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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 145 records · Page 8
All sourcesNASA NTRSDOE OSTI
NASA NTRS1974

Optimal high- and low-thrust geocentric transfer

A computer code which rapidly calculates time optimal combined high- and low-thrust transfers between two geocentric orbits in the presence of a strong gravitational field has been developed as a mission analysis tool. The low-thrust portion of the transfer can be between any two arbitrary ellipses. There is an option for including the effect of two initial high-thrust impulses which would raise the spacecraft from a low, initially circular orbit to the initial orbit for the low-thrust portion of the transfer. In addition, the effect of a single final impulse after the low-thrust portion of the transfer may be included. The total Delta V for the initial two impulses must be specified as well as the Delta V for the final impulse. Either solar electric or nuclear electric propulsion can be assumed for the low-thrust phase of the transfer.

Sackett, L. L.↗
NASA NTRS1974

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

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.

Hambly, D.↗
NASA NTRS1976

Aerospike thrust chamber program

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.

Campbell, J., Jr.↗
NASA NTRS1977

Experimental investigations of thrust vectoring systems for VTOL aircraft

This paper presents a summary of two technology programs sponsored by NASA to investigate the characteristics of two thrust vectoring schemes for V/STOL aircraft. The operational capability of the VTOL aircraft is dependent on maximum utilization of the installed thrust in both the cruise and powered lift modes of flight. An effective thrust vectoring system on the cruise propulsion unit is therefore essential to provide maximum payload in hover and STOL plus minimum specific fuel consumption in loiter and cruise. Introducing a high by-pass ratio fan system, augmenting the gas generator thrust, as the propulsion system for VTOL aircraft places increased significance on the performance of the relatively short coupled thrust vectoring systems. The two programs discussed herein include both large-scale and small-scale tests of a vectoring hood system with a vented, internal-lip and swivel nozzle systems. These tests indicated that a satisfactory thrust vectoring system can be developed.

Rolls, L. S.↗
NASA NTRS1979

Effect of forward velocity and crosswind on the reverse-thrust performance of a variable-pitch fan engine

A full-size variable-pitch fan engine was tested in the Ames 40 by 80 foot wind tunnel to determine the effect of forward velocity and crosswind on reverse-thrust performance. Two flight-type inlet configurations were tested, and a flared fan nozzle was installed as an inlet for reverse-thrust operation. Steady-state reverse-thrust performance was obtained up to 54 m/s (105 knots). An abrupt decrease in reverse thrust occurred at about 30 m/s (60 knots). Reverse thrust was established following forward-to-reverse thrust transients both statically and with forward velocities only up to 30 m/s.

Reemsnyder, D. C.↗
NASA NTRS1979

Effect of forward velocity and crosswind on the reverse-thrust performance of a variable-pitch fan engine

Variable-pitch-fan engines may be attractive for future short-haul aircraft if sufficient reverse thrust is available for aircraft deceleration after touchdown. Thrust reversal is obtained in these engines by changing fan blade pitch about 90 deg, which causes the fan airflow to enter the fan duct nozzle and exhaust through the fan inlet. This capability would eliminate the heavy and costly thrust reverser system required for current fixed-pitch turbofan engines. NASA has, therefore, supported the development of advanced technology for a quiet, clean, high-bypass-ratio turbofan engine for future short-haul aircraft. In connection with this program, tests were conducted to determine the effect of forward velocity and angle of attack on steady-state reverse-thrust performance. Other objectives of the tests were related to the determination of the effect of forward velocity on forward-to-reverse thrust transient performance and the determination of the effectiveness of an overshoot blade angle technique to establish reverse thrust during a transient. The results of the tests are discussed.

Reemsnyder, D. C.↗
NASA NTRS1980

Reverse thrust performance of the QCSEE variable pitch turbofan engine

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

Samanich, N. E.↗
NASA NTRS1980

An analysis of thrust-induced effects on the longitudinal aerodynamics of STOL fighter configurations

An analysis of the thrust-induced longitudinal aerodynamic characteristics of three fighter-type configurations is presented. A brief discussion of the take-off and landing requirements for the next generation fighter aircraft leads to the conclusion that advances in lift coefficient and thrust reversing will be required to allow short-field operation. Typical power-on longitudinal aerodynamic data for the three fighter configurations, indicating different approaches to meeting the high-lift coefficient requirement, are discussed. Thrust reversing is not addressed in this paper. The power-on data are analyzed to determine what power effects are present, that is, direct thrust, boundary-layer control, vortex flows, or supercirculation. The results of the analysis indicate, for the configurations studied, that induced effects are small compared to direct thrust and that boundary-layer control, rather than leading-edge vortex flows or supercirculation, is the only significant thrust-induced effect.

Paulson, J. W., Jr.↗
NASA NTRS1980

Low thrust vehicle concept study

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.

Source record↗
NASA NTRS1980

Reverse thrust performance of the QCSEE variable pitch turbofan engine

Results of steady-state reverse and forward-to-reverse thrust transient performance tests are presented. The original QCSEE 4-segment variable fan nozzle was retested in reverse and compared with a continuous, 30-deg half-angle conical exlet. Data indicated that the significantly more stable, higher pressure recovery flow with the fixed 30-deg exlet resulted in lower engine vibrations, lower fan blade stress and approximately a 20% improvement in reverse thrust. Objective reverse thrust of 35% of takeoff thrust was reached. Thrust response of less than 1.5 sec was achieved for the approach and the takeoff-to-reverse thrust transients.

Samanich, N. E.↗
NASA NTRS1984

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

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

Rowe, R. K.↗
NASA NTRS1986

Polyphase thrust tectonic in the Barberton greenstone belt

In the circa 3.5 by-old Barberton greenstone belt, the supracrustal rocks form a thick and strongly deformed thrust complex. Structural studies in the southern part of the belt have shown that 2 separate phases of over-thrusting (D sub 1 and D sub 2) successively dismembered the original stratigraphy. Thrust nappes were subsequently refolded during later deformations (D sub 3 and D sub 4). This report deals with the second thrusting event which, in the study region appears to be dominant, and (unlike the earlier thrusting), affects the entire supracrustal pile. The supracrustal rocks form a predominantly NE/SW oriented, SE dipping tectonic fan (the D sub 2 fan) in which tectonic slices of ophiolitic-like rocks are interleaved with younger sedimentary sequences of the Diepgezet and malalotcha groups. Structural and sedimentological data indicate that the D sub 2 tectonic fan was formed during a prolonged, multi-stage regional horizontal shortening event during which several types of internal deformation mechanisms were successively and/or simultaneously active. Movement appears to have been predominantly to the NW and to the N. During D sub 2, periods of quiescence and sedimentation followed periods of thrust propagation. Although the exact kinematics which led to the formation of this fan is not yet known, paleoenvironmental interpretations together with structural data suggest that D sub 2 was probably related to (an) Archean collision(s).

Paris, I. A.↗
NASA NTRS1987

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)

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.

Smith, Tamara A.↗
NASA NTRS1984

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

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

Johns, Albert L.↗
NASA NTRS1989

Thrust vectoring effect on time-optimal 90 degrees angle of attack pitch up maneuvers of a high alpha fighter aircraft

Using thrust vectoring, the problem of pitching up a high alpha fighter aircraft to 90 degrees angle of attack in minimum time is considered. Pontryagin's maximum principle is used together with a two-point boundary value numerical algorithm to derive open-loop controls for various parameterized limits on thrust vectoring angles. Without thrust vectoring, a high alpha fighter with initial conditions 0.6 Mach number and 15,000 feet requires 7.0 seconds to pitch-up to 90 degrees. However, with 20-degree angle thrust vectoring, the pitch-up takes only 1.5 seconds with a final Mach number of 0.47. Also, the gain in altitude is less than 100 feet. More important, additional thrust vectoring yields litle benefit beyond the 20-degree angles. This paper describes control solutions in detail for thrust vectoring angles between 0 and 90 degrees.

Stalford, Harold↗
NASA NTRS1990

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

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

Taylor, John G.↗
NASA NTRS1990

Test plan pressure fed thrust chamber technology

Aerojet is developing the technology for the design of a reliable, low cost, efficient, and lightweight LOX/RP-1 pressure fed engine. This technology program is a direct result of Aerojet's liquid rocket booster (LRB) study and previous NASA studies that identified liquid engines using high bulk density hydrocarbon fuels as very attractive for a space transportation system (STS). Previous large thrust LOX/RP-1 engine development programs were characterized by costly development problems due to combustion instability damage. The combustion stability solution was typically obtained through trial and error methods of minimizing instability damage by degrading engine performance. The approach to this program was to utilize existing and newly developed combustion analysis models and design methodology to create a thrust chamber design with features having the potential of producing reliable and efficient operation. This process resulted in an engine design with a unique high thrust-per-element OFO triplet injector utilizing a low cost modular approach. Cost efficient ablative materials are baselined for the injector face and chamber. Technology demonstration will be accomplished through a hot fire test program using appropriately sized subscale hardware. This subscale testing will provide a data base to supplement the current industry data bank and to anchor and validate the applied analysis models and design methodology. Once anchored and validated, these analysis models and design methodology can be applied with greatly increased confidence to design and characterize a large scale pressure fed LOX/RP-1 thrust chamber. The objective of this test program is to generate a data base that can be used to anchor and validate existing analysis models and design methodologies and to provide early concept demonstration of a low cost, efficient LOX/RP-1 thrust chamber. Test conditions and hardware instrumentation were defined to provide data sufficient to characterize combustion stability, performance, and thermal operation over a wide thrust chamber throttling range.

Dunn, Glenn↗
NASA NTRS1991

Design of thrust vectoring exhaust nozzles for real-time applications using neural networks

Thrust vectoring continues to be an important issue in military aircraft system designs. A recently developed concept of vectoring aircraft thrust makes use of flexible exhaust nozzles. Subtle modifications in the nozzle wall contours produce a non-uniform flow field containing a complex pattern of shock and expansion waves. The end result, due to the asymmetric velocity and pressure distributions, is vectored thrust. Specification of the nozzle contours required for a desired thrust vector angle (an inverse design problem) has been achieved with genetic algorithms. This approach is computationally intensive and prevents the nozzles from being designed in real-time, which is necessary for an operational aircraft system. An investigation was conducted into using genetic algorithms to train a neural network in an attempt to obtain, in real-time, two-dimensional nozzle contours. Results show that genetic algorithm trained neural networks provide a viable, real-time alternative for designing thrust vectoring nozzles contours. Thrust vector angles up to 20 deg were obtained within an average error of 0.0914 deg. The error surfaces encountered were highly degenerate and thus the robustness of genetic algorithms was well suited for minimizing global errors.

Prasanth, Ravi K.↗