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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 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS1920

The problem of the turbo-compressor

In terminating the study of the adaptation of the engine to the airplane, we will examine the problem of the turbo-compressor,the first realization of which dates from the war; this will form an addition to the indications already given on supercharging at various altitudes. This subject is of great importance for the application of the turbo-compressor worked by the exhaust gases. As a matter of fact, a compressor increasing the pressure in the admission manifold may be controlled by the engine shaft by means of multiplication gear or by a turbine operated by the exhaust gas. Assuming that the increase of pressure in the admission manifold is the same in both cases, the pressure in the exhaust manifold would be greater in the case in which the compressor is worked by the exhaust gas and there would result a certain reduction of engine power which we must be able to calculate. On the other hand , if the compressor is controlled by the engine shaft, a certain fraction of the excess power supplied is utilized for the rotation of the compressor. In order to compare the two systems, it is there-fore necessary to determine the value of the reduction of power due to back pressure when the turbine is employed.

Devillers, Rene↗
NASA NTRS1945

Air-consumption parameters for automatic mixture control of aircraft engines

Data obtained from Navy calibration tests of an 18-cylinder, two-row, radial engine of 3350-cubic-inch displacement and a 14-cylinder, two-row, radial engine of 2600-cubic-inch displacement (carburetor types) were analyzed to show the correlation between the air consumption of these engines and the parameters that evaluate the air consumption from intake-manifold temperature and pressure, exhaust back pressure, and engine speed.

Shames, Sidney J↗
NASA NTRS1977

Modeling the complete Otto cycle: Preliminary version

A description is given of the equations and the computer program being developed to model the complete Otto cycle. The program incorporates such important features as: (1) heat transfer, (2) finite combustion rates, (3) complete chemical kinetics in the burned gas, (4) exhaust gas recirculation, and (5) manifold vacuum or supercharging. Changes in thermodynamic, kinetic and transport data as well as model parameters can be made without reprogramming. Preliminary calculations indicate that: (1) chemistry and heat transfer significantly affect composition and performance, (2) there seems to be a strong interaction among model parameters, and (3) a number of cycles must be calculated in order to obtain steady-state conditions.

Zeleznik, F. J.↗
NASA NTRS1977

Modeling the complete Otto cycle - Preliminary version

A description is given of the equations and the computer program being developed to model the complete Otto cycle. The program incorporates such important features as: (1) heat transfer, (2) finite combustion rates, (3) complete chemical kinetics in the burned gas, (4) exhaust gas recirculation, and (5) manifold vacuum or supercharging. Changes in thermodynamic, kinetic and transport data as well as model parameters can be made without reprogramming. Preliminary calculations indicate that: (1) chemistry and heat transfer significantly affect composition and performance, (2) there seems to be a strong interaction among model parameters, and (3) a number of cycles must be calculated in order to obtain steady-state conditions.

Zeleznik, F. J.↗
NASA NTRS1947

Effect of Exhaust Pressure on the Cooling Characteristics of a Liquid-Cooled Engine

Data for a liquid-cooled engine with a displacement volume of 1710 cubic inches were analyzed to determine the effect of exhaust pressure on the engine cooling characteristics. The data covered a range of exhaust pressures from 7 to 62 inches of mercury absolute, inlet-manifold pressures from 30 to 50 inches of mercury absolute, engine speeds from 1600 to 3000 rpm, and fuel-air ratios from 0.063 to 0.100. The effect of exhaust pressure on engine cooling was satisfactorily incorporated in the NACA cooling-correlation method as a variation in effective gas temperature with exhaust pressure. Large variations of cylinder-head temperature with exhaust pressure were obtained for operation at constant charge flow. At a constant charge flow of 2 pounds per second (approximately 1000 bhp) and a fuel-air ratio of 0.085, an increase in exhaust pressure from 10 to 60 inches of mercury absolute resulted in an increase of 40 F in average cylinder-head temperature. For operation at constant engine speed and inlet-manifold pressure and variable exhaust pressure (variable charge flow), however, the effect of exhaust pressure on cylinder-head temperature is small. For example, at an inlet-manifold pressure of 40 inches of mercury absolute, an engine speed of 2400 rpm.- and a fuel-air ratio of 0.085, the average cylinder-head temperature was about the same at exhaust pressures of 10 and 60 inches of,mercury absolute; a rise and a subsequent decrease of about 70 occurred between these extremes.

Doyle, Ronald B.↗
NASA NTRS1921

Performance of a Liberty 12 airplane engine

In cooperation with the Engineering Division of the Air Service of the United States Army, a Liberty-12 engine has been tested at the Bureau of Standards. The program of tests was planned to yield that information considered most important in determining the value of the engine for aviation. Full power runs were made at the ground, at 25,000 feet, and at several intermediate altitudes. To determine the mechanical efficiency of the engine, friction horsepower was measured at the ground and at 15,000 feet. As a basis for predicting engine performance with a propeller, a series of tests was made in which the dynamometer load and engine throttle were adjusted at each speed to simulate the engine load which would be imposed at that speed by a propeller operating under normal full load at 1,700 r.p.m. Among the quantities calculated from the test measurements are: brake horsepower; break mean effective pressure; fuel consumption; mixture ratio; mechanical, thermal, and volumetric efficiency; and the percentage of the heat in the fuel appearing in the jacket water and in the exhaust. Jacket water temperature, oil temperature, manifold pressure, etc., are recorded to show the conditions under which the test was made.

Sparrow, S W↗
NASA NTRS1992

Rotary engine performance limits predicted by a zero-dimensional model

A parametric study was performed to determine the performance limits of a rotary combustion engine. This study shows how well increasing the combustion rate, insulating, and turbocharging increase brake power and decrease fuel consumption. Several generalizations can be made from the findings. First, it was shown that the fastest combustion rate is not necessarily the best combustion rate. Second, several engine insulation schemes were employed for a turbocharged engine. Performance improved only for a highly insulated engine. Finally, the variability of turbocompounding and the influence of exhaust port shape were calculated. Rotary engines performance was predicted by an improved zero-dimensional computer model based on a model developed at the Massachusetts Institute of Technology in the 1980's. Independent variables in the study include turbocharging, manifold pressures, wall thermal properties, leakage area, and exhaust port geometry. Additions to the computer programs since its results were last published include turbocharging, manifold modeling, and improved friction power loss calculation. The baseline engine for this study is a single rotor 650 cc direct-injection stratified-charge engine with aluminum housings and a stainless steel rotor. Engine maps are provided for the baseline and turbocharged versions of the engine.

Bartrand, Timothy A.↗
NASA NTRS1978

Heat Resistant Paint

The racing car shown is one of many coated with an inorganic paint that protects "hot parts" of automotive vehicles. Developed and manufactured by Sperex Corporation, Gardena, California, the durable, heat-resistant paint is used on car and truck exhaust systems, firewalls, brake drums and engine manifolds. NASA technology contributed to development of the paint. Sperex was provided a technical support packa'ge detailing the research of Goddard Space Flight Center on long-life inorganic coatings. The information helped Sperex perfect its own formulations.

Source record↗
NASA NTRS2011

Modeling of Supersonic Film Cooling on the J-2X Nozzle Extension

Supersonic film cooling (SSFC) of nozzles has been used in several liquid rocket engine designs, and is being applied to the nozzle extension (NE) of the J-2X upper stage engine currently under development. Turbine exhaust gas (TEG) is injected tangentially from a manifold along the NE, and provides a thermal barrier from the core nozzle flow for the NE. As the TEG stream mixes with the nozzle flow, the effectiveness of the thermal barrier is reduced. This paper documents computational fluid dynamics (CFD) analysis work performed by NASA Marshall Space Flight Center (MSFC) to model the flow of the TEG through the manifold, into the nozzle, and the subsequent mixing of the TEG stream with the core flow. The geometry and grid of the TEG manifold, structural support ribs, and the NE wall will be shown, and the CFD boundary conditions described. The Loci-CHEM CFD code used in this work will also be briefly described. A unique approach to modeling the combined TEG manifold/thrust chamber assembly (TCA) was employed, as it was not practical to model the entire 360 circumferential range in one simulation. Prior CFD validation work modeling Calspan SSFC experiments in the early 1990s, documented in a previous AIAA paper, will also be briefly discussed. The fluid dynamics of the TEG flow through the manifold, into and between the structural support ribs, and into the nozzlette that feeds the TCA will be described. Significant swirl and non-uniformities are present, which along with the wakes from the ribs, act to degrade the film cooling effectiveness compared to idealized injection of TEG gas. The effect of these flow characteristics on the adiabatic wall temperature profile on the NE will be discussed.

Ruf, Joseph H.↗
NASA NTRS1949

Correlation of cylinder-head temperatures and coolant heat rejections of a multicylinder, liquid-cooled engine of 1710-cubic-inch displacement

Data obtained from an extensive investigation of the cooling characteristics of four multicylinder, liquid-cooled engines have been analyzed and a correlation of both the cylinder-head temperatures and the coolant heat rejections with the primary engine and coolant variables was obtained. The method of correlation was previously developed by the NACA from an analysis of the cooling processes involved in a liquid-cooled-engine cylinder and is based on the theory of nonboiling, forced-convection heat transfer. The data correlated included engine power outputs from 275 to 1860 brake horsepower; coolant flows from 50 to 320 gallons per minute; coolants varying in composition from 100 percent water to 97 percent ethylene glycol and 3 percent water; and ranges of engine speed, manifold pressure, carburetor-air temperature, fuel-air ratio, exhaust-gas pressure, ignition timing, and coolant temperature. The effect on engine cooling of scale formation on the coolant passages of the engine and of boiling of the coolant under various operating conditions is also discussed.

Lundin, Bruce T↗
NASA NTRSDate not supplied

Overview of the SLS Core Stage Thrust Vector Control System Design

The Space Launch System (SLS) Core Stage (CS) Thrust Vector Control (TVC) consists of four independent hydraulic systems. The SLS CS TVC system is comprised of 8 mechanical feedback Shuttle heritage Type III TVC actuators and four RS-25 engines, each attached to a Shuttle heritage gimbal block/bearing. Each hydraulic system nominally provides hydraulic power to one RS-25 engine and two actuators. Additionally, each system provides redundant control capability to one actuator on each of its neighboring systems. The RS-25 uses hydraulic power to control propellant valves, and the TVC actuators are used to move the engine in the pitch and yaw gimbal planes. The TVC system design leverages hardware from the Space Shuttle program as well as new hardware designed specifically for the Core Stage. The Space Shuttle heritage hardware directly reused on SLS includes the Orbiter TVC hydraulic servo-actuators (with two slight design modifications), the Orbiter hydraulic circulation pumps, the Orbiter gimbal block/bearing, and the Solid Rocket Booster hydraulic pumps. The Solid Rocket Booster APU turbines are powered by hot gas produced by a catalyzed hydrazine decomposition. The SLS Core Auxiliary Power Unit (CAPU) is derived from the Space Shuttle Orbiter Auxiliary Power Unit (APU); on the SLS Core Stage, the CAPU turbine is spun using cold gas tapped-off from the RS-25 to CS liquid hydrogen autogenous pressurization line. The remaining hardware in the TVC system (hydraulic Filter Manifold (FM), hydraulic Supply Accumulator (SA), hydraulic Return Accumulator (RA), Hydraulic Reservoir, Exhaust Gas Heat Exchanger (EGHE)) as well as the avionics providing control and telemetry (TVC Actuator Controller (TAC) and CAPU Controller (CAPUC) are new components developed for SLS. This paper is the first installment in a seven-paper series surveying the design, engineering, test validation, and flight performance of the Core Stage Thrust Vector Control system. In this paper, the overall design architecture of the CS TVC is presented, with a focus on the interfaces between the TVC actuators, the engines, their hydraulic power systems, and the avionics that provide commands from the SLS Vehicle Management (VM) software to effect stable and robust flight control for the integrated SLS launch vehicle.

Thrust Vector Control↗
NASA NTRSDate not supplied

Overview of the SLS Core Stage Thrust Vector Control System Design

The Space Launch System (SLS) Core Stage (CS) Thrust Vector Control (TVC) consists of four independent hydraulic systems. The SLS CS TVC system is comprised of 8 mechanical feedback Shuttle heritage Type III TVC actuators and four RS-25 engines, each attached to a Shuttle heritage gimbal block/bearing. Each hydraulic system nominally provides hydraulic power to one RS-25 engine and two actuators. Additionally, each system provides redundant control capability to one actuator on each of its neighboring systems. The RS-25 uses hydraulic power to control propellant valves, and the TVC actuators are used to move the engine in the pitch and yaw gimbal planes. The TVC system design leverages hardware from the Space Shuttle program as well as new hardware designed specifically for the Core Stage. The Space Shuttle heritage hardware directly reused on SLS includes the Orbiter TVC hydraulic servo-actuators (with two slight design modifications), the Orbiter hydraulic circulation pumps, the Orbiter gimbal block/bearing, and the Solid Rocket Booster hydraulic pumps. The Solid Rocket Booster APU turbines are powered by hot gas produced by a catalyzed hydrazine decomposition. The SLS Core Auxiliary Power Unit (CAPU) is derived from the Space Shuttle Orbiter Auxiliary Power Unit (APU); on the SLS Core Stage, the CAPU turbine is spun using cold gas tapped-off from the RS-25 to CS liquid hydrogen autogenous pressurization line. The remaining hardware in the TVC system (hydraulic Filter Manifold (FM), hydraulic Supply Accumulator (SA), hydraulic Return Accumulator (RA), Hydraulic Reservoir, Exhaust Gas Heat Exchanger (EGHE)) as well as the avionics providing control and telemetry (TVC Actuator Controller (TAC) and CAPU Controller (CAPUC) are new components developed for SLS. This paper is the first installment in a seven-paper series surveying the design, engineering, test validation, and flight performance of the Core Stage Thrust Vector Control system. In this paper, the overall design architecture of the CS TVC is presented, with a focus on the interfaces between the TVC actuators, the engines, their hydraulic power systems, and the avionics that provide commands from the SLS Vehicle Management (VM) software to effect stable and robust flight control for the integrated SLS launch vehicle.

Thrust Vector Control↗
NASA NTRS1940

Flight Tests of Exhaust Gas Jet Propulsion, Special Report

Flight test s were conducted on the XP-41 airplane, equipped with a Pratt & Whitney R1830-19, 14-cylinder, air-cooled engine, to determine the increase in flight speed obtainable by the use of individual exhaust stacks directed rearwardly to obtain exhaust-gas thrust. Speed increases up to 18 miles per hour at 20,000 feet altitude were obtained using stacks having an exit area of 3.42 square inches for each cylinder. A slight increase in engine power and decrease in cylinder temperature at a given manifold pressure were obtained with the individual stacks as compared with a collector-ring installation. Exhaust-flame visibility was quite low, particularly in the rich range of fuel-air ratios.

Pnkel, Benjamin↗
NASA NTRS1966

Elimination of rocket engine asymmetric loads during tests at sea level

Secondary injection concept eliminates asymmetric loads and may increase thrust rocket engine loads during sea level tests. The concept uses either a tubular manifold with evenly spaced injection ports or secondary fluid injected at the turbine exhaust inlet to the thrust chamber.

Johnson, J. R.↗
NASA NTRS1985

Calculation of flow about posts and powerhead model

A three dimensional analysis of the non-uniform flow around the liquid oxygen (LOX) posts in the Space Shuttle Main Engine (SSME) powerhead was performed to determine possible factors contributing to the failure of the posts. Also performed was three dimensional numerical fluid flow analysis of the high pressure fuel turbopump (HPFTP) exhaust system, consisting of the turnaround duct (TAD), two-duct hot gas manifold (HGM), and the Version B transfer ducts. The analysis was conducted in the following manner: (1) modeling the flow around a single and small clusters (2 to 10) of posts; (2) modeling the velocity field in the cross plane; and (3) modeling the entire flow region with a three dimensional network type model. Shear stress functions which will permit viscous analysis without requiring excessive numbers of computational grid points were developed. These wall functions, laminar and turbulent, have been compared to standard Blasius solutions and are directly applicable to the cylinder in cross flow class of problems to which the LOX post problem belongs.

Anderson, P. G.↗
NASA NTRS2001

NASA's Atmospheric Effects of Aviation Project: Results of the August 1999 Aerosol Measurement Intercomparison Workshop, Laboratory Phase

During August 1-14, 1999, NASA's Atmospheric Effects of Aviation Project (AEAP) convened a workshop at the NASA Langley Research Center to try to determine why such a wide variation in aerosol emissions indices and chemical and physical properties have been reported by various independent AEAP-supported research teams trying to characterize the exhaust emissions of subsonic commercial aircraft. This workshop was divided into two phases, a laboratory phase and a field phase. The laboratory phase consisted of supplying known particle number densities (concentrations) and particle size distributions to a common manifold for the participating research teams to sample and analyze. The field phase was conducted on an aircraft run-up pad. Participating teams actually sampled aircraft exhaust generated by a Langley T-38 Talon aircraft at 1 and 9 m behind the engine at engine powers ranging from 48 to 100 percent. Results from the laboratory phase of this intercomparison workshop are reported in this paper.

Cofer, W. Randy, III↗
NASA NTRS2001

NASA's Atmospheric Effects of Aviation Project

During August 1-14, 1999, NASA's Atmospheric Effects of Aviation Project (AEAP) convened a workshop at the NASA Langley Research Center to try to determine why such a wide variation in aerosol emissions indices and chemical and physical properties has been reported by various independent AEAP-supported research teams trying to characterize the exhaust emissions of subsonic commercial aircraft. This workshop was divided into two phases, a laboratory phase and a field phase. The laboratory phase consisted of supplying known particle number densities (concentrations) and particle size distributions to a common manifold for the participating research teams to sample and analyze. The field phase was conducted on an aircraft run-up pad. Participating teams actually sampled aircraft exhaust generated by a Langley T-38 Talon aircraft at 1 and 9 m behind the engine at engine powers ranging from 48 to 100 percent. Results from the laboratory phase of this intercomparison workshop are reported in this paper.

Cofer, W. Randy, III↗