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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 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS1975

Comparison of parametric duct-burning turbofan and non-afterburning turbojet engines in a Mach 2.7 transport

A parametric study was made of duct-burning turbofan and suppressed dry turbojet engines installed in a supersonic transport. A range of fan pressure ratios was considered for the separate-flow-fan engines. The turbofan engines were studied both with and without jet noise suppressors. Single- as well as dual-stream suppression was considered. Attention was concentrated on designs yielding sideline noises of FAR 36 (108 EPNdB) and below. Trades were made between thrust and wing area for a constant takeoff field length. The turbofans produced lower airplane gross weights than the turbojets at FAR 36 and below. The advantage for the turbofans increased as the sideline noise limit was reduced. Jet noise suppression, especially for the duct stream, was very beneficial for the turbofan engines as long as duct burning was permitted during takeoff. The maximum dry unsuppressed takeoff mode, however, yielded better results at extremely low noise levels. Noise levels as low as FAR 36-11 EPNdB were obtained with a turbofan in this takeoff mode, but at a considerable gross weight penalty relative to the best FAR 36 results.

Whitlow, J. B., Jr.↗
NASA NTRS1972

Modularized instrument system for turbojet engine test facilities

A modular instrument system is being developed to handle the many data channels 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; in addition, each module provides a controlled environment for its contents. A minicomputer in the control room gathers the data, performs 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 minimization of manual adjustments; and (4) reduced cost through minimization of wiring and simplification of control room display.

Nieberding, W. C.↗
NASA NTRS1956

Effect of Inlet-air-flow Distortions on Steady-state Performance of J65-B-3 Turbojet Engine

The effects of inlet-air-flow distortions on the performance of the J65-B-3 turbojet engine were determined over a range of altitudes from 15,000 to 50,000 feet at a flight Mach number of 0.8. Radial inlet-air-flow distortions apparently do not affect the radial distribution of pressure after the first few compressor stages, while the circumferential inlet-air-flow distortion carried completely through the engine. For the distortions investigated, at rated exhaust-gas temperature and fixed-area exhaust-nozzle operation, the primary effect of the radial inlet-air-flow distortions was to reduce the engine air flow, and the primary effect of the circumferential distortion was to impose a temperature profile on the turbine, both resulting in reduction of thrust.

Smith, Ivan D↗
NASA NTRS1949

Performance of J33 turbojet engine with shaft-power extraction III : turbine performance

The performance of the turbine component of a J33 turbojet engine was determined over a range of turbine speeds from 8000 to 11,500 rpm.Turbine-inlet temperature was varied from the minimum required to drive the compressor to a maximum of approximately 2000 degrees R at each of several intermediate turbine speeds. Data are presented that show the horsepower developed by the turbine per pound of gas flow. The relation between turbine-inlet stagnation pressure, turbine-outlet stagnation pressure, and turbine-outlet static pressure was established. The turbine-weight-flow parameter varied from 39.2 to 43.6. The maximum turbine efficiency measured was 0.86 at a pressure ratio of 3.5 and a ratio of blade speed to theoretical nozzle velocity of 0.39. A generalized performance map of the turbine-horsepower parameter plotted against the turbine-speed parameter indicated that the best turbine efficiency is obtained when the turbine power is 10 percent greater than the compressor horsepower. The variation of efficiency with the ratio of blade speed to nozzle velocity indicated that the turbine operates at a speed above that for maximum efficiency when the engine is operated normally with the 19-inch-diameter jet nozzle.

Huppert, M C↗
NASA NTRS1975

Experience with integrally-cast compressor and turbine components for a small, low-cost, expendable-type turbojet engine

Experiences with integrally-cast compressor and turbine components during fabrication and testing of four engine assemblies of a small (29 cm (11 1/2 in.) maximum diameter) experimental turbojet engine design for an expendable application are discussed. Various operations such as metal removal, welding, and re-shaping of these components were performed in preparation of full-scale engine tests. Engines with these components were operated for a total of 157 hours at engine speeds as high as 38,000 rpm and at turbine inlet temperatures as high as 1256 K (1800 F).

Dengler, R. P.↗
NASA NTRS1951

Altitude Performance of J35-A-17 Turbojet Engine in an Altitude Chamber

An investigation of the altitude performance characteristics of an Allison J35-A-17 turbojet engines have been conducted in an altitude chamber at the NACA Lewis laboratory. Engine performance was obtained over a range of altitudes from 20,000 to 60,000 feet at a flight Mach number of 0.62 and a range of flight Mach numbers from 0.42 to 1.22 at an altitude of 30,000 feet. The performance of the engine over the range investigated could be generalized up to an altitude of 30,000 feet. Performance of the engine at any flight Mach number in the range investigated can be predicted for those operating condition a t which critical flow exits in the exhaust nozzle with the exception of the variables corrected net thrust, and net-thrust specific fuel consumption.

Vincent, K. R.↗
NASA NTRS1958

Altitude Performance of the Afterburner on the Iroquois Turbojet Engine. Coord. No. AF-P-6

The performance and operational characteristics of two afterburner configurations for the Iroquois turbojet engine were evaluated in an altitude test chamber over a range of afterburner equivalence ratios at afterburner-inlet pressures from 733 to 3186 pounds per square foot absolute. These conditions correspond to an altitude range from 38,700 to 66,800 feet at a flight Mach number of 1.5. The only difference between the two afterburner configurations was in the pattern of afterburner fuel injection. At an afterburner-inlet pressure of approximately 3100 pounds per square foot absolute, corresponding to an altitude of 38,700 feet and a_ flight Mach number of 1.5, the combustion efficiency of both configurations reached peak values of 0.80 to 0.85 at equivalence ratios of 0.35 to 0.40. However, further reduction in the afterburner-inlet pressure severely affected combustion efficiency. For example, at an afterburner inlet pressure level of 700 to 1000 pounds per square foot absolute, the efficiency for both configurations was 0.20 to 0.40.

Groesbeck, Donald E.↗
NASA NTRS1977

ENGINEL: A single rotor turbojet engine cycle match performance program

ENGINEL is a computer program which was developed to generate the design and off-design performance of a single rotor turbojet engine with or without afterburning using a cycle match procedure. It is capable of producing engine performance over a wide range of altitudes and Mach numbers. The flexibility, of operating with a variable geometry turbine, for improved off-design fuel consumption or with a fixed geometry turbine as in conventional turbojets, has been incorporated. In addition, the option of generation engine performance with JP4, liquid hydrogen or methane as fuel is provided.

Lovell, W. A.↗
NASA NTRS1975

Experience with integrally-cast compressor and turbine components for a small, low-cost, expendable-type turbojet engine

A discussion regarding experiences with integrally-cast compressor and turbine components during fabrication and testing of four engine assemblies of a small (29 cm (11-1/2 in.) maximum diameter) experimental turbojet engine design for an expendable application is presented. Various operations such as metal removal, welding, and re-shaping of these components are performed in preparation of full-scale engine tests. Engines with these components have been operated for a total of 157 hours at engine speeds as high as 38,000 rpm and at turbine inlet temperatures as high as 1256 K (1800 F).

Dengler, R. P.↗
NASA NTRS1979

Effects of steady-state pressure distortion on the stall margin of a J85-21 turbojet engine

The effects of the inlet pressure distortions, induced by five screen patterns, on the performance of a J85-21 turbojet engine was conducted at the NASA Lewis Research Center. Testing was in support of the HiMAT RPRV program at Dryden Flight Research Center. Distortion patterns were chosen based on anticipated application of test results of the HiMAT installation. Tests were conducted at a simulated Mach number and altitude condition of 0.9 and 10 973 meters. Results are presented in terms of distortion levels and standard compressor performance parameters.

Bobula, G. A.↗
NASA NTRS1947

Investigation of Thrust Augmentation of a 1600-pound Thrust Centrifugal-flow-type Turbojet Engine by Injection of Refrigerants at Compressor Inlets

Investigations were conducted to determine effectiveness of refrigerants in increasing thrust of turbojet engines. Mixtures of water an alcohol were injected for a range of total flows up to 2.2 lb/sec. Kerosene was injected into inlets covering a range of injected flows up to approximately 30% of normal engine fuel flow. Injection of 2.0 lb/sec of water alone produced an increase in thrust of 35.8% of rate engine conditions and kerosene produced a negligible increase in thrust. Carbon dioxide increased thrust 23.5 percent.

Jones, William L.↗
NASA NTRS1955

Temperature in a J47-25 Turbojet-engine Combustor and Turbine Sections During Steady-state and Transient Operation in a Sea-level Test Stand

In order to determine the conditions of engine operation causing the most severe thermal stresses in the hot parts of a turbojet engine, a J47-25 engine was instrumented with thermocouples and operated to obtain engine material temperatures under steady-state and transient conditions. Temperatures measured during rated take-off conditions of nozzle guide vanes downstream of a single combustor differed on the order of 400 degrees F depending on the relation of the blades position to the highest temperature zone of the burner. Under the same operation conditions, measured midspan temperatures in a nozzle guide vane in the highest temperature zone of a combustor wake ranged from approximately 1670 degrees F at leading and trailing edges to 1340 degrees F at midchord on the convex side of the blade. The maximum measured nozzle-guide-vane temperature of 1920degrees at the trailing edge occurred during a rapid acceleration from idle to rated take-off speed following which the tail-pipe gas temperature exceeded maximum allowable temperature by 125 degrees F.

Morse, C R↗
NASA NTRS1948

Altitude-Wind-Tunnel Investigation of a 3000-Pound-Thrust Axial-Flow Turbojet Engine: Operational Characteristics - 4

An investigation was conducted 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. 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 were evaluated.

Hawkins, W. Kent↗
NASA NTRS1948

Altitude-Wind-Tunnel Investigation of a 4000-Pound-Thrust Axial-Flow Turbojet Engine: Analysis of Turbine Performance - 5

Performance characteristics of the turbine of a 4000-pound-thrust axial-flow turbojet engine was determined in investigations of the complete engine in the NACA Cleveland altitude wind tunnel. Characteristics are presented as functions of the total-pressure ratio across the turbine and of turbine speed and gas flow corrected to sea-level conditions. Three turbine nozzles of different areas were used to determine the area that gave optimum performance. Inasmuch as tail-pipe nozzles of different diameters were investigated in combination with the standard turbine nozzle, the effect of varying discharge conditions on turbine operation could be observed. The investigations covered a range of pressure attitudes from 5000 to 40,000 feet. The engine was investigated over the entire operable range of speeds at each altitude. At pressure altitude of 30,000 feet, the effect on turbine operation of varying the ram pressure ration over a range from 1.10 to 1.77 was evaluated. An altitude effect was apparent when turbine pressure ratio was plotted against corrected turbine speed but it was so slight as to be negligible insofar as the turbine efficiencies were concerned. A maximum turbine efficiency of slightly more than 82 percent was obtained with the configuration using the standard turbine nozzle and the low-flow compressor. This efficiency, which is somewhat lower than the actual turbine efficiency, is uncorrected for accessories drive power, bearing friction, tail-pipe pressure drop, compressor thermal radiation, and introduction of turbine-disk cooling air into the gas stream. Changes in the ram pressure ratio had a negligible effect on the turbine efficiency.

Krebs, Richard P.↗
NASA NTRS1954

Preliminary Altitude Performance Characteristics of the J57-P-1 Turbojet Engine with Fixed-area Exhaust Nozzle

An investigation to determine the altitude performance of the J57-P-1 turbojet engine and components was conducted at the NACA Lewis altitude wind tunnel. Data were obtained over a corrected inboard rotor speed range from 56 to 106 percent of rated speed, with intercompressor bleeds both open and closed, at altitudes from 15,000 to 50,000 feet and at a flight Mach number of 0.81. The corresponding range of Reynolds number indices was from 0.858 to 0.213. All data presented were obtained with a fixed-area exhaust nozzle sized according to the manufacturer's specification. Over-all engine performance parameters are presented as functions of inboard rotor speed corrected on the basis of engine inlet temperature. Component parameters are presented as functions of their respective corrected rotor speeds. A tabulation of all performance data is included in addition to the graphical presentation. Corrected net thrust is unusually sensitive to changes in corrected inboard rotor speed in the high speed region. A change of 1 percent in speed, at sated speed, produced a change of 6 percent in corrected net thrust . At rated engine speed, increasing the altitude from 15,000 to 50,000 feet at a constant flight Mach number of 0.81 increased the specific fuel consumption 13 percent but did not affect corrected net thrust.

Bloomer, Harry E.↗
NASA NTRS1950

Preliminary Performance Data on Westinghouse Electronic Power Regulator Operating on J34-WE-32 Turbojet Engine in Altitude Wind Tunnel

The behavior of the Westinghouse electronic power regulator operating on a J34-WE-32 turbojet engine was investigated in the NACA Lewis altitude wind tunnel at the request of the Bureau of Aeronautics, Department of the Navy. The object of the program was to determine the, steady-state stability and transient characteristics of the engine under control at various altitudes and ram pressure ratios, without afterburning. Recordings of the response of the following parameters to step changes in power lever position throughout the available operating range of the engine were obtained; ram pressure ratio, compressor-discharge pressure, exhaust-nozzle area, engine speed, turbine-outlet temperature, fuel-valve position, jet thrust, air flow, turbine-discharge pressure, fuel flow, throttle position, and boost-pump pressure. Representative preliminary data showing the actual time response of these variables are presented. These data are presented in the form of reproductions of oscillographic traces.

Ketchum, James R.↗