modest effort in the S&T Program could lead to large savings in the lifetime repositorymore » total cost and significantly enhanced understanding of the behavior of the proposed Yucca Mountain repository, without safety being compromised, and in some instances being enhanced. An overall strength of the S&T Program is the significant amount of integration that has already been achieved after two years of research. As an example (illustrated in Figure 1), our understanding of the behavior of the total waste isolation system has been enhanced through integration of the Source Term, Materials Performance, and Natural Barriers Thrust areas. All three thrust areas contribute to the integration of different processes in the in-drift environment. These processes include seepage into the drift, dust accumulation on the waste package, brine formation and precipitation on the waste package, mass transfer through the fuel cladding, changes in the seepage-water chemical composition, and transport of released radionuclides through the invert and natural barriers. During FY2005, each of our program areas assembled a team of external experts to conduct an independent review of their respective projects, research directions, and emphasis. In addition, the S&T Program as a whole was independently reviewed by the S&T Programmatic Evaluation Panel. As a result of these reviews, adjustments to the S&T Program will be implemented in FY2006 to ensure that the Program is properly aligned with OCRWM's priorities. Also during FY2005, several programmatic documents were published, including the Science and Technology Program Strategic Plan, the Science and Technology Program Management Plan, and the Science and Technology Program Plan. These and other communication products are available on the OCRWM web site under the Science and Technology section (http://www.ocrwm.doe.gov/osti/index.shtml).« less

Nasa's Emerging Productivity Program

NASA Technical Reports Server (NTRS)

Braunstein, D. R.

1984-01-01

The goals, membership, and organizational structure of the NASA Productivity Steering Committee are described as well as steps taken to make NASA a leader in the development and application of productivity and quality concepts at every level of agency management. The overall strategy for the Productivity Improvement and Quality Enhancement (PIQE) Program is through employee involvement, both civil servant and contractor, in all phases of agency-wide activity. Elements of the PIQE program and initial thrusts are examined.

Particle Swarm Optimization of Low-Thrust, Geocentric-to-Halo-Orbit Transfers

NASA Astrophysics Data System (ADS)

Abraham, Andrew J.

Missions to Lagrange points are becoming increasingly popular amongst spacecraft mission planners. Lagrange points are locations in space where the gravity force from two bodies, and the centrifugal force acting on a third body, cancel. To date, all spacecraft that have visited a Lagrange point have done so using high-thrust, chemical propulsion. Due to the increasing availability of low-thrust (high efficiency) propulsive devices, and their increasing capability in terms of fuel efficiency and instantaneous thrust, it has now become possible for a spacecraft to reach a Lagrange point orbit without the aid of chemical propellant. While at any given time there are many paths for a low-thrust trajectory to take, only one is optimal. The traditional approach to spacecraft trajectory optimization utilizes some form of gradient-based algorithm. While these algorithms offer numerous advantages, they also have a few significant shortcomings. The three most significant shortcomings are: (1) the fact that an initial guess solution is required to initialize the algorithm, (2) the radius of convergence can be quite small and can allow the algorithm to become trapped in local minima, and (3) gradient information is not always assessable nor always trustworthy for a given problem. To avoid these problems, this dissertation is focused on optimizing a low-thrust transfer trajectory from a geocentric orbit to an Earth-Moon, L1, Lagrange point orbit using the method of Particle Swarm Optimization (PSO). The PSO method is an evolutionary heuristic that was originally written to model birds swarming to locate hidden food sources. This PSO method will enable the exploration of the invariant stable manifold of the target Lagrange point orbit in an effort to optimize the spacecraft's low-thrust trajectory. Examples of these optimized trajectories are presented and contrasted with those found using traditional, gradient-based approaches. In summary, the results of this dissertation find that the PSO method does, indeed, successfully optimize the low-thrust trajectory transfer problem without the need for initial guessing. Furthermore, a two-degree-of-freedom PSO problem formulation significantly outperformed a one-degree-of-freedom formulation by at least an order of magnitude, in terms of CPU time. Finally, the PSO method is also used to solve a traditional, two-burn, impulsive transfer to a Lagrange point orbit using a hybrid optimization algorithm that incorporates a gradient-based shooting algorithm as a pre-optimizer. Surprisingly, the results of this study show that "fast" transfers outperform "slow" transfers in terms of both Deltav and time of flight.

Localized fluid discharge in subduction zones: Insights from tension veins around an ancient megasplay fault (Nobeoka Thrust, SW Japan)

NASA Astrophysics Data System (ADS)

Otsubo, M.; Hardebeck, J.; Miyakawa, A.; Yamaguchi, A.; Kimura, G.

2017-12-01

Fluid-rock interactions along seismogenic faults are of great importance to understand fault mechanics. The fluid loss by the formation of mode I cracks (tension cracks) increases the fault strength and creates drainage asperities along the plate interface (Sibson, 2013, Tectonophysics). The Nobeoka Thrust, in southwestern Japan, is an on-land example of an ancient megasplay fault and provides an excellent record of deformation and fluid flow at seismogenic depths of a subduction zone (Kondo et al., 2005, Tectonics). We focus on (1) Pore fluid pressure loss, (2) Amount of fault strength recovery, and (3) Fluid circulation by the formation of mode I cracks in the post-seismic period around the fault zone of the Nobeoka Thrust. Many quartz veins that filled mode I crack at the coastal outcrops suggest a normal faulting stress regime after faulting of the Nobeoka Thrust (Otsubo et al., 2016, Island Arc). We estimated the decrease of the pore fluid pressure by the formation of the mode I cracks around the Nobeoka Thrust in the post-seismic period. When the pore fluid pressure exceeds σ3, veins filling mode I cracks are constructed (Jolly and Sanderson, 1997, Jour. Struct. Geol.). We call the pore fluid pressure that exceeds σ3 "pore fluid over pressure". The differential stress in the post-seismic period and the driving pore fluid pressure ratio P* (P* = (Pf - σ3) / (σ1 - σ3), Pf: pore fluid pressure) are parameters to estimate the pore fluid over pressure. In the case of the Nobeoka Thrust (P* = 0.4, Otsubo et al., 2016, Island Arc), the pore fluid over pressure is up to 20 MPa (assuming tensile strength = 10 MPa). 20 MPa is equivalent to <10% of the total pore fluid pressure around the Nobeoka Thrust (depth = 10 km, density = 2.7 kg/m3). When the pore fluid pressure decreases by 4%, the normalized pore pressure ratio λ* (λ* = (Pf - Ph) / (Pl - Ph), Pl: lithostatic pressure; Ph: hydrostatic pressure) changes from 0.95 to 0.86. In the case of the Nobeoka Thrust, the fault strength can increase by up to 10 MPa (assuming frictional coefficient = 0.6). 10 MPa is almost equivalent to the stress drop values in large trench type earthquakes. Hence, we suggest that the fluid loss caused by the formation of mode I cracks in the post-seismic period may play an important role by increasing frictional strength along the megasplay fault.

Temperature micro-mapping and redox conditions of a chlorite zoning pattern in green-schist facies fault zone

NASA Astrophysics Data System (ADS)

Trincal, Vincent; Lanari, Pierre; Lacroix, Brice; Buatier, Martine D.; Charpentier, Delphine; Labaume, Pierre; Muñoz, Manuel

2014-05-01

Faults are major discontinuities driving fluid flows and playing a major role in precipitation of ore deposits. Mineral paragenesis and crystal chemistry depend on Temperature (T) condition, fluid composition but also on the redox environment of precipitation. The studied samples come from the Pic de Port Vieux thrust sheet, a minor thrust sheet associated to Gavarnie thrust fault zone (Central Pyrenees). The Pic de Port Vieux Thrust sheet comprises a 1-20 meter thick layer of Triassic red beds and mylonitized Cretaceous limestone. The thrust sheet is affected by faults and cleavage; the other important deformation product is a set of veins filled by quartz and chlorite. Microstructural and mineralogical investigations were performed based on the previous work of Grant (1992). The crystallization of chlorite is syn-tectonic and strongly controlled by the fluid circulation during the Gavarnie thrust sheet emplacement. Chlorite precipitated in extension veins, crack-seal shear veins or in open cavities. The chlorite filling the open cavities occurs as pseudo-uniaxial plates arranged in rosette-shaped aggregates. These aggregates appear to have developed as a result of radial growth of the chlorite platelets. According to point and microprobe X-ray images, these chlorites display oscillatory chemical zoning patterns with alternating iron rich and magnesium rich bands. The chlorite composition ranges from Fe rich pole (Si2.62Al1.38O10(Al1.47Fe1.87Mg2.61)6(OH)8) to Mg rich pole (Si2.68Al1.31O10(Al1.45Fe1.41Mg3.06)6(OH)8). In metamorphic rocks, zoning pattern or rimmed minerals results for varying P or T conditions and can be used to unravel the P-T history of the sample. In the present study, temperature maps are derived from standardized microprobe X-ray images using the program XMapTools (Lanari et al 2014). The (Fe3+/Fetot) value in chlorite was directly measured using μXANES spot analyses collected at the Fe-K edge. The results indicate a homogeneous temperature of 300-350° C throughout the crystallization. This result excludes the T as the main parameter to explain the Fe and Mg zoning patterns. Several other origins can be proposed and discussed in order to understand zoning patterns such as fluid chemistry, pressure, pH or redox variations of the fluid. Grant, N.T., 1992. Post-emplacement extension within a thrust sheet from the central Pyrenees. Journal of the Geological Society 149, 775-792. Lanari, P., Vidal, O., De Andrade, V., Dubacq, B., Lewin, E., Grosch, E.G., Schwartz, S., 2014. XMapTools: A MATLAB©-based program for electron microprobe X-ray image processing and geothermobarometry. Computers & Geosciences 62, 227-240.

Development and Hot-fire Testing of Additively Manufactured Copper Combustion Chambers for Liquid Rocket Engine Applications

NASA Technical Reports Server (NTRS)

Gradl, Paul R.; Greene, Sandy Elam; Protz, Christopher S.; Ellis, David L.; Lerch, Bradley A.; Locci, Ivan E.

2017-01-01

NASA and industry partners are working towards fabrication process development to reduce costs and schedules associated with manufacturing liquid rocket engine components with the goal of reducing overall mission costs. One such technique being evaluated is powder-bed fusion or selective laser melting (SLM), commonly referred to as additive manufacturing (AM). The NASA Low Cost Upper Stage Propulsion (LCUSP) program was designed to develop processes and material characterization for GRCop-84 (a NASA Glenn Research Center-developed copper, chrome, niobium alloy) commensurate with powder-bed AM, evaluate bimetallic deposition, and complete testing of a full scale combustion chamber. As part of this development, the process has been transferred to industry partners to enable a long-term supply chain of monolithic copper combustion chambers. To advance the processes further and allow for optimization with multiple materials, NASA is also investigating the feasibility of bimetallic AM chambers. In addition to the LCUSP program, NASA has completed a series of development programs and hot-fire tests to demonstrate SLM GRCop-84 and other AM techniques. NASA's efforts include a 4K lbf thrust liquid oxygen/methane (LOX/CH4) combustion chamber and subscale thrust chambers for 1.2K lbf LOX/hydrogen (H2) applications that have been designed and fabricated with SLM GRCop-84. The same technologies for these lower thrust applications are being applied to 25-35K lbf main combustion chamber (MCC) designs. This paper describes the design, development, manufacturing and testing of these numerous combustion chambers, and the associated lessons learned throughout their design and development processes.

Operation and Performance Measurement on Engines in Sea Level Test Facilities

DTIC Science & Technology

1984-03-01

progressively larger collector to ’:7.. *: capture the efflux, but secondary airflow increases rapidly as collector area .-- increases. Therefore...1 + Dbm + Dc + Dts + Dbt - WeVe + (Pe - P 2 )Ae where the terms above and to follow are defined as Fa - measured thrust from load cell Pn - net thrust...Dbt - buoyancy (boat-tail) drag on exhaust nozzle. Considering that . (Pe-P" 2 )Ae + WeVe - (Pe-P..)Ae + (P-1-P- 2 )Ae + WeVe and .. .* Pg (Pe-PŖ)Ae

The ENABLER - Based on proven NERVA technology

NASA Astrophysics Data System (ADS)

Livingston, Julie M.; Pierce, Bill L.

The ENABLER reactor for use in a nuclear thermal propulsion engine uses the technology developed in the NERVA/Rover program, updated to incorporate advances in the technology. Using composite fuel, higher power densities per fuel element, improved radiation resistant control components and the advancements in use of carbon-carbon materials; the ENABLER can provide a specific impulse of 925 seconds, an engine thrust to weight (excluding reactor shield) approaching five, an improved initial mass in low Earth orbit and a consequent reduction in launch costs and logistics problems. This paper describes the 75,000 lbs thrust ENABLER design which is a low cost, low risk approach to meeting tommorrow's space propulsion needs.

Solar electric geocentric transfer with attitude constraints: Analysis

NASA Technical Reports Server (NTRS)

Sackett, L. L.; Malchow, H. L.; Delbaum, T. N.

1975-01-01

A time optimal or nearly time optimal trajectory program was developed for solar electric geocentric transfer with or without attitude constraints and with an optional initial high thrust stage. The method of averaging reduces computation time. A nonsingular set of orbital elements is used. The constraints, which are those of one of the SERT-C designs, introduce complexities into the analysis and the solution yields possible discontinuous changes in thrust direction. The power degradation due to VanAllen radiation is modeled analytically. A wide range of solar cell characteristics is assumed. Effects such as oblateness and shadowing are included. The analysis and the results of many example runs are included.

Hydrogen-oxygen auxiliary propulsion for the space shuttle. Volume 2: Low pressure thrusters

NASA Technical Reports Server (NTRS)

1973-01-01

An abbreviated program was conducted to investigate igniter, injector, and thrust chamber technology for a 10.3 N/cm2 (15 psia) chamber pressure, 6660 N (1500 lbf) gaseous H2/O2 APS thruster for the Space Shuttle Vehicle. Successful catalytic igniter tests were conducted with ambient and cold propellants. Injector testing with a heat sink chamber (MR = 2.5, area ratio = 5.0) gave a measured specific impulse of 386 sec with 11% of the fuel used as film coolant. This coolant flow rate was demonstrated to be more than adequate to cool a spun adiabatic wall, flightweight thrust chamber.

Measurements of atmospheric turbulence effects on tail rotor acoustics

NASA Technical Reports Server (NTRS)

Hagen, Martin J.; Yamauchi, Gloria K.; Signor, David B.; Mosher, Marianne

1994-01-01

Results from an outdoor hover test of a full-scale Lynx tail rotor are presented. The investigation was designed to further the understanding of the acoustics of an isolated tail rotor hovering out-of-ground effect in atmospheric turbulence, without the effects of the main rotor wake or other helicopter components. Measurements include simultaneous rotor performance, noise, inflow, and far-field atmospheric turbulence. Results with grid-generated inflow turbulence are also presented. The effects of atmospheric turbulence ingestion on rotor noise are quantified. In contradiction to current theories, increasing rotor inflow and rotor thrust were found to increase turbulence ingestion noise. This is the final report of Task 13A--Helicopter Tail Rotor Noise, of the NASA/United Kingdom Defense Research Agency cooperative Aeronautics Research Program.

Saturn Apollo Program

NASA Image and Video Library

1960-01-01

This cutaway illustrates the S-I stage, the first stage of the Saturn I vehicle developed by the Marshall Space Flight Center (MSFC). The stage was propelled by a cluster of eight H-1 engines, capable of producing 1,500,000 pounds of thrust.

Rocket thrust chamber thermal barrier coatings

NASA Technical Reports Server (NTRS)

Quentmeyer, R. J.

1985-01-01

Subscale rocket thrust chamber tests were conducted to evaluate the effectiveness and durability of thin yttria stabilized zirconium oxide coatings applied to the thrust chamber hot-gas side wall. The fabrication consisted of arc plasma spraying the ceramic coating and bond coat onto a mandrell and then electrodepositing the copper thrust chamber wall around the coating. Chambers were fabricated with coatings .008, and .005 and .003 inches thick. The chambers were thermally cycled at a chamber pressure of 600 psia using oxygen-hydrogen as propellants and liquid hydrogen as the coolant. The thicker coatings tended to delaminate, early in the cyclic testing, down to a uniform sublayer which remained well adhered during the remaining cycles. Two chambers with .003 inch coatings were subjected to 1500 thermal cycles with no coating loss in the throat region, which represents a tenfold increase in life over identical chambers having no coatings. An analysis is presented which shows that the heat lost to the coolant due to the coating, in a rocket thrust chamber design having a coating only in the throat region, can be recovered by adding only one inch to the combustion chamber length.

Age and source of water in springs associated with the Jacksonville Thrust Fault Complex, Calhoun County, Alabama

USGS Publications Warehouse

Robinson, James L.

2004-01-01

Water from wells and springs accounts for more than 90 percent of the public water supply in Calhoun County, Alabama. Springs associated with the Jacksonville Thrust Fault Complex are used for public water supply for the cities of Anniston and Jacksonville. The largest ground-water supply is Coldwater Spring, the primary source of water for Anniston, Alabama. The average discharge of Coldwater Spring is about 32 million gallons per day, and the variability of discharge is about 75 percent. Water-quality samples were collected from 6 springs and 15 wells in Calhoun County from November 2001 to January 2003. The pH of the ground water typically was greater than 6.0, and specific conductance was less than 300 microsiemens per centimeter. The water chemistry was dominated by calcium, carbonate, and bicarbonate ions. The hydrogen and oxygen isotopic composition of the water samples indicates the occurrence of a low-temperature, water-rock weathering reaction known as silicate hydrolysis. The residence time of the ground water, or ground-water age, was estimated by using analysis of chlorofluorocarbon, sulfur hexafluoride, and regression modeling. Estimated ground-water ages ranged from less than 10 to approximately 40 years, with a median age of about 18 years. The Spearman rho test was used to identify statistically significant covariance among selected physical properties and constituents in the ground water. The alkalinity, specific conductance, and dissolved solids increased as age increased; these correlations reflect common changes in ground-water quality that occur with increasing residence time and support the accuracy of the age estimates. The concentration of sodium and chloride increased as age increased; the correlation of these constituents is interpreted to indicate natural sources for chloride and sodium. The concentration of silica increased as the concentration of potassium increased; this correlation, in addition to the isotopic data, is evidence that silicate hydrolysis of clay minerals occurred. The geochemical modeling program NETPATH was used to investigate possible mixing scenarios that could yield the chemical composition of water collected from springs associated with the Jacksonville Thrust Fault Complex. The results of NETPATH modeling suggest that the primary source of water in Coldwater Spring is a deep aquifer, and only small amounts of rainwater from nearby sources are discharged from the spring. Starting with Piedmont Sports Spring and moving southwest along a conceptual ground-water flow path that parallels the Jacksonville Thrust Fault Complex, NETPATH simulated the observed water quality of each spring, in succession, by mixing rainwater and water from the spring just to the northeast of the spring being modeled. The percentage of rainwater and ground water needed to simulate the quality of water flowing from the springs ranged from 1 to 25 percent rainwater and 75 to 99 percent ground water.

Progress on Variable Cycle Engines

NASA Technical Reports Server (NTRS)

Westmoreland, J. S.; Howlett, R. A.; Lohmann, R. P.

1979-01-01

Progress in the development and future requirements of the Variable Stream Control Engine (VSCE) are presented. The two most critical components of this advanced system for future supersonic transports, the high performance duct burner for thrust augmentation, and the low jet coannular nozzle were studied. Nozzle model tests substantiated the jet noise benefit associated with the unique velocity profile possible with a coannular nozzle system on a VSCE. Additional nozzle model performance tests have established high thrust efficiency levels only at takeoff and supersonic cruise for this nozzle system. An experimental program involving both isolated component and complete engine tests has been conducted for the high performance, low emissions duct burner with good results and large scale testing of these two components is being conducted using a F100 engine as the testbed for simulating the VSCE. Future work includes application of computer programs for supersonic flow fields to coannular nozzle geometries, further experimental testing with the duct burner segment rig, and the use of the Variable Cycle Engine (VCE) Testbed Program for evaluating the VSCE duct burner and coannular nozzle technologies.

Analysis of the flow field generated near an aircraft engine operating in reverse thrust. M.S. Thesis

NASA Technical Reports Server (NTRS)

Ledwith, W. A., Jr.

1972-01-01

A computer solution is developed to the exhaust gas reingestion problem for aircraft operating in the reverse thrust mode on a crosswind-free runway. The computer program determines the location of the inlet flow pattern, whether the exhaust efflux lies within the inlet flow pattern or not, and if so, the approximate time before the reversed flow reaches the engine inlet. The program is written so that the user is free to select discrete runway speeds or to study the entire aircraft deceleration process for both the far field and cross-ingestion problems. While developed with STOL applications in mind, the solution is equally applicable to conventional designs. The inlet and reversed jet flow fields involved in the problem are assumed to be noninteracting. The nacelle model used in determining the inlet flow field is generated using an iterative solution to the Neuman problem from potential flow theory while the reversed jet flow field is adapted using an empirical correlation from the literature. Sample results obtained using the program are included.

An air bearing system for small high speed gas turbines

NASA Astrophysics Data System (ADS)

Turner, A. B.; Davies, S. J.; Nimir, Y. L.

1994-03-01

This paper describes the second phase of an experimental program concerning the application of air bearings to small turbomachinery test rigs and small gas turbines. The first phase examined externally pressurized (EP) journal bearings, with a novel EP thrust bearing, for application to 'warm air' test rigs, and was entirely successful at rotational speeds in excess of 100,000 rpm. This second phase examined several designs of tilting pad-spiring journal bearings, one with a novel form of externally pressurized pad, but all using the original EP thrust bearing. The designs tested are described, including some oscillogram traces, for tests up to a maximum of 70,000 rpm; the most successful using a carbon pad-titanium beam spring arrangement. The thrust bearing which gave trouble-free operation throughout, is also described. The results of an original experiment to measure the 'runway speed' of a radial inflow turbine are also presented, which show that overspeeds of 58 percent above the design speed can result from free-power turbine coupling failure.

Metrics for NASA Aeronautics Research Mission Directorate (ARMD) Strategic Thrust 3B Vertical Lift Strategic Direction

NASA Technical Reports Server (NTRS)

Hochstetler, Ronald D.; Salvano, Dan; Gorton, Susan A.

2017-01-01

The NASA Aeronautics Research Mission Directorate (ARMD) Strategic Implementation Plan details an ambitious plan for aeronautical research for the next quarter century and beyond. It includes a number of advanced technologies needed to address requirements of the overall aviation community (domestic and international), with an emphasis on safety, efficiency, operational flexibility, and alternative propulsion air transport options. The six ARMD Strategic Thrust Areas (STAs) represent a specific set of multi-decade research agendas for creating the global aviation improvements most in demand by the aviation service consumers and the general public. To provide NASA with a measurement of the preeminent value of these research areas, it was necessary to identify and quantify the measurable benefits to the aviation community from capabilities delivered by the research programs. This paper will describe the processes used and the conclusions reached in defining the principal metrics for ARMD Strategic Thrust Area 3B "Vertical Lift Strategic Direction."

Saturn Apollo Program

NASA Image and Video Library

1967-01-01

This is a cutaway illustration of the Saturn V launch vehicle with callouts of the major components. The Saturn V is the largest and most powerful launch vehicle developed in the United States. It was a three stage rocket, 363 feet in height, used for sending American astronauts to the moon and for placing the Skylab in Earth orbit. The Saturn V was designed to perform Earth orbital missions through the use of the first two stages, while all three stages were used for lunar expeditions. The S-IC stage (first stage) was powered by five F- engines, which burned kerosene and liquid oxygen to produce more than 7,500,000 pounds of thrust. The S-II (second) stage was powered by five J-2 engines, that burned liquid hydrogen and liquid oxygen and produced 1,150,000 pounds thrust. The S-IVB (third) stage used one J-2 engine, producing 230,000 pounds of thrust, with a re-start capability. The Marshall Space Flight Center and its contractors designed, developed, and assembled the Saturn V launch vehicle stages.

Criteria for design of integrated flight/propulsion control systems for STOVL fighter aircraft

NASA Technical Reports Server (NTRS)

Franklin, James A.

1993-01-01

As part of NASA's program to develop technology for short takeoff and vertical landing (STOVL) fighter aircraft, control system designs have been developed for a conceptual STOVL aircraft. This aircraft is representative of the class of mixed-flow remote-lift concepts that was identified as the preferred design approach by the U.S./U.K. STOVL Joint Assessment and Ranking Team. The control system designs have been evaluated throughout the powered-lift flight envelope on the Vertical Motion Simulator (VMS) at Ames Research Center. Items assessed in the control system evaluation were: maximum control power used in transition and vertical flight, control system dynamic response associated with thrust transfer for attitude control, thrust margin in the presence of ground effect and hot-gas ingestion, and dynamic thrust response for the engine core. Effects of wind, turbulence, and ship airwake disturbances are incorporated in the evaluation. Results provide the basis for a reassessment of existing flying-qualities design criteria applied to STOVL aircraft.

Design criteria for integrated flight/propulsion control systems for STOVL fighter aircraft

NASA Technical Reports Server (NTRS)

Franklin, James A.

1993-01-01

As part of NASA's program to develop technology for short takeoff and vertical landing (STOVL) fighter aircraft, control system designs have been developed for a conceptual STOVL aircraft. This aircraft is representative of the class of mixed-flow remote-lift concepts that was identified as the preferred design approach by the US/UK STOVL Joint Assessment and Ranking Team. The control system designs have been evaluated throughout the powered-lift flight envelope on Ames Research Center's Vertical Motion Simulator. Items assessed in the control system evaluation were: maximum control power used in transition and vertical flight, control system dynamic response associated with thrust transfer for attitude control, thrust margin in the presence of ground effect and hot gas ingestion, and dynamic thrust response for the engine core. Effects of wind, turbulence, and ship airwake disturbances are incorporated in the evaluation. Results provide the basis for a reassessment of existing flying qualities design criteria applied to STOVL aircraft.

Force-motion phase relations and aerodynamic performance of a plunging plate

NASA Astrophysics Data System (ADS)

Son, Onur; Cetiner, Oksan

2018-02-01

Due to the unsteady motion of a plunging plate, forces acting on the body experience a phase difference with respect to the motion. These phase relations are investigated experimentally for a harmonically plunging plate within an amplitude range of 0.05≤ {a/c}≤ 0.6, reduced frequency range of 0.78<{k}<7.06, and at a constant Reynolds number of 10,000. Both streamwise and cross-stream force components are found to have a phase lag following the motion; however, their variations are different. The phase lag of the force on the cross-stream direction increases as the amplitude increases. Drag-thrust transition has an influence on the streamwise force phase lags, which starts to increase when the thrust starts to be produced. Particle image velocimetry measurements are also performed to reveal the relations between vortex structures and force measurements. Leading edge vortex shedding characteristics are observed to be changing from drag occurring cases to thrust producing cases in parallel with the increment in phase lags.

Conical Magnetic Bearings Developed for Active Stall Control in Gas Turbine Engines

NASA Technical Reports Server (NTRS)

Trudell, Jeffrey J.; Kascak, Albert F.; Provenza, Andrew J.; Buccieri, Carl J.

2004-01-01

Active stall control is a current research area at the NASA Glenn Research Center that offers a great benefit in specific fuel consumption by allowing the gas turbine to operate beyond the onset of stall. Magnetic bearings are being investigated as a new method to perform active stall control. This enabling global aviation safety technology would result in improved fuel efficiency and decreased carbon dioxide emissions, as well as improve safety and reliability by eliminating oil-related delays and failures of engine components, which account for 40 percent of the commercial aircraft departure delays. Active stall control works by perturbing the flow in front of the compressor stage such that it cancels the pressure wave, which causes the compressor to go into stall. Radial magnetic bearings are able to whirl the shaft so that variations in blade tip leakage would flow upstream causing a perturbation wave that could cancel the rotating stall cell. Axial or thrust magnetic bearings cannot be used to cancel the surge mode in the compressor because they have a very low bandwidth and thus cannot modulate at a high enough frequency. Frequency response is limited because the thrust runner cannot be laminated. To improve the bandwidth of magnetic thrust bearings, researchers must use laminations to suppress the eddy currents. A conical magnetic bearing can be laminated, resulting in increased bandwidth in the axial direction. In addition, this design can produce both radial and thrust force in a single bearing, simplifying the installation. The proposed solution combines the radial and thrust bearing into one design that can be laminated--a conical magnetic bearing. The new conical magnetic bearing test rig, funded by a Glenn fiscal year 2002 Director's Discretionary Fund, was needed because none of the existing rigs has an axial degree of freedom. The rotor bearing configuration will simulate that of the main shaft on a gas turbine engine. One conical magnetic bearing replaces the ball bearing in front of the compressor, and the second replaces the roller bearing behind the burner. The rig was made operational to 10,000 rpm under Smart Efficient Components funding, and both position and current adaptive vibration control have been demonstrated. Upon program completion, recommendations will be made as to the efficacy of the conical magnetic bearing for active stall control.

A reevaluation of the age of the Vincent-Chocolate Mountains thrust system, southern California

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

Jacobsen, C.E.; Barth, A.P.

1993-04-01

The Vincent-Chocolate Mountains (VCM) thrust superposes Mesozoic arc plutons and associated Precambrian country rock above subduction-related Pelona-Orocopia schist. The thrust is disrupted in many areas by postmetamorphic deformation, but appears to be intact in the San Gabriel Mountains. Two Rb-Sr mineral-isochron ages from Pelona Schist and mylonite in the San Gabriel Mountains led Ehlig (1981) to conclude that the original thrusting event occurred at c. 60 Ma. However, biotite K-Ar ages determined by Miller and Morton (1980) for upper plate in the same area caused Dillon (1986) to reach a different conclusion. The biotite ages range mainly from 74--60 Mamore » and increase structurally upward from the VCM thrust. Dillon (1986) inferred that the age gradient was due to uplift and cooling of the upper plate during underthrusting of Pelona Schist. This would indicate that the VCM thrust was at least 74 Ma in age. An alternative to the interpretation of Dillon (1986) is that the biotite age gradient largely predates the VCM thrust. Upward heat flow, leading to older ages at higher structural levels, could have resulted from either static cooling of Cretaceous plutons or uplift and erosion induced by crustal thickening during possible west-directed intra-arc thrusting at c. 88--78 Ma (May and Walker, 1989). Subsequent underthrusting of Pelona Schist would establish a cold lower boundary to the crust and cause the closure of isotopic systems in the base of the upper plate. A 60 Ma time of thrusting is also suggested by two amphibole [sup 40]Ar/[sup 39]Ar ages from the Pelona Schist of the San Gabriel Mountains. Peak metamorphic temperature in this area was below 480 C and amphibole ages should thus indicate time of crystallization rather than subsequent cooling. Four phengite [sup 40]Ar/[sup 39]Ar ages of 55--61 Ma from Pelona Schist and mylonite indicate rapid cooling from peak metamorphic temperatures, consistent with subduction refrigeration.« less

Kinematic evolution of a regional-scale gravity-driven deepwater fold-and-thrust belt: The Lamu Basin case-history (East Africa)

NASA Astrophysics Data System (ADS)

Cruciani, F.; Barchi, M. R.; Koyi, H. A.; Porreca, M.

2017-08-01

The deepwater fold-and-thrust belts (DWFTBs) are geological structures recently explored thanks to advances in offshore seismic imaging by oil industry. In this study we present a kinematic analysis based on three balanced cross-sections of depth-converted, 2-D seismic profiles along the offshore Lamu Basin (East African passive margin). This margin is characterized by a regional-scale DWFTB (> 450 km long), which is the product of gravity-driven contraction on the shelf that exhibits complex structural styles and differing amount of shortening along strike. Net shortening is up to 48 km in the northern wider part of the fold-and-thrust belt (≈ 180 km), diminishing to < 15 km toward the south, where the belt is markedly narrower (≈ 50 km). The three balanced profiles show a shortening percentage around 20% (comparable with the maximum values documented in other gravity-driven DWFTBs), with a significant variability along dip: higher values are achieved in the outer (i.e. down-dip) portion of the system, dominated by basinward-verging, imbricate thrust sheets. Fold wavelength increases landward, where doubly-verging structures and symmetric detachment folds accommodate a lower amount of shortening. Similar to other cases, a linear and systematic relationship between sedimentary thickness and fold wavelength is observed. Reconstruction of the rate of shortening through time within a fold-and-thrust belt shows that after an early phase of slow activation (Late Cretaceous), > 95% of net shortening was produced in < 10 Myr (during Paleocene). During this acme phase, which followed a period of high sedimentation rate, thrusts were largely synchronous and the shortening rate reached a maximum value of 5 mm/yr. The kinematic evolution reconstructed in this study suggests that the structural evolution of gravity-driven fold-and-thrust belts differs from the accretionary wedges and the collisional fold-and-thrust belts, where thrusts propagate in-sequence and shortening is uniformly accommodated along dip.