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At least 865 records · Page 48

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NASA NTRS1993

Large-scale density structures in the outer heliosphere

The Plasma Science experiment on the Voyager 2 spacecraft has measured the solar wind density from 1 to 38 AU. Over this distance, the solar wind density decreases as the inverse square of the heliocentric distance. However, there are large variations in this density at a given radius. Such changes in density are the dominant cause of changes in the solar wind ram pressure in the outer heliosphere and can cause large perturbations in the location of the termination shock of the solar wind. Following a simple model suggested by Suess, we study the non-equilibrium, dynamic location of the termination shock as it responds to these pressure changes. The results of this study suggest that the termination shock is rarely if ever at its equilibrium distance and may depart from that distance by as much as 50 AU at times.

Belcher, J. W.↗

NASA NTRS1992

Analysis of a space-exposed thermoplastic resin

The chemical characterization of a thermoplastic resin exposed to the low Earth orbit (LEO) environment for 10 months and for 5.8 years is reported. The resin, processed as a thick film and as a matrix for a graphite fiber reinforced composite, few exposed in the RAM direction on the NASA Long Duration Exposure Facility (LDEF). Differences attributable to environmental exposure were detected in infrared spectra and in various molecular weight parameters of film after 10 months in LEO. Those effects were not as apparent in composites after 5.8 years in LEO. Increased exposure to atomic oxygen toward the end of the LDEF mission probably scrubbed these effects from specimens exposed for 5.8 years. The intent of this study is to increase our fundamental understanding of space environmental effects on polymeric materials and to develop a benchmark for enhancing our methodology for and understanding of the ground-based simulation of space environmental effects.

Young, Philip R.↗

NASA NTRS1948

Ice Protection of Turbojet Engines by Inertia Separation of Water III : Annular Submerged Inlets

Aerodynamic and icing studies were conducted on a one-half-scale model of an annular submerged inlet for use with axial-flow turbojet engines. Pressure recoveries, screen radial-velocity profiles, circumferential mass-flow variations, and icing characteristics were determined at the compressor inlet. In order to be effective in maintaining water-free induction air, the inlet gap must be extremely small and ram-pressure recoveries consequently are low, the highest achieved being 65 percent at inlet-velocity ratio of 0.86. All inlets exhibited considerable screen icing. Severe mass-flow shifts occurred at angles of attack.

ICE PREVENTION AND REMOVAL - PROPULSION SYSTEMS↗

NASA NTRS1950

Altitude-chamber performance of British Rolls-Royce Nene II engine III : 18.00-inch-diameter jet nozzle

An altitude-chamber investigation of British Rolls-Royce Nene II turbojet engine was conducted over range of altitudes from sea level to 65,000 feet and ram pressure ratios from 1.10 to 3.50, using an 18.00-inch-diameter jet nozzle. The 18.00-inch-diameter jet nozzle gave slightly lower values of net-thrust specific fuel consumption than either the 18.41- or the standard 18.75-inch-diameter jet nozzles at high flight speeds. At low flight speeds, the 18.41-inch-diameter jet nozzle gave the lowest value of net-thrust specific fuel consumption.

ENGINES, TURBOJET↗

NASA NTRS1952

Performance and Operational Characteristics of a Python Turbine-propeller Engine at Simulated Altitude Conditions / Carl L. Meyer and Lavern A. Johnson

The performance and operational characteristics of a Python turbine propeller engine have been investigated over a range of engine operating conditions at simulated altitudes from 5000 to 40,000 feet in the NACA Lewis altitude wind tunnel . For the performance phase of the investigation, a single cowl-inlet ram pressure ratio was maintained and independent control of engine speed and fuel flow was used to obtain a range of power at each engine speed. Engine performance data obtained at a given altitude could not be used to predict performance accurately at other altitudes by use of the standard air pressure and temperature generalizing factors. Specific fuel consumption based on shaft horsepower decreased as the shaft horsepower increased at a given engine speed. At a given altitude condition and turbine -inlet total temperature, there was an optimum engine speed at which maximum shaft horsepower was obtained. At a given engine speed and turbine-inlet total temperature, a greater portion of the total available energy was converted to propulsive power as the altitude increased. Windmilling starts were made at altitudes up to 40,000 feet; the minimum true airspeed required for successful windmilling starts increased as the altitude was raised. In general, the engine control prevented large overshoots or undershoots of engine speed during starts, accelerations, and decelerations.

Engines - Armstrong-Siddely Python Turbo-Propeller↗

NASA NTRS1953

Analysis of several methods of pumping cooling air for turbojet engine afterburners

Several methods of pumping air to an annular cooling passage surrounding a typical axial-flow turbojet-engine afterburner were evaluated and compared on the basis of thrust and specific fuel consumption of the systems. Each system was analyzed over a range of afterburner-wall temperatures, flight Mach numbers, and exhaust-gas temperatures at sea level and 35,000 feet. Ram pressure recovery, boundary-layer pressure recovery, and the engine-jet actuated ejector appear to be satisfactory systems at high Mach numbers. Cooling with compressor-exit air bleed was found to be unsatisfactory,but the use of compressor-exit bleed air as the primary fluid in a high-performance ejector was satisfactory. The use of an auxiliary compressor driven from the engine shaft increased the thrust and decreased the specific fuel consumption of the engine for many of the conditions investigated.

PUMPS, JET AND THRUST AUGMENTORS↗

NASA NTRS1954

A flow calorimeter for determining combustion efficiency from residual enthalpy of exhaust gases

A flow calorimeter for determining the combustion efficiency of turbojet and ram-jet combustors from measurement of the residual enthalpy of combustion of the exhaust gas is described. Briefly, the calorimeter catalytically oxidizes the combustible constituents of exhaust-gas samples, and the resultant temperature rise is measured. This temperature rise is related to the residual enthalpy of combustion of the sample by previous calibration of the calorimeter. Combustion efficiency can be calculated from a knowledge of the residual enthalpy of the exhaust gas and the combustor input enthalpy. An accuracy of +-0.2 Btu per cubic foot was obtained with prepared fuel-air mixtures, and the combustion efficiencies of single turbojet combustors measured by both the flow-calorimeter and heat-balance methods compared within 3 percentage units. Flow calorimetry appears to be a suitable method for determining combustion efficiencies at high combustor temperatures where ordinary thermocouples cannot be used. The method is fundamentally more accurate than heat-balance methods at high combustion efficiencies and can be used to verify near-100-percent efficiency data.

Evans, Albert↗

NASA NTRS1954

Aspects of internal-flow-system design for helicopter propulsive units

Pertinent items relating to the design of internal-flow systems for reciprocating engine, turbine engine, and pressure-jet installations in helicopters are discussed. The importance of the following items is emphasized: controllable exit and ram recovery for reciprocating-engine cooling; performance penalties possible in turbine-engine installations attributed to distortions in the flow distribution at the face of the axial compressor; the effects of high subsonic flow velocities on the performance of duct elements; and the effects of centrifugal forces on the flow in a ducted helicopter rotor. Data on the maximum flow capacities of bends are included.

John R Henry↗

NASA NTRS1957

Some Operating Experience and Problems Encountered During Operation of a Free-jet Facility

During a free-jet investigation of a 28-inch ram-jet engine at a Mach number of 2.35, flow pulsation at the engine inlet were discovered which proved to have an effect on the engine performance and operational characteristics, particularly the engine rich blowout limits. This report discusses the finding of the flow pulsations, their elimination, and effect. Other facility characteristics, such as the establishment of flow simulation and the degree of subcritical operation of the diffuser, are also explained.

Mcaulay, John E↗

NASA NTRS1952

Evaluation of five conical center-body supersonic diffusers at several angles of attack

Five supersonic inlets having conical center bodies were investigated in a 16-inch ram jet engine with and without combustion at Mach numbers of 1.7, 1.9, and 2.0 and angles of attack from 0 degrees to 10 degrees. Four of the inlets were of the low-mass-flow-ratio type whereas the fifth inlet was of the high-mass-flow-ratio type. The stable operating range of the four low-nass-flow-ratio inlets decreased considerably as the angle of attack was increased despite an increase of the cone angle or the use of boundary-layer bleed. The stable range of the high-mass-flow-ratio inlet increased considerably both at 0 degrees and at greater angles of attack as the free-stream Mach number was decreased. Free-stream Mach number, angle of attack, and inlet type had negligible effects on combustion efficiency, temperature ratio, or combustion-chamber inlet Mach number in the supercritical region.

Englert, Gerald W↗

NASA NTRS1945

A method for the calculation of external lift, moment, and pressure drag of slender open-nose bodies of revolution at supersonic speeds

An approximate method is presented for the calculation of the external lift, moment, and pressure drag of slender open-nose bodies of revolution of supersonic speeds. The lift, moment, and pressure drag of a typical ram-jet body shape are calculated at Mach numbers 1.45, 1.60, 1.75, and 3.00; and the lift and moment results are compared with available experimental data. The agreement of the calculated lift and moment data with the experimental data is excellent. The pressure-drag comparison was not presented because of the uncertainty of the amount of skin-friction drag present in the experimental results.

Brown, Clinton E↗

NASA NTRS1957

Cloud-Droplet Ingestion in Engine Inlets with Inlet Velocity Ratios of 1.0 and 0.7

The paths of cloud droplets into two engine inlets have been calculated for a wide range of meteorological and flight conditions. The amount of water in droplet form ingested by the inlets and the amount and distribution of water impinging on the inlet walls are obtained from these droplet-trajectory calculations. In both types of inlet, a prolate ellipsoid of revolution represents either part or all of the forebody at the center of an annular inlet to an engine. The configurations can also represent a fuselage of an airplane with side ram-scoop inlets. The studies were made at an angle of attack of 0 degree. The principal difference between the two inlets studied is that the inlet-air velocity of one is 0.7 that of the other. The studies of the two velocity ratios lead to some important general concepts of water ingestion in inlets.

Brun, Rinaldo J↗

NASA NTRS1947

Performance Charts for a Turbojet System

Convenient charts are presented for computing the thrust, fuel consumption, and other performance values of a turbojet system. These charts take into account the effects of ram pressure, compressor pressure ratio, ratio of combustion-chamber-outlet temperature to atmospheric temperature, compressor efficiency, turbine efficiency, combustion efficiency, discharge-nozzle coefficient, losses in total pressure in the inlet to the jet-propulsion unit and in the combustion chamber, and variation in specific heats with temperature. The principal performance charts show clearly the effects of the primary variables and correction charts provide the effects of the secondary variables. The performance of illustrative cases of turbojet systems is given. It is shown that maximum thrust per unit mass rate of air flow occurs at a lower compressor pressure ratio than minimum specific fuel consumption. The thrust per unit mass rate of air flow increases as the combustion-chamber discharge temperature increases. For minimum specific fuel consumption, however, an optimum combustion-chamber discharge temperature exists, which in some cases may be less than the limiting temperature imposed by the strength temperature characteristics of present materials.

CHARTS-PERFORMANCE-ENGINES, TURBOJET↗

NASA NTRS1948

Altitude-wind-tunnel investigation of a 4000-pound-thrust axial-flow turbojet engine VI : combustion-chamber performance

An analysis of the performance of the types A, B, and C combustion chambers of the 4000-pound-thrust axial-flow turbojet engine is presented. The data were obtained from investigations of the complete engine over a range of pressure altitudes from 5000 to 40,000 feet and ram pressure ratios from 1.00 to 1.86. The combustion-chamber pressure losses, the effect of the losses on cycle efficiency, and the combustion efficiency are discussed.

Pinkel, I Irving↗

NASA NTRS1949

Spark ignition of flowing gases I : energies to ignite propane-air mixtures in pressure range of 2 to 4 inches mercury absolute

Ignition studies of flowing gases were made to obtain information applicable to ignition problems in gas-turbine and ram-jet aircraft propulsion systems operating at altitude conditions.Spark energies required for ignition of a flowing propane-air mixture were determined for pressure of 2 to 4 inches mercury absolute, gas velocities of 5.0 to 54.2 feet per second, fuel-air ratios of 0.0607 to 0.1245, and spark durations of 1.5 to 24,400 microseconds. The results showed that at a pressure of 3 inches mercury absolute the minimum energy required for ignition occurred at fuel-air ratios of 0.08 to 0.095. The energy required for ignition increased almost linearly with increasing gas velocity. Shortening the spark duration from approximately 25,000 to 125 microseconds decreased the amount of energy required for ignition. A spark produced by the discharge of a condenser directly into the spark gap and having a duration of 1.5 microseconds required ignition energies larger than most of the long-duration sparks.

Swett, Clyde C , Jr↗

NASA NTRS1953

Air admixture to exhaust jets

The problem of thrust increase by air admixture to exhaust jets of rockets, turbojet, ram- and pulse-jet engines is investigated theoretically. The optimum ratio of mixing chamber pressure to ambient pressure and speed range for thrust increase due to air admixture is determined for each type of jet engine.

ENGINES, PULSE-JET↗

NASA NTRS1952

Pressure recovery, drag, and subcritical stability characteristics of conical supersonic diffusers with boundary-layer removal

A study of two 20 degrees half-angle, low mass-flow ratio conical supersonic inlets with cone boundary-layer bleed was made on a 16-inch ram-jet engine in the Lewis 8- by 6-foot supersonic wind tunnel. A greater stable subcritical range of operation was obtained with the bleed inlets than with the corresponding inlet without boundary-layer bleed. The drag added by the bleed system was small.

Obey, Leonard T↗

NASA NTRS1990

SEU hardening of CMOS memory circuit

This paper reports a design technique to harden CMOS memory circuits against Single Event Upset (SEU) in the space environment. A RAM cell and Flip Flop design are presented to demonstrate the method. The Flip Flop was used in the control circuitry for a Reed Solomon encoder designed for the Space Station.

Whitaker, S.↗