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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 487 records · Page 27
All sourcesNASA NTRSDOE OSTI
NASA NTRS1995

Robotic Planetary Science Missions Enabled with Small NTR Engine/Stage Technologies

The high specific impulse (Isp) and engine thrust-to-weight ratio of liquid hydrogen (LH2)-cooled nuclear thermal rocket (NTR) engines makes them ideal for upper stage applications to difficult robotic planetary science missions. A small 15 thousand pound force (klbf) NTR engine using a uranium-zirconium-niobium 'ternary carbide' fuel (Isp approximately 960 seconds at approximately 3025K) developed in the Commonwealth of Independent States (CIS) is examined and its use on an expendable injection stage is shown to provide major increases in payload delivered to the outer planets (Saturn, Uranus, Neptune and Pluto). Using a single 'Titan IV-class' launch vehicle, with a lift capability to low Earth orbit (LEO) of approximately 20 metric tons (t), an expendable NTR upper stage can inject two Pluto 'Fast Flyby' spacecraft (PFF/SC) plus support equipment-combined mass of approximately 508 kg--on high energy, '6.5-9.2 year' direct trajectory missions to Pluto. A conventional chemical propulsion mission would use a liquid oxygen (LOX)/LH2 'Centaur' upper stage and two solid rocket 'kick motors' to inject a single PFF/SC on the same Titan IV launch vehicle. For follow on Pluto missions, the NTR injection stage would utilize a Jupiter 'gravity assist' (JGA) maneuver to launch a LOX/liquid methane (CH4) capture stage (Isp approximately 375 seconds) and a Pluto 'orbiter' spacecraft weighing between approximately 167-312 kg. With chemical propulsion, a Pluto orbiter mission is not a viable option because c inadequate delivered mass. Using a 'standardized' NTR injection stage and the same single Titan IV launch scenario, 'direct flight' (no gravity assist) orbiter missions to Saturn, Uranus and Neptune are also enabled with transit times of 2.3, 6.6, and 12.6 years, respectively. Injected mass includes a storable, nitrogen tetroxide/monomethyl hydrazine (N2O4/MMH) capture stage (Isp approximately 330 seconds) and orbiter payloads 340 to 820% larger than that achievable using a LOX/LH2-fueled injection stage. The paper discusses NTR technology and mission characteristics, shows NTR stage and payload accommodations within the 26.2 m long Titan IV payload fairing, and discusses NTR stage performance as a function of assumed cryogenic tank technology.

Borowski, Stanley K.↗
NASA NTRS1998

The Control System for the X-33 Linear Aerospike Engine

The linear aerospike engine is being developed for single-stage -to-orbit (SSTO) applications. The primary advantages of a linear aerospike engine over a conventional bell nozzle engine include altitude compensation, which provides enhanced performance, and lower vehicle weight resulting from the integration of the engine into the vehicle structure. A feature of this integration is the ability to provide thrust vector control (TVC) by differential throttling of the engine combustion elements, rather than the more conventional approach of gimballing the entire engine. An analysis of the X-33 flight trajectories has shown that it is necessary to provide +/- 15% roll, pitch and yaw TVC authority with an optional capability of +/- 30% pitch at select times during the mission. The TVC performance requirements for X-33 engine became a major driver in the design of the engine control system. The thrust level of the X-33 engine as well as the amount of TVC are managed by a control system which consists of electronic, instrumentation, propellant valves, electro-mechanical actuators, spark igniters, and harnesses. The engine control system is responsible for the thrust control, mixture ratio control, thrust vector control, engine health monitoring, and communication to the vehicle during all operational modes of the engine (checkout, pre-start, start, main-stage, shutdown and post shutdown). The methodology for thrust vector control, the health monitoring approach which includes failure detection, isolation, and response, and the basic control system design are the topic of this paper. As an additional point of interest a brief description of the X-33 engine system will be included in this paper.

Jackson, Jerry E.↗
NASA NTRS1959

Flight Tests of A 1/8-Scale Model of the Bell D-188A Jet VTOL Airplane

The Bell D-188A VTOL airplane is a horizontal-attitude VTOL fighter with tilting engine nacelles at the tips of a low-aspect-ratio unswept wing and additional engines in the fuselage. The model could be flown smoothly in hovering and transition flight. In forward flight the model could be flown smoothly at the lower angles of attack but experienced an uncontrollable directional divergence at angles of attack above about 16 deg.

Smith, Charles C., Jr.↗
NASA NTRS1987

RL10A-3-3B high mixture ratio qualification program

The results of the high mixture ratio qualification testing of the RL10 engine for the Shuttle/Centaur program are presented. The objective of the engine qualification test was to demonstrate the suitability of the RL10A-3-3B engine for space vehicle flight by subjecting it to the testing specified in RL10A-3-3B Model Specification Number 2295 dated February 1986. The applicable section of the specification is presented. Due to payload volume advantages which can be achieved by increasing the operating mixture ratio of the RL10, a decision was made to qualify the engine to run at a higher mixture ratio. A program was created to qualify the RL10 engine for operation at 15,000 pounds thrust and a nominal 6.0 to 1 mixture ratio. This model of the engine was designated the RL10A-3-3B. The qualification program included three test series as follows: (1) hardware durability and limits test in which the engine completed 23 firings and 4605.7 seconds with 1588.7 seconds at less than 6.6 mixture ratio; (2) preliminary qualification test in which the engine completed 26 firings and 5750 seconds; and (3) qualification test in which the engine completed 26 hot firings and 5693.4 seconds with 905.9 seconds at 6.7 mixture ratio. Several changes in engine hardware were required for operation of the RL10A-3-3B engine in the Space Shuttle which include a duel pressure switch ignition, an oxidizer flow control, and helium plumbing changes.

Vogel, T.↗
NASA NTRS1994

Measurements of Forward Flight Effects on the Advanced Ducted Propulsion Demonstrator Engine

The performance of the Pratt & Whitney Advanced Ducted Propulsion (ADP) UHB concept has been recently evaluated with studies of a 17 in. diameter fan simulator. Following the model scale tests, a 118 in. diameter demonstrator was tested at the NASA Ames 40- by 80-Foot Wind Tunnel. The 18 blade fan was driven by the low compressor shaft of a PW2037 core through a reduction gear system fabricated by Fiat with approximately 1:3.7 reduction ratio. ne variable pitch fan was hydraulically actuated with settings for take-off, cruise, feather, and reverse thrust. The low-pressure turbine was built by MTU to provide higher shaft power in comparison with the standard PW2037. The demonstrator was provided with 45 vanes located 2.6 fan chords downstream of the rotor, and 10 case struts approximately 1 fan chord downstream of the vanes. The inlet, mid-duct, and exhaust linings were acoustically treated. Acoustic surveys were taken in the for-ward thrust mode for fan speeds of 898, 1120, 1205, and 1302 R.P.M., and at tunnel speeds of 25, 50, 100, and 140 kts. The lowest speed was achieved with the wind tunnel fans at flat pitch, but with the engine pumping the test section Microphone signals were recorded for 30 seconds at 5 deg. increments. These measurements will be used to assess the effects of forward speed on UHB engines, to compare these effects with the corresponding characteristics of conventional bypass ratio engines, and to discuss the various aspects of testing large engines in the wind tunnel.

Horne, W. C.↗
NASA NTRS1993

Two stage turbine for rockets

The aerodynamic design and rig test evaluation of a small counter-rotating turbine system is described. The advanced turbine airfoils were designed and tested by Pratt & Whitney. The technology represented by this turbine is being developed for a turbopump to be used in an advanced upper stage rocket engine. The advanced engine will use a hydrogen expander cycle and achieve high performance through efficient combustion of hydrogen/oxygen propellants, high combustion pressure, and high area ratio exhaust nozzle expansion. Engine performance goals require that the turbopump drive turbines achieve high efficiency at low gas flow rates. The low mass flow rates and high operating pressures result in very small airfoil heights and diameters. The high efficiency and small size requirements present a challenging turbine design problem. The shrouded axial turbine blades are 50 percent reaction with a maximum thickness to chord ratio near 1. At 6 deg from the tangential direction, the nozzle and blade exit flow angles are well below the traditional design minimum limits. The blade turning angle of 160 deg also exceeds the maximum limits used in traditional turbine designs.

Veres, Joseph P.↗
NASA NTRS1987

Unducted-Fan Engine

Unducted-fan (UDF(TM) engine advanced counterrotating-pusher-propeller propulsion system for high-subsonic (mach 0.7 to 0.85) aircraft. Combines modern high-pressure-ratio gas-turbine engine with multistage counterrotating power turbine directly coupled to counterrotating, advanced high-speed propellers. Key feature of system is unique direct turbine drive eliminating need for gearbox to transmit power to propeller blades.

Meleason, Edward T.↗
NASA NTRS2004

Advanced Health Management System for the Space Shuttle Main Engine

Boeing-Canoga Park (BCP) and NASA-Marshall Space Flight Center (NASA-MSFC) are developing an Advanced Health Management System (AHMS) for use on the Space Shuttle Main Engine (SSME) that will improve Shuttle safety by reducing the probability of catastrophic engine failures during the powered ascent phase of a Shuttle mission. This is a phased approach that consists of an upgrade to the current Space Shuttle Main Engine Controller (SSMEC) to add turbomachinery synchronous vibration protection and addition of a separate Health Management Computer (HMC) that will utilize advanced algorithms to detect and mitigate predefined engine anomalies. The purpose of the Shuttle AHMS is twofold; one is to increase the probability of successfully placing the Orbiter into the intended orbit, and the other is to increase the probability of being able to safely execute an abort of a Space Transportation System (STS) launch. Both objectives are achieved by increasing the useful work envelope of a Space Shuttle Main Engine after it has developed anomalous performance during launch and the ascent phase of the mission. This increase in work envelope will be the result of two new anomaly mitigation options, in addition to existing engine shutdown, that were previously unavailable. The added anomaly mitigation options include engine throttle-down and performance correction (adjustment of engine oxidizer to fuel ratio), as well as enhanced sensor disqualification capability. The HMC is intended to provide the computing power necessary to diagnose selected anomalous engine behaviors and for making recommendations to the engine controller for anomaly mitigation. Independent auditors have assessed the reduction in Shuttle ascent risk to be on the order of 40% with the combined system and a three times improvement in mission success.

Davidson, Matt↗
NASA NTRS1986

A parametric study of a gas-generator airturbo ramjet (ATR)

Parametric engine performance calculations were carried out for an airturbo ramjet (ATR). A LOX-LH2 rocket powered turbine powered the compressor. The engine was flown over a typical flight path up to Mach 5 to show the effect of engine off design operation. The compressor design efficiency, compressor pressure ratio, rocket turbine efficiency, rocket turbine inlet temperature, and rocket chamber pressure were varied to show their effect on engine net thrust and specific impulse at Mach 5 cruise. Estimates of engine weights as a fucntion of the ratio of compressor air to rocket propellant flow and rocket champer pressure are also included. In general, the Mach 5 results indicate that increasing the amount of rocket gas produced increased thrust but decreased the specific impulse. The engine performance was fairly sensitive to rocket chamber pressure, especially at higher compressor pressure ratios. At higher compressor pressure ratios, the engine thrust was sensitive to turbine inlet temperature. At all compressor pressure ratios, the engine performance was not sensitive to compressor or turbine efficiency.

Snyder, C. A.↗
NASA NTRS1986

A parametric study of a gas-generator airturbo ramjet (ATR)

Parametric engine performance calculations were carried out for an airturbo ramjet (ATR). A LOX-LH2 rocket powered turbine powered the compressor. The engine was flown over a typical flight path up to Mach 5 to show the effect of engine off design operation. The compressor design efficiency, compressor pressure ratio, rocket turbine efficiency, rocket turbine inlet temperature, and rocket chamber pressure were varied to show their effect on engine net thrust and specific impulse at Mach 5 cruise. Estimates of engine weights as a function of the ratio of compressor air to rocket propellant flow and rocket chamber pressure are also included. In general, the Mach 5 results indicate that increasing the amount of rocket gas produced increased thrust but decreased the specific impulse. The engine performance was fairly sensitive to rocket chamber pressure, especially at higher compressor pressure ratios. At higher compressor pressure ratios, the engine thrust was sensitive to turbine inlet temperature. At all compressor pressure ratios, the engine performance was not sensitive to compressor or turbine efficiency.

Snyder, Christopher A.↗
NASA NTRS1947

Investigation of the I-40 Jet-Propulsion Engine in the Cleveland Altitude Wind Tunnel. V - Operational Characteristics: Operational Characteristics - 5

An investigation has been conducted in the Cleveland altitude wind tunnel to determine the operational characteristics of the I-40 jet-propulsion engine over a range of pressure altitudes from 10,000 to 50,000 feet and ram-pressure ratios from 1.00 to 1.76. Engine operational data were obtained with the engine in the standard configuration and with various modifications of the fuel system, the electrical system, and the combustion chambers. The effects of altitude and airspeed on operating speed range, starting, windmilli.ng, acceleration, speed regulation, cooling, and vibration of the standard and modified engines were determined, and damage to parts was noted. Maximum engine speed was obtainable at all altitudes and airspeeds wi th each fuel-control system investigated. The minimum idling speed was raised by increases in altitude and airspeed. The lowest minimum stable speeds were obtained with the standard configuration using 40-gallon nozzles with individual metering plugs. The engine was started normally at altitudes as high as 20,000 feet with all of the fuel systems and ignition combinations except one. Ignition at 70,000 feet was difficult and, although successful ignition occurred, acceleration was slow and usually characterized by excessive tail-pipe temperature. During windmilling investigations of the engine equipped with the standard fuel system, the engine could not be started at ram-pressure ratios of 1.1 to 1.7 at altitudes of 10,000, 20,000 and 30,000 feet. When equipped with the production barometric and Monarch 40-gallon nozzles, the engine accelerated in 12 seconds from an engine speed of 6000 rpm to 11,000 rpm at 20,000 feet and an average tail-pipe temperature of 11000 F. At the same altitude and temperature, all the engine configurations had approximately the same rate of acceleration. The Woodward governor produced the safest accelerations, inasmuch as it could be adjusted to automatically prevent acceleration blow out. The engine speed was held constant by the Woodward governor and the Edwards regulator during simulated dives and climbs at constant throttle position. The bearing cooling system was satisfactory at all altitudes and airspeeds. The engines operated without serious failure, although the exhaust cone, the tail pipe, and the airplane fuselage were damaged during altitude starts.

Golladay, Richard L.↗
NASA NTRS1984

Evaluation of dissociated and steam-reformed methanol as automotive engine fuels

Dissociated and steam reformed methanol were evaluated as automotive engine fuels. Advantages and disadvantages in using methanol in the reformed rather than liquid state were discussed. Engine dynamometer tests were conducted with a four cylinder, 2.3 liter, spark ignition automotive engine to determine performance and emission characteristics operating on simulated dissociated and steam reformed methanol (2H2 + CO and 3H2 + CO2 respectively), and liquid methanol. Results are presented for engine performance and emissions as functions of equivalence ratio, at various throttle settings and engine speeds. Operation on dissociated and steam reformed methanol was characterized by flashback (violent propagation of a flame into the intake manifold) which limited operation to lower power output than was obtainable using liquid methanol. It was concluded that: an automobile could not be operated solely on dissociated or steam reformed methanol over the entire required power range - a supplementary fuel system or power source would be necessary to attain higher powers; the use of reformed mechanol, compared to liquid methanol, may result in a small improvement in thermal efficiency in the low power range; dissociated methanol is a better fuel than steam reformed methanol for use in a spark ignition engine; and use of dissociated or steam reformed methanol may result in lower exhaust emissions compared to liquid methanol.

Lalk, T. R.↗
NASA NTRS1986

HYTESS-Hypothetical Turbofan-Engine Simplified Simulation

Simulated characteristics mimic those of advanced turbofan engines. Computer program developed to offer those interested in engine dynamics and controls research efficient, realistic, and easily-used engine simulation. Simulation developed from linearized operating-point models but still retains essential nonlinear engine effects. Representative of hypothetical, low-bypass-ratio, twin-spool, axialflow turbofan engine. Program written in FORTRAN IV.

Merrill, W.↗
NASA NTRS1985

Engineering and programming manual: Two-dimensional kinetic reference computer program (TDK)

The Two Dimensional Kinetics (TDK) computer program is a primary tool in applying the JANNAF liquid rocket thrust chamber performance prediction methodology. The development of a methodology that includes all aspects of rocket engine performance from analytical calculation to test measurements, that is physically accurate and consistent, and that serves as an industry and government reference is presented. Recent interest in rocket engines that operate at high expansion ratio, such as most Orbit Transfer Vehicle (OTV) engine designs, has required an extension of the analytical methods used by the TDK computer program. Thus, the version of TDK that is described in this manual is in many respects different from the 1973 version of the program. This new material reflects the new capabilities of the TDK computer program, the most important of which are described.

Nickerson, G. R.↗
NASA NTRSDate not supplied

Flametube Evaluation of a Lean-Lean Combustor Concept Developed for Supersonic Cruise Aircraft

Gaseous emissions were measured in single-cup flametube tests of an advanced low-NOx combustor concept at simulated supersonic cruise conditions. The combustor concept is a low technology readiness level (TRL), lean front-end design developed under the NASA Fundamental Aeronautics/Supersonics project to minimize NOx emissions at supersonic cruise. The flametube conditions matched or approached combustor conditions at supersonic cruise, with combustor inlet temperatures up to 920 K, inlet pressures up to 19 bar, and combusted gas temperatures up to 2,120 K. Whether these conditions met or just approached supersonic cruise conditions depended on the type of engine the combustor would be installed in. Two types of engines were considered here: a “derivative” engine based on a current technology and an “advanced” engine with a higher operating pressure ratio and higher temperature limits. For the “derivative” engine, the combustor is expected to be at least close to meeting the NASA NOx emissions goal of 10 g-NOx/kg-fuel at supersonic cruise. However, with the higher combustor inlet and flame temperatures of the advanced engine, NOx emissions are expected to be well above the goal.

combustion↗
NASA NTRS2022

Flametube Evaluation of a Lean-Lean Combustor Concept Developed for Supersonic Cruise Aircraft

Gaseous emissions were measured in single-cup flametube tests of an advanced low-NO x combustor concept at simulated supersonic cruise conditions. The combustor concept is a low technology readiness level (TRL), lean front-end design developed under the NASA Fundamental Aeronautics/Supersonics project to minimize NO x emissions at supersonic cruise. The flametube conditions matched or approached combustor conditions at supersonic cruise, with combustor inlet temperatures up to 920 K, inlet pressures up to 19 bar, and combusted gas temperatures up to 2,120 K. Whether these conditions met or just approached supersonic cruise conditions depended on the type of engine the combustor would be installed in. Two types of engines were considered here: a "derivative" engine based on a current technology and an "advanced" engine with a higher operating pressure ratio and higher temperature limits. For the "derivative" engine, the combustor is expected to be at least close to meeting the NASA NO x emissions goal of 10 g-NO x /kg-fuel at supersonic cruise. However, with the higher combustor inlet and flame temperatures of the advanced engine, NO x emissions are expected to be well above the goal.

supersonic cruise↗
NASA NTRS1987

SPRE 1 free-piston Stirling engine testing at NASA Lewis Research Center

As part of the NASA funded portion of the SP-100 Advanced Technology Program the Space Power Research Engine (SPRE 1) was designed and built to serve as a research tool for evaluation and development of advanced Stirling engine concepts. The SPRE 1 is designed to produce 12.5 kW electrical power when operated with helium at 15 MPa and with an absolute temperature ratio of two. The engine is now under test in a new test facility which was designed and built at NASA Lewis specifically to test the SPRE 1. The SPRE 1, the NASA test facility, the initial SPRE 1 test results, and future SPRE 1 test plans are described.

Cairelli, James E.↗