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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 307 records · Page 17
All sourcesNASA NTRSDOE OSTI
NASA NTRS1949

Altitude-wind-tunnel investigation of AN-F-58 fuel in experimental version of J47 turbojet engine

An altitude-wind-tunnel investigation of the performance of AN-F-58 and AN-F-32 fuels in an experimental turbojet engine was conducted over a range of simulated altitudes and flight Mach numbers. Combustion efficiencies obtained with AN-F-58 and AN-F-32 fuels were approximately equal. The minimum-speed altitude operational limit was essentially the same with either AN-F-58 or AN-F-32 fuel. Starting characteristics of the two fuels were approximately the same at low wind milling speeds. Visual observation showed no apparent differences in the carbon-deposition rates of AN-F-58 and AN-F-32 fuels.

CHARGING AND CONTROL - GAS TURBINE ENGINES↗
NASA NTRS1949

Altitude performance of AN-F-58 fuels in J35-C-3 single combustor

Three fuel blends conforming to AN-F-58 specifications were tested in order to determine the influence of boiling temperatures and aromatic content on altitude performance in single combustor of 4000-pound-thrust turbojet engine. At simulated engine conditions from an altitude of 30,000 to an altitude of 60,000 feet and 85-percent rated engine speed, the three AN-F-58 fuels showed little difference in performance.

COMBUSTION - EFFECT OF ENGINE OPERATING CONDITIONS↗
NASA NTRS1950

Altitude-test-chamber Investigation of Performance of a 28-inch Ram-jet Engine I : Combustion and Operational Performance of Four Combustion-chamber Configurations

An altitude-test-chamber investigation of a 28-inch-diameter ram-jet engine at a simulated flight Mach number of approximately 2.0 for altitudes of 40,000 to 50,000 feet was conducted at the NACA Lewis laboratory. Three different flame holders, varying in the number and size of the annular gutters, in conjunction with several fuel-injection systems were investigated. The combustion efficiency for the flame-holder fuel-injection system that provided the best over-all operational fuel-air-ratio range (0.03 to 0.075) was over 0.9 at a fuel-air ratio of about 0.065 for the altitude range investigated.

MISSILES, SPECIFIC TYPES↗
NASA NTRS1950

Altitude Performance Characteristics of Tail-pipe Burner with Convergingconical Burner Section on J47 Turbojet Engine

An investigation of turbojet-engine thrust augmentation by means of tail-pipe burning was conducted in the NACA Lewis altitude wind tunnel. Performance data were obtained with a tail-pipe burner having a converging conical burner section installed on an axial-flow-compressor type turbojet engine over a range of simulated flight conditions and tail-pipe fuel-air ratios with a fixed-area exhaust nozzle. A maximum tail-pipe combustion efficiency of 0.86 was obtained at an altitude of 15,000 feet and a flight Mach number of 0.23. Tail-pipe burner operation was possible up to an altitude of 45,000 feet at a flight Mach number of 0.23.

Prince, William R↗
NASA NTRS1950

Effect of fuel volatility on altitude starting limits of a turbojet engine

The effect of fuel volatility on altitude starting limits of an axial-flow-compressor-type turbojet engine was investigated using fuels with Reid vapor pressures of 1.1 and 5.4 pounds per square inch. At flight Mach numbers from 0.40 to 0.85, the AN-F-58 fuel allowed consistent windmilling at altitudes 2000 to 8000 feet higher than was obtained with the 1.1-pound Reid vapor pressure fuel. At a flight Mach number of 0.25, ignition could not be established at any altitude with the lower-volatility fuel.

COMBUSTION - IGNITION OF GASES↗
NASA NTRS1950

Altitude Investigation of Performance of Turbine-propeller Engine and Its Components

An investigation was conducted on a turbine-propeller engine in the NACA Lewis altitude wind tunnel at altitudes from 5000 to 35,000 feet. The applicability of generalized parameters to turbine-propeller engine data, analyses of the compressor, the combustion chambers, and the turbine, and a study of the over-all engine performance are reported. Engine performance data obtained at sea-level static conditions could be used to predict static performance at altitudes up to 35,000 feet by use of the standard generalized parameters.

Wallner, Lewis E↗
NASA NTRS1951

Calculated effects of turbine rotor-blade cooling-air flow, altitude, and compressor bleed point on performance of a turbojet engine

Effects of air-cooling turbine rotor blades on performance of a turbojet engine were calculated for a range of altitudes from sea level to 40,000 feet and a range of coolant flows up to 3 percent of compressor air flow, for two conditions of coolant bleed from the compressor. Bleeding at required coolant pressure resulted in a sea-level thrust reduction approximately twice the percentage coolant flow and in an increase in specific fuel consumption approximately equal to percentage coolant flow. For any fixed value of coolant flow ratio the percentage thrust reduction and percentage increase in specific fuel consumption decreased with altitude. Bleeding coolant at the compressor discharge resulted in an additional 1 percent loss in performance at sea level and in smaller increase in loss of performance at higher altitudes.

TURBINE COOLING↗
NASA NTRS1952

Altitude Performance of Annular Combustor Type Turbojet Engine with JFC-2 Fuel /james W. Useller, James L. Harp, Jr. and Zelmar Barson

An investigation was made comparing the performance of JFC-2 fuel and unleaded, clear gasoline in a 3000-pound-thrust turbojet engine. The JFC-2 fuel was a blend of percent diesel fuel and 25 percent aviation gasoline. Engine combustion efficiency was equal to that obtained with gasoline at rated engine speed and altitudes up to 35,000 feet, but at lower engine speeds or at higher altitudes the JFC-2 fuel gave lower combustion efficiency. No discernible difference was obtained in starting or low-speed combustiion blow-out characteristics of the two fuels. Turbine-discharge radial temperature profiles were nearly the same at altitudes up to 35,000 feet.

Useller, James W↗
NASA NTRS1951

Preliminary results of turbojet-engine altitude-starting investigation

A spark energy of 2.13 joules per spark at 1 spark per second produced ignition to an altitude of 50,000 feet at a flight Mach number of 0.6. The minimum power requirements for ignition were obtained from a combination of low spark repetition rates and high spark energy. The altitude-ignition limit was also increased by increasing spark-gap immersion, fuel temperature, inlet-air temperature, and fuel volatility, and by decreasing flight Mach number. The maximum altitude at which flame propagation was accomplished from combusters with spark plugs to combusters without spark plugs to combustors without spark plugs was increased about 5000 feet by increasing fuel volatility.

COMBUSTION - IGNITION OF GASES↗
NASA NTRS1953

Effect of Diffuser Design, Diffuser-exit Velocity Profile and Fuel Distribution on Altitude Performance of Several Afterburner Configurations

An investigation was conducted in the NACA Lewis altitude wind tunnel to improve the altitude performance and operational characteristics of an afterburner primarily by modifying the diffuser-exit velocity profile by changes in diffuser design and by changing the fuel distribution and the flame holder. Twenty configurations, consisting of combinations of six diffuser geometries, six flame-holder types, and twelve fuel systems, were investigated. Data were obtained over a range of afterburner fuel-air ratios at diffuser-inlet total pressures from 2750 to 620 pounds per square foot. Changes in fuel distribution affected the fuel-air ratio at which peak combustion efficiency occurred as well as the efficiency level. Screeching combustion, which was most prevalent at low altitudes and medium-to-high fuel-air ratios, imposed a restriction on the operable range of a number of configurations.

Diffusers, Subsonic↗
NASA NTRS1920

Effect of altitude on radiator performance

As an airplane rises to high altitudes the decrease in the density and the temperature of the air have important effects on the performance of the radiator. This report gives the results of a study of the effect of reduced pressure and temperature upon the capacity of airplane radiators. A method is presented by which the performance of a radiator at an altitude may be estimated for a particular speed of the airplane at a particular altitude.

James, W S↗
NASA NTRS1948

Altitude-wind-tunnel Investigation of Operational Characteristics of Westinghouse X24C-4B Axial Flow Turbojet Engine

An investigation has been conducted in the NACA Cleveland altitude wind tunnel to evaluate the operational characteristics of a 3000-pound-thrust axial-flow turbojet engine over a range of simulated altitudes from 2000 to 50,000 feet and simulated flight Mach numbers from 0 to 1.04 throughout the operable range of engine speeds. Operational characteristics investigated include engine operating range, acceleration, deceleration, starting, altitude and flight-Mach-number compensation of the fuel-control system, and operation of the lubrication system at high and low ambient-air temperatures.

Hawkins, W Kent↗
NASA NTRS1949

Investigation of Altitude Performance of AN-F-58 Fuels in Annular Combustor of J34-WE-22 Engine

An investigation was conducted in the combustor of a 3000-pound-thrust turbojet engine to determine the altitude performance characteristics of AN-F-58 fuels. Three fuels conforming to AN-F-58 specification were prepared in order to determine the influence of boiling-temperature range and aromatic content on altitude performance. The performance of the three AN-F-58 fuels was compared with the performance of AN-F-48 (grade 100/130) fuel in the range of simulated altitudes from 20,000 to 55,000 feet, corrected engine speeds from 6000to 12,500 rpm, and flight Mach numbers of 0.2 and 0.6.

Tischler, Adelbert O↗
NASA NTRS1948

Altitude-Wind-Tunnel investigation of Westinghouse 19B-2, 19B-8, and 19XB-1 jet-propulsion engines V : combustion chamber performance

Pressure losses through the combustion chamber and the combustion efficiency of the 19B-2 and 19B-8 jet-propulsion engines and the combustion efficiency of the 19XB-1 jet-propulsion engine are presented.Data were obtained from an investigation of the complete engine in the NACA Cleveland altitude wind tunnel over a range of simulated altitudes from 5000 to 30,000 feet and tunnel Mach numbers from less than 0.100 to 0.455. The combustion-chamber pressure loss due to friction was higher for the 19B-2 combustion chamber than for the 19B-8. The 19B-2 combustion chamber had a screen of 40-percent open area interposed between the compressor outlet and the combustion-chamber inlet. The screen for the 19B-8 combustion chamber had a 60-percent open area, which except for a small difference in tail-pipe-nozzle outlet area represents the only point of difference between the standard 19B-2 and 19B-8 combustion chambers. The pressure loss due to heat addition to the flowing gases in the combustion chamber was approximately the same for the 19B-2 and 19B-8 configurations. Altitude and tunnel Mach number had no significant effect on the over-all total-pressure loss through the combustion chamber. A decrease in tail-pipe-nozzle outlet area (tail cone out) resulted in a decrease in combustion-chamber total-pressure loss at high engine speeds.

Boyd, Bemrose↗
NASA NTRS1948

Altitude-Wind-Tunnel investigation of Westinghouse 19B-2, 19B-8, and 19XB-1 Jet-Propulsion Engines IV : analysis of compressor performance

Investigations were conducted in the NACA Cleveland altitude wind tunnel to determine the performance and operational characteristics of the 19B-2, 19B-6, and 19XB-1 Turbojet Engines. One objective of the investigations was to determine the effect of altitude, flight Mach number, and tail-pipe-nozzle area on the performance characteristics of the six-stage and ten-stage axial-flow compressors of the 19B-8 and 19XB-1 engines, respectively. The data were obtained over a range of simulated altitudes and flight Mach numbers. At each simulated flight condition the engine was run over its full operable range of speeds. Performance characteristics of the 19B-8 and 19XB-1 compressors for the range of operation obtainable in the turbojet-engine installation are presented. Compressor characteristics are presented as functions of air flow corrected to sea-level conditions, compressor Mach number, and compressor load coefficient.

Dietz, Robert O↗
NASA NTRS1989

High altitude reconnaissance aircraft design

A Universities Space Research Association (USRA) sponsored (undergraduate) study is presented on the feasibility and design of a high altitude reconnaissance/research aircraft. The aircraft mission was to carry 1,000-3,000 lb of atmospheric pollutant monitoring equipment for 1-5 hr at an altitude of 100,000-130,000 ft. Three configurations subject to the same mission requirements were studied in detail. The three designs analyzed were the tandem-wing-twin-boom, joined wing, and conventional twin-boom configurations. The performance of the three proposed configurations is presented and shows that high altitude flight is possible with current technology. Different possible propulsion systems were investigated and suggestions are made for further investigation and better optimization of the designs.

Poladian, D.↗
NASA NTRS1993

Use of ultrasound in altitude decompression modeling

A model that predicts the probability of developing decompression sickness (DCS) with various denitrogenation schedules is being developed by the Armstrong Laboratory, using human data from previous exposures. It was noted that refinements are needed to improve the accuracy and scope of the model. A commercially developed ultrasonic echo imaging system is being used in this model development. Using this technique, bubbles images from a subject at altitude can be seen in the gall bladder, hepatic veins, vena cava, and chambers of the heart. As judged by their motion and appearance in the vena cava, venous bubbles near the heart range in size from 30 to 300 M. The larger bubbles skim along the top, whereas the smaller ones appear as faint images near the bottom of the vessel. Images from growing bubbles in a model altitude chamber indicate that they grow rapidly, going from 20 to 100 M in 3 sec near 30,000 ft altitude. Information such as this is valuable in verifying those aspects of the DCS model dealing with bubble size, their growth rate, and their site of origin.

Olson, Robert M.↗
NASA NTRS1993

The distribution of large volcanoes on Venus as a function of height and altitude

Theory predicts that the slower cooling of lava flows on Venus should result in lava flows that are typically 20 percent longer than their terrestrial counterparts and that the development of neutral buoyancy zones (NBZ) on Venus may be strongly influenced by altitude-controlled variations in surface pressure. Observations that support these predictions would include relatively low heights for Venus volcanoes, and an increase in both the number and development of large edifices with increasing basal altitude. The results of an analysis of the height and altitude distribution of 123 large (diameter greater than 100 km) volcanoes made using Magellan image and altimetry data are presented and these results are used to begin to test the predications of the above theories.

Keddie, S. T.↗