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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 181 records · Page 10
All sourcesNASA NTRSDOE OSTI
NASA NTRS1947

Effect of Fuel on Performance of a Single Combustor of an I-16 Turbojet Engine at Simulated Altitude Conditions

As part of a study of the effects of fuel composition on the combustor performance of a turbojet engine, an investigation was made in a single I-16 combustor with the standard I-16 injection nozzle, supplied by the engine manufacturer, at simulated altitude conditions. The 10 fuels investigated included hydrocarbons of the paraffin olefin, naphthene, and aromatic classes having a boiling range from 113 degrees to 655 degrees F. They were hot-acid octane, diisobutylene, methylcyclohexane, benzene, xylene, 62-octane gasoline, kerosene, solvent 2, and Diesel fuel oil. The fuels were tested at combustor conditions simulating I-16 turbojet operation at an altitude of 45,000 feet and at a rotor speed of 12,200 rpm. At these conditions the combustor-inlet air temperature, static pressure, and velocity were 60 degrees F., 12.3 inches of mercury absolute, and 112 feet per second respectively, and were held approximately constant for the investigation. The reproducibility of the data is shown by check runs taken each day during the investigation. The combustion in the exhaust elbow was visually observed for each fuel investigated.

Zettle, Eugene V↗
NASA NTRS1974

Modularized instrument system for turbojet engine test facilities

A new modular instrument system is being developed to handle the many channels of data commonly encountered in turbojet engine testing. Each module contains a group of transducers and all the signal conditioning, multiplexing, and digitizing electronics necessary for direct interface with a digital computer. The digital interface within each module is the same for all modules. Each module provides a controlled environment for its contents. A minicomputer in the control room gathers the data, performs some on-line calculation and display, and interfaces with a shared recording and computing system. The advantages of this system are (1) reduced manpower for system installation, setup, and checkout; (2) standardized equipment interfaces; (3) increased reliability through automatic system testing and through minimization of manual adjustments; and (4) reduced cost through minimization of wiring and simplification of control room display.

Nieberding, W. C.↗
NASA NTRS1976

Effects of forward velocity on noise for a J85 turbojet engine with multitube suppressor from wind tunnel and flight tests

Flight and wind tunnel noise tests were conducted using a J85 turbojet engine as a part of comprehensive programs to obtain an understanding of forward velocity effects on jet exhaust noise. Nozzle configurations of primary interest were a 104-tube suppressor with and without an acoustically-treated shroud. The installed configuration of the engine was as similar as possible in the flight and wind tunnel tests. Exact simultaneous matching of engine speed, exhaust velocity, and exhaust temperature was not possible, and the wind tunnel maximum Mach number was approximately 0.27, while the flight Mach number was approximately 0.37. The nominal jet velocity range was 450 to 640 m/sec. For both experiments, background noise limited the jet velocity range for which significant data could be obtained. In the present tests the observed directivity and forward velocity effects for the suppressor are more similar to predicted trends for internally-generated noise than unsuppressed jet noise.

Stone, J. R.↗
NASA NTRS1982

Performance of a 2D-CD nonaxisymmetric exhaust nozzle on a turbojet engine at altitude

Baseline thrust and cooling data obtained with a 2D-CD versatile research exhaust nozzle mounted on a turbojet engine in an altitude chamber are presented. The tests covered a range of nozzle pressure ratios, nozzle pressure ratios, nozzle throat areas, and internal expansion area ratios. The thrust data obtained show good agreement with theory and scale model results after correcting the data for leakage and bypass cooling flows. Additional work is needed to improve predictability of cooling performance.

Straight, D. M.↗
NASA NTRS1952

Performance of the Components of the XJ34-WE-32 Turbojet Engine over a Range of Engine and Flight Conditions

Performance of the compressor, combustor, and turbine operating as integral parts of the XJ34-WE-32 turbojet engine was determined in the Lewis altitude wind tunnel over a range of altitudes from 5000 to 55,000 feet and flight Mach numbers from 0.28 to 1.05. Data were obtained for each of four exhaust-nozzle areas and are presented in graphical and tabular form.

ENGINES, TURBO-JET↗
NASA NTRS1956

Analysis of turbojet-engine controls for afterburning starting

A simulation procedure is developed for studying the effects of an afterburner start on a controlled turbojet engine. The afterburner start is represented by introducing a step decrease in the effective exhaust-nozzle area, after which the control returns the controlled engine variables to their initial values. The degree and speed with which the control acts are a measure of the effectiveness of the particular control system. Data are presented from five systems investigated using an electronic analog computer and the developed simulation procedure. These systems are compared with respect to steady-state errors, speed of response, and transient deviations of the system variables.

Phillips, W E , Jr↗
NASA NTRS1949

Altitude-wind-tunnel Investigation of J47 Turbojet-engine Performance

An investigation has been conducted in the NACA Lewis altitude wind tunnel to evaluate the performance of the J47 turbojet engine over a range of simulated altitudes from 5000 to 50,000 feet, simulated flight Mach numbers from 0.21 to 0.97, and a complete range of engine speeds. Data are presented to show the effects of altitude at a flight Mach number of 0.21 and of flight Mach number at an altitude of 25,000 feet. The performance data are generalized by two methods to determine the range of flight conditions for which engine performance may be predicted from performance data obtained at a given flight condition.

Conrad, E William↗
NASA NTRS1954

Preliminary Data on the Effects of Inlet Pressure Distortions on the J57-P-1 Turbojet Engine

An investigation to determine the steady-state and surge characteristics of the J57-P-1 two-spool turbojet engine with various inlet air-flow distortions was conducted in the altitude wind tunnel at the NACA Lewis laboratory. Along with a uniform inlet total-pressure distribution, one circumferential and three radial pressure distortions were investigated. Data were obtained over a complete range of compressor speeds both with and without intercompressor air bleed at a flight Mach number of 0.8 and at altitudes of 35,000 and 50,000 feet. Total-pressure distortions of the magnitudes investigated had very little effect on the steady-state operating line for either the outer or inner compressor. The small radial distortions investigated also had engine over that obtained with the uniform inlet pressure distribution. The circumferential distortion, however, raised the minimum speed at which the engine could operate without encountering surge when the intercompressor bleeds were closed. This increase in minimum speed resulted in a substantial reduction in the operable speed range accompanied by a reduction in the altitude operating limit.

Wallner, Lewis E.↗
NASA NTRS1961

Aerodynamic Characteristics, Temperature, and Noise Measurements of a Large-Scale External-Flow Jet-Augmented-Flap Model with Turbojet Engines Operating

An investigation has been conducted in the Langley full-scale tunnel on a large-scale model powered by turbojet engines with flattened rectangular nozzles. The wing had 35 deg. sweep of the leading edge, an aspect ratio of 6.5, a taper ratio of 0.31, and NACA 65(1)-412 and 65-408 airfoils at the root and tip. The investigation included measurements of the longitudinal aerodynamic characteristics of the model with half-span and full-span flaps and measurements of the sound pressure and skin temperature on the portions of the lower surface of the wing immersed in the jet flow. The tests were conducted over a range or angles of attack from -8 to 16 deg. for Reynolds numbers from 1.8 x 10(exp 6) to 4.4 x 10(exp 6) and a range of momentum coefficients from 0 to 2.0. In general, the aerodynamic results of this investigation made with a large-scale hot-jet model verified the results of previous investigations with small models powered by compressed-air jets. Although blowing was only done over the inboard portion of the wing, substantial amounts of induced lift were also obtained over the outboard portion of the wing. Skin temperatures were about 340 F and wing heating could be handled with available materials without cooling. Random acoustic loadings on the wing surface were high enough to indicate that fatigue failure from this source would require special consideration in the design of an external-flow jet flap system for an airplane.

Fink, Marvin P.↗
NASA NTRS1947

Preliminary Results of Altitude-Wind-Tunnel Investigation of X34C-4B Turbojet Engine. III - Compressor Performance: Compressor Performance - 3

The performance of the 11-stage axial-flow compressor in the X24C-4B turbojet engine was analyzed on the basis of results obtained from an investigation of the complete engine in the NACA Cleveland altitude wind tunnel. The engine was operated with four, exhaust nozzles of different outlet area over a range of engine speeds from 6000 to 12,500 rpm, corrected engine speeds from approximately 6100 to 13,600 rpm, and compressor Mach numbers from 0.45 to 1.00. Data are presented for engine operation over a range of simulated altitudes from 15,000 to 45,000 feet and simulated flight Mach numbers from 0.24 to 1.08.

Dupree, David T.↗
NASA NTRS1948

Performance of J33-A-23 Turbojet-Engine Compressor: Over-All Performance Characteristics of Compressor with 17-Blade Impeller - Part 1

The production-model 333-A-23 turbojet-engine compressor with a 17-blade impeller was operated at ambient and 0 F inlet temperatures and at inlet pressures of 14 and 5 inches mercury absolute for equivalent impeller speeds from 6000 to 12,750 rpm. The results of this investigation are compared with those of the 533-A-21 compressor. At the design equivalent speed of 11,750 rpm the maximum pressure ratio was 4.39. This occurred at the surge point at which the equivalent weight flow was 80.8 pounds per second, ana the adiabatic temperature-rise efficiency was 0.757. The maximum flow at the design equivalent speed was 88.0 pounds per second. The maximum adiabatic temperature-rise efficiency of 0.799 was obtained at an equivalent speed of 10,000 rpm, and equivalent weight flow of 62.9 pounds per second, and a pressure ratio of 3.20. At the maximum equivalent speed investigated (12,750 rpm), a peak pressure ratio of 4.90 was attained at an equivalent weight flow of 85.4 pounds per second and an efficiency of 0.680.

Beede, William L.↗
NASA NTRS1955

Over-All Performance of J65-B3 Turbojet Engine for Reynolds Number Indices from 0.8 to 0.2

The steady-state over-all performance characteristics of the J65-B3 turbojet engine were determined in an altitude test chamber for four exhaust-nozzle areas at Reynolds number indices of 0.8, 0.4, and 0.2. This range of Reynolds number indices corresponds to a range of altitudes from about sea level to 51,500 feet at a flight Mach number of 0.8. Generalized data are presented to allow calculation of engine performance at any flight condition corresponding to a Reynolds number index within the range investigated. Engine performance calculated from these generalized data is presented for seven altitudes over a range of flight speeds from zero to about 1100 knots. The use of an exhaust nozzle sized to give rated perforce at sea level would permit operation near the point of minimum specific fuel consumption for a wide range of flight conditions and engine speeds.

Fenn, David B.↗
NASA NTRS1948

Performance of Compressor of XJ-41-V Turbojet Engine: Performance Analysis Over Range of Compressor Speeds from 5000 to 10,000 RPM - 4

An investigation of the XJ-41-V turbojet-engine compressor was conducted to determine the performance of the compressor and to obtain fundamental information on the aerodynamic problems associated with large centrifugal-type compressors. The results of the research conducted on the original compressor indicated the compressor would not meet the desired engine-design air-flow requirements because of an air-flow restriction in the vaned collector. The compressor air-flow choking point occurred near the entrance to the vaned-collector passage and was instigated by a poor mass-flow distribution at the vane entrance and from relatively large negative angles of attack of the air stream along the entrance edges of the vanes at the outer passage wall and large positive angles of attack at the inner passage wall. As a result of the analysis, a design change of the vaned collector entrance is recommended for improving the maximum flow capacity of the compressor.

John W R Creagh↗