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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 685 records · Page 38
All sourcesNASA NTRSDOE OSTI
NASA NTRS1994

Experience with Ada on the F-18 High Alpha Research Vehicle flight test program

Considerable experience has been acquired with Ada at the NASA Dryden Flight Research Facility during the on-going High Alpha Technology Program. In this program, an F-18 aircraft has been highly modified by the addition of thrust-vectoring vanes to the airframe. In addition, substantial alteration was made in the original quadruplex flight control system. The result is the High Alpha Research Vehicle. An additional research flight control computer was incorporated in each of the four channels. Software for the research flight control computer was written Ada. To date, six releases of this software have been flown. This paper provides a detailed description of the modifications to the research flight control system. Efficient ground-testing of the software was accomplished by using simulations that used the Ada for portions of their software. These simulations are also described. Modifying and transferring the Ada flight software to the software simulation configuration has allowed evaluation of this language. This paper also discusses such significant issues in using Ada as portability, modifiability, and testability as well as documentation requirements.

Regenie, Victoria A.↗
NASA NTRS1996

Computational and Experimental Efforts in Gravity Probe B Microthruster Analysis

Gravity Probe B, an experiment to test the theory of relativity, will be launched near the turn of the millennium. Due to the precise pointing requirements need to successfully carry out this experiment, the satellite will use sixteen proportionally controlled microthrusters as a main component of the attitude control system. These microthrusters use the helium boil-off from the on-board dewar as propellant. Marshall Space Flight Center, verified the design of the thruster flow path by both computational and experimental methods. The flow performance of the thruster has been adequately characterized. Graphs show specific impulse, thrust coefficient, discharge coefficient, and mass flow rate trends. Value was added to the program through gained confidence in the design of the thruster and through valuation of some design trade-offs.

Droege, Alan↗
NASA NTRS2010

Acoustic Measurements of Small Solid Rocket Motor

Rocket acoustic noise can induce loads and vibration on the vehicle as well as the surrounding structures. Models have been developed to predict these acoustic loads based on scaling existing solid rocket motor data. The NASA Marshall Space Flight Center acoustics team has measured several small solid rocket motors (thrust below 150,000 lbf) to anchor prediction models. This data will provide NASA the capability to predict the acoustic environments and consequent vibro-acoustic response of larger rockets (thrust above 1,000,000 lbf) such as those planned for the NASA Constellation program. This paper presents the methods used to measure acoustic data during the static firing of small solid rocket motors and the trends found in the data.

Vargas, Magda B.↗
NASA NTRS2005

SOLID SOLUTION CARBIDES ARE THE KEY FUELS FOR FUTURE NUCLEAR THERMAL PROPULSION

Nuclear thermal propulsion uses nuclear energy to directly heat a propellant (such as liquid hydrogen) to generate thrust for space transportation. In the 1960 s, the early Rover/Nuclear Engine for Rocket Propulsion Application (NERVA) program showed very encouraging test results for space nuclear propulsion but, in recent years, fuel research has been dismal. With NASA s renewed interest in long-term space exploration, fuel researchers are now revisiting the RoverMERVA findings, which indicated several problems with such fuels (such as erosion, chemical reaction of the fuel with propellant, fuel cracking, and cladding issues) that must be addressed. It is also well known that the higher the temperature reached by a propellant, the larger the thrust generated from the same weight of propellant. Better use of fuel and propellant requires development of fuels capable of reaching very high temperatures. Carbides have the highest melting points of any known material. Efforts are underway to develop carbide mixtures and solid solutions that contain uranium carbide, in order to achieve very high fuel temperatures. Binary solid solution carbides (U, Zr)C have proven to be very effective in this regard. Ternary carbides such as (U, Zr, X) carbides (where X represents Nb, Ta, W, and Hf) also hold great promise as fuel material, since the carbide mixtures in solid solution generate a very hard and tough compact material. This paper highlights past experience with early fuel materials and bi-carbides, technical problems associated with consolidation of the ingredients, and current techniques being developed to consolidate ternary carbides as fuel materials.

Panda, Binayak↗
NASA NTRS1975

Effect of attitude constraints on solar-electric geocentric transfers

The present work assesses the increase in flight time and fuel consumption due to introducing attitude constraints on both the thrust vector and the plane of the solar cell arrays on geocentrically orbiting spacecraft. A modified version of the SECKSPOT computer program calculates nearly time-optimal trajectories for the constrained case of zero pitch and roll. Unconstrained cases are generated with the SECKSPOT code. It is concluded that with a pitch constraint but without a roll constraint, power would not be a function of thrust direction, and so the time-optimal thruster direction would be along the projection of the primer vector in the plane normal to the radius vector. The roll constraint would cause power to become a function of thrust angle and sun angle. For certain sun angles the locus of the ratio of power to maximum power is concave and thus there may be jumps in the control angle. Comparisons are made for a SERT-C type mission between constrained and unconstrained cases in an inverse square gravity field.

Sackett, L. L.↗
NASA NTRS2000

Flying Qualities Evaluation of a Commuter Aircraft With an Ice Contaminated Tailplane

During the NASA/FAA (Federal Aviation Administration) Tailplane Icing Program, pilot evaluations of aircraft flying qualities were conducted with various ice shapes attached to the horizontal tailplane of the NASA Twin Otter Icing Research Aircraft. Initially, only NASA pilots conducted these evaluations, assessing the differences in longitudinal flight characteristics between the baseline or clean aircraft, and the aircraft configured with an Ice Contaminated Tailplane (ICT). Longitudinal tests included Constant Airspeed Flap Transitions, Constant Airspeed Thrust Transitions, zero-G Pushovers, Repeat Elevator Doublets, and Simulated Approach and Go-Around tasks. Later in the program, guest pilots from government and industry were invited to fly the NASA Twin Otter configured with a single full-span artificial ice shape attached to the leading edge of the horizontal tailplane. This shape represented ice formed due to a 'Failed Boot' condition, and was generated from tests in the Glenn Icing Research Tunnel on a full-scale tailplane model. Guest pilots performed longitudinal handling tests, similar to those conducted by the NASA pilots, to evaluate the ICT condition. In general, all pilots agreed that longitudinal flying qualities were degraded as flaps were lowered, and further degraded at high thrust settings. Repeat elevator doublets demonstrated reduced pitch damping effects due to ICT, which is a characteristic that results in degraded flying qualities. Pilots identified elevator control force reversals (CFR) in zero-G pushovers at a 20 deg flap setting, a characteristic that fails the FAR 25 no CFR certification requirement. However, when the same pilots used the Cooper-Harper rating scale to perform a simulated approach and go-around task at the 20 deg flap setting, they rated the airplane as having Level I and Level II flying qualities respectively. By comparison, the same task conducted at the 30 deg flap setting, resulted in Level II flying qualities for the approach portion, and Level III for the go-around portion.The results of this program indicate that safe and acceptable flying qualities with an ICT condition, can be effectively assessed by task-oriented pilot maneuvers. In addition, other maneuvers such as repeat elevator doublets provide good qualitative and quantitative assessments of pitch damping and elevator effectiveness, which are characteristics that correlate well with pilot task ratings. The results of this testing indicate that the FAR 25 zero-G pushover maneuver, which requires no CFR during its execution, may be an overly conservative pass/fail criteria for aircraft certification.

Ranaudo, Richard J.↗
NASA NTRS1972

A method for calculating transient thrust and flow-rate levels for Mariner type attitude control nitrogen gas jets

The purpose of this report is to define and program the transient pneumatic flow equations necessary to determine, for a given set of conditions (geometry, pressures, temperatures, valve on time, etc.), the total nitrogen impulse and mass flow per pulse for the single pulsing of a Mariner type reaction control assembly valve. The rates of opening and closing of the valves are modeled, and electrical pulse durations from 20 to 100 ms are investigated. In developing the transient flow analysis, maximum use was made of the steady-state analysis. The impulse results are also compared to an equivalent square-wave impulse for both the Mariner Mars 1971 (MM'71) and Mariner Mars 1964 (MM'64) systems. It is demonstrated that, whereas in the MM'64 system, the actual impulse was as much as 56 percent higher than an assumed impulse (which is the product of the steady-state thrust and value on time i.e., the square wave), in the MM'71 system, these two values were in error in the same direction by only approximately 4 percent because of the larger nozzle areas and shorter valve stroke used.

Ferrera, J. D.↗
NASA NTRS2005

Geometry of Thrust Faults Beneath Amenthes Rupes, Mars

Amenthes Rupes is a 380 km-long lobate fault scarp located in the eastern hemisphere of Mars near the dichotomy boundary. The scarp is marked by about 1 km of vertical separation across a northeast dipping thrust fault (top to the SW) and offsets heavily-cratered terrain of Late Noachian age, the visible portion of which was in place by 3.92 Ga and the buried portion in place between 4.08 and 4.27 Ga. The timing of scarp formation is difficult to closely constrain. Previous geologic mapping shows that near the northern end of Amenthes Rupes, Hesperian age basalts terminate at the scarp, suggesting that fault slip predated the emplacement of these flows at 3.69 to 3.9 Ga. Maxwell and McGill also suggest the faulting ceased before the final emplacement of the Late Hesperian lavas on Isidis Planitia. The trend of the faults at Amenthes, like many thrust faults at the dichotomy boundary, parallels the boundary itself. Schultz and Watters used a dislocation modeling program to match surface topography and vertical offset of the scarp at Amenthes Rupes, varying the dip and depth of faulting, assuming a slip of 1.5 km on the fault. They modeled faulting below Amenthes Rupes as having a dip of between 25 and 30 degrees and a depth of 25 to 35 km, based on the best match to topography. Assuming a 25 degree dip and surface measurements of vertical offset of between 0.3 and 1.2 km, Watters later estimated the maximum displacement on the Amenthes Rupes fault to be 2.90 km. However, these studies did not determine the geometry of the thrust using quantitative constraints that included shortening estimates. Amenthes Rupes deforms large preexisting impact craters. We use these craters to constrain shortening across the scarp and combine this with vertical separation to infer fault geometry. Fault dip was also estimated using measurements of scarp morphology. Measurements were based on 460 m (1/128 per pixel) digital elevation data from the Mars Orbiter Laser Altimeter (MOLA), an instrument on the Mars Global Surveyor (MGS) satellite.

Vidal, A.↗
NASA NTRS2011

Solid Rocket Booster (SRB) - Evolution and Lessons Learned During the Shuttle Program

The Solid Rocket Booster (SRB) element integrates all the subsystems needed for ascent flight, entry, and recovery of the combined Booster and Motor system. These include the structures, avionics, thrust vector control, pyrotechnic, range safety, deceleration, thermal protection, and retrieval systems. This represents the only human-rated, recoverable and refurbishable solid rocket ever developed and flown. Challenges included subsystem integration, thermal environments and severe loads (including water impact), sometimes resulting in hardware attrition. Several of the subsystems evolved during the program through design changes. These included the thermal protection system, range safety system, parachute/recovery system, and others. Obsolescence issues occasionally required component recertification. Because the system was recovered, the SRB was ideal for data and imagery acquisition, which proved essential for understanding loads and system response. The three main parachutes that lower the SRBs to the ocean are the largest parachutes ever designed, and the SRBs are the largest structures ever to be lowered by parachutes. SRB recovery from the ocean was a unique process and represented a significant operational challenge; requiring personnel, facilities, transportation, and ground support equipment. The SRB element achieved reliability via extensive system testing and checkout, redundancy management, and a thorough postflight assessment process. Assembly and integration of the booster subsystems was a unique process and acceptance testing of reused hardware components was required for each build. Extensive testing was done to assure hardware functionality at each level of stage integration. Because the booster element is recoverable, subsystems were available for inspection and testing postflight, unique to the Shuttle launch vehicle. Problems were noted and corrective actions were implemented as needed. The postflight assessment process was quite detailed and a significant portion of flight operations. The SRBs provided fully redundant critical systems including thrust vector control, mission critical pyrotechnics, avionics, and parachute recovery system. The design intent was to lift off with full redundancy. On occasion, the redundancy management scheme was needed during flight operations. This paper describes some of the design challenges, how the design evolved with time, and key areas where hardware reusability contributed to improved system level understanding.

Kanner, Howard S.↗
NASA NTRS1991

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

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.

Berton, Jeffrey J.↗
NASA NTRS1986

Experiments on high speed ejectors

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.

Wu, J. J.↗
NASA NTRS1990

Developments in REDES: The rocket engine design expert system

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.

Davidian, Kenneth O.↗
NASA NTRS1990

Developments in REDES: The Rocket Engine Design Expert System

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.

Davidian, Kenneth O.↗
NASA NTRS1974

Operations analysis (study 2.1). Program SEPSIM (solar electric propulsion stage simulation)

Program SEPSIM is a FORTRAN program which performs deployment, servicing, and retrieval missions to synchronous equatorial orbit using a space tug with a continuous low thrust upper stage known as a solar electric propulsion stage (SEPS). The SEPS ferries payloads back and forth between an intermediate orbit and synchronous orbit, and performs the necessary servicing maneuvers in synchronous orbit. The tug carries payloads between the orbiter and the intermediate orbit, deploys fully fueled SEPS vehicles, and retrieves exhausted SEPS vehicles when, and if, required. The program is presently contained in subroutine form in the Logistical On-orbit VEhicle Servicing (LOVES) Program, but can also be run independently with the addition of a simple driver program.

Lang, T. J.↗
NASA NTRS1978

Multibody system applications and simulations at the Jet Propulsion Laboratory

The historical development of the Jet Propulsion Laboratory (JPL) of generic computer programs for solving the H-M-H equations of motion of point-connected sets of rigid bodies in a topological tree is traced, as well as the application of these programs and the multibody modelling approach to the design of spacecraft control systems. These include thrust vector control and science instrument articulation on such vehicles as Mariner 9, Mariner 10, Viking Orbiter, and Voyager.

Fleischer, G. E.↗
NASA NTRS1979

Progress on Variable Cycle Engines

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.

Westmoreland, J. S.↗
NASA NTRS1984

NASA's life sciences and space radiation biology

Plans for the various missions in which men and women are expected to participate during the next 10 years are outlined. Such missions include flights of up to three months duration in low earth orbit as well as possible short excursions to geosynchronous orbit. Research activities are described which cover the full spectrum of physiological and psychological responses to space flight. These activities are shown to contribute to the ongoing Shuttle program and the future Space Station. The paper includes a summary of the major technical thrusts needed to support extended habitation in space.

Rambaut, P.↗