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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 469 records · Page 26
All sourcesNASA NTRSDOE OSTI
NASA NTRS1953

Force and pressure-recovery characteristics at supersonic speeds of a conical nose inlet with bypasses discharging outward from the body axis

Aerodynamic and performance characteristics of a conical spike nacelle-type inlet with two bypasses are presented at Mach numbers of 1.6, 1.8, and 2.0 for angles of attach up to 90 degrees. The bypasses were located 6 inlet diameters downstream of the inlet and were designed to discharge the bypass mass flow outward from the body axis. The inlet was designed to attain a mass-flow ratio of unity at a Mach number of 2.0. It is shown that discharging the bypass mass flow outward from the body nearly doubles the critical drag of a similar configuration but with bypass discharge in an axial direction. As a result of this greater drag, the net force on the model in the flight direction is reduced when comparison is made with the axial discharge case. The lift and pitching-moment coefficients are slightly higher than those for a configuration without bypasses. Approximately 25 % of the maximum inlet mass flow was discharged through the bypasses, and the pressure-recovery and mass-flow characteristics were in qualitative and quantitative agreement with the results of an investigation of a similar configuration with axial discharge.

DIFFUSERS, SUBSONIC↗
NASA NTRS1997

Bleed Hole Flow Phenomena Studied

Boundary-layer bleed is an invaluable tool for controlling the airflow in supersonic aircraft engine inlets. Incoming air is decelerated to subsonic speeds prior to entering the compressor via a series of oblique shocks. The low momentum flow in the boundary layer interacts with these shocks, growing in thickness and, under some conditions, leading to flow separation. To remedy this, bleed holes are strategically located to remove mass from the boundary layer, reducing its thickness and helping to maintain uniform flow to the compressor. The bleed requirements for any inlet design are unique and must be validated by extensive wind tunnel testing to optimize performance and efficiency. To accelerate this process and reduce cost, researchers at the NASA Lewis Research Center initiated an experimental program to study the flow phenomena associated with bleed holes. Knowledge of these flow properties will be incorporated into computational fluid dynamics (CFD) models that will aid engine inlet designers in optimizing bleed configurations before any hardware is fabricated. This ongoing investigation is currently examining two hole geometries, 90 and 20 (both with 5-mm diameters), and various flow features.

Source record↗
NASA NTRS1951

Theoretical and experimental investigation of additive drag

The significance of additive drag is discussed and equations for determining its approximate value are derived. Charts are presented giving values of additive drag for open-nose inlets and for annular-nose inlets with conical flow at the inlet. The effects of variable inlet total-pressure recovery and static pressures on the center body are investigated, and an analytical method of predicting the variation of pressure on the center body with mass-flow ratio is given. Experimental values of additive drag are compared with values predicted by the methods.

AIR INLETS↗
NASA NTRS1948

Simulated Altitude Performance of Combustors for the Westinghouse 24C Jet Engine I-24C-2 Combustor

A Westinghouse 24C-2 combustor was investigated at conditions simulating operation of the 24C Jet engine at zero ram over ranges of altitude and engine speed. The investigation was conducted to determine the altitude operational limits, that is, the maximum altitude for various engine speeds at which an average combustor-outlet gas temperature sufficient for operation of the jet engine could be obtained. Information was also obtained regarding the character of the flames, the combustion efficiency, the combustor-outlet gas temperature and velocity distributions, the extent of afterburning, the flow characteristics of the fuel manifolds, the combustor inlet-to-outlet total-pressure drop, and the durability of the combustor basket. The results of the investigation indicated that the altitude operational limits for zero ram decreased from 12,000 feet at an engine speed of 4000 rpm to a minimum of 9000 feet at 6000 rpm, and thence increased to 49,000 feet at 12,000 rpm.. At altitudes below the operational limits, flames were essentially steady, but, at altitudes above the operational limits, flames were often cycling and either blew out or caused violent explosions and vibrations. At conditions on the altitude operational limits the type of combustion varied from steady to cycling with increasing fuel-air ratio and the reverse occurred with decreasing fuel-air ratio. In the region of operation investigated, the combustion efficiency ranged from 75 to 95 percent at altitudes below the operational limits and dropped to 55 percent or less at some altitudes above the operational limits. The deviations in the local combustor-outlet gas temperatures were within +20 to -30 percent of the mean combustor temperature rise for inlet-air temperatures at the low end of the range investigated, but became more uneven (up to +/-100 percent) with increasing inlet-air temperatures. The distribution of the combustor-outlet gas velocity followed a similar trend. Practically no afterburning downstream of the combustor outlet occurred. At conditions of high inlet-air temperature several factors indicated that fuel vapor or air formed in the fuel manifolds and adversely affected combustion. The combustor inlet-to-outlet total-pressure drop can be correlated as a function of the ratio of the combustion-air inlet density to outlet density and of the inlet dynamic pressure. The walls of the combustor basket were warped and burned during 29 hours of operation.

Manganiello, Eugene J.↗
NASA NTRS1953

Investigation of Pressure Recovery of a Single-conical-shock Nose Inlet at Mach Number 5.4

An experimental investigation of the performance of a single-conical-shock diffuser was conducted at a Mach number of 5.4 and a Reynolds number based on model diameter of 375,000. Total-pressure recoveries of 13.7 and 13.1 percent were obtained at angles of attack of 0 degrees and 3 degrees, respectively. The corresponding kinetic energy efficiencies were 86.4 percent at an angle of attack of 0 degrees and 86.0 percent at an angle of attack of 3 degrees.

DIFFUSERS, SUPERSONIC↗
NASA NTRS1999

NASA Glenn 1-by 1-Foot Supersonic Wind Tunnel User Manual

This manual describes the NASA Glenn Research Center's 1 - by 1 -Foot Supersonic Wind Tunnel and provides information for customers who wish to conduct experiments in this facility. Tunnel performance envelopes of total pressure, total temperature, and dynamic pressure as a function of test section Mach number are presented. For each Mach number, maps are presented of Reynolds number per foot as a function of the total air temperature at the test section inlet for constant total air pressure at the inlet. General support systems-such as the service air, combustion air, altitude exhaust system, auxiliary bleed system, model hydraulic system, schlieren system, model pressure-sensitive paint, and laser sheet system are discussed. In addition, instrumentation and data processing, acquisition systems are described, pretest meeting formats and schedules are outlined, and customer responsibilities and personnel safety are addressed.

Seablom, Kirk D.↗
NASA NTRS1952

Experimental Investigation of a Two-dimensional Split-wing Ram-jet Inlet at Mach Number of 3.85

Performance characteristics of a two-dimensional isentropic diffuser have been experimentally determined at a Mach number of 3.85. At zero angle of attack, a maximum total-pressure recovery of 0.41 was obtained with a supercritical mass-flow ratio of 0.95. As a consequence of the twin-duct arrangement of the diffuser, a large discontinuity in pressure recovery and mass flow with a characteristic hysteresis was encountered between critical and subcritical operation. An asymmetric shock pattern with large-scale separation and flow reversal in one of the passages occurred at reduced mass flows. Pressure and force data presented for an angle-of-attack range from zero to 4 degrees.

DIFFUSERS, SUPERSONIC↗
NASA NTRS1951

Preliminary Investigation of a New Type of Supersonic Inlet

A supersonic inlet with supersonic deceleration of the flow entirely outside of the inlet is considered. A particular arrangement with fixed geometry having a central body with a circular annular intake is analyzed, and it is shown theoretically that this arrangement gives high pressure recovery for a large range of Mach number and mass flow and therefore is practical for use on supersonic airplanes and missiles. For some Mach numbers the drag coefficient for this type of inlet is larger than the drag coefficient for the type of inlet with supersonic compression entirely inside, but the pressure recovery is larger for all flight conditions. The differences in drag can be eliminated for the design Mach number. Experimental results confirm the results of the theoretical analysis and show that pressure recoveries of 95 percent for Mach numbers of 1.33 and 1.52, 92 percent for a Mach number of 1.72, and 86 percent for a Mach number of 2.10 are possible, with the configurations considered. If the mass flow decreases, the total drag coefficient increases gradually and the pressure recovery does not change appreciably. The results of this work were first presented in a classified document issued in 1946.

DIFFUSERS, SUPERSONIC↗
NASA NTRS1977

Static aerodynamic characteristics of a 0.035-scale model of a modified NKC-135 airplane at a Mach number of 0.28

A 0.035-scale model fo a modified NKC-135 airplane was tested in 12-foot pressure wind tunnel to determine the effects on the static aerodynamic characteristics of modifications to the basic aircraft. Modifications investigated included: nose, lower fuselage, and upper fuselage radomes; wing pylons and pods; overwing probe; and air conditioning inlets. The investigation was performed at a Mach number of 0.28 over a Reynolds number range from 6.6 to 26.2 million per meter. Angles of attack and sideslip varied from -8 deg to 20 deg and from -18 deg to 8 deg, respectively, for various combinations of flap, aileron, and rudder deflections. A limited analysis of the test results indicates that the addition of the radomes reduces lateral-directional stability and control effectiveness of the basic aircraft.

Hedstrom, E.↗
NASA NTRS1980

Experimental evaluation of catalytic combustion with heat removal at near stoichiometric conditions

Two concentric tube configurations were tested. Tests were conducted at an inlet pressure of 150,000 Pa, inlet fuel air mixture temperatures from 780 to 960 K, combustion air flow rates from 0.78 to 1.5 g/s, equivalence ratios up to 0.90, and a range of cooling air flow rates. Propane and propylene fuels were used. Both configurations used air flowing through the center tube for cooling and combustion in the annulus on the catalytic surface. One configuration had the catalyst applied to the outside surface of the inner tube. Conversion of the fuel was very low for this configuration. The other configuration had the catalyst applied to the inside surface of the outer tube. Conversion of the fuel was considerably better in this configuration.

Bulzan, D. L.↗
NASA NTRS1984

Aerodynamic characteristics of the 40- by 80/80- by 120-foot wind tunnel at NASA Ames Research Center

The design and testing of vane sets and air-exchange inlet for the 40 x 80/80 x 120-ft wind tunnel at NASA Ames are reported. Boundary-layer analysis and 2D and 3D inviscid panel codes are employed in computer models of the system, and a 1/10-scale 2D facility and a 1/50-scale 3D model of the entire wind tunnel are used in experimental testing of the vane sets. The results are presented in graphs, photographs, drawings, and diagrams are discussed. Generally good agreement is found between the predicted and measured performance.

Corsiglia, V. R.↗
NASA NTRS1984

Aerodynamic characteristics of the 40- by 80-/80- by 120-ft wind tunnel at NASA-Ames Research Center

The design and testing of vane sets and air-exchange inlet for the 40 x 80/80 x 120-ft wind tunnel at NASA Ames are reported. Boundary-layer analysis and 2D and 3D inviscid panel codes are employed in computer models of the system, and a 1/10-scale 2D facility and a 1/50-scale 3D model of the entire wind tunnel are used in experimental testing of the vane sets. The results are presented in graphs, photographs, drawings, and diagrams and discussed. Generally good agreement is found between the predicted and measured performance.

Corsiglia, V. R.↗
NASA NTRS1950

Characteristics of Perforated Diffusers at Free-stream Mach Number 1.90

An investigation was conducted at Mach number 1.90 to determine pressure recovery and mass-flow characteristics of series of perforated convergent-divergent supersonic diffusers. Pressure recoveries as high as 96 percent were obtained, but at reduced mass flows through the diffuser. Theoretical considerations of effect of perforation distribution on shock stability in converging section of diffuser are presented and correlated with experimental data. A method of estimating relative importance of pressure recovery and mass flow on internal thrust coefficient basis is given and a comparison of various diffusers investigated is made.

FLOW, SUPERSONIC↗
NASA NTRS1953

Free-flight Performance of a Rocket-boosted, Air-launched 16-inch-diameter Ram-jet Engine at Mach Numbers up to 2.20

The investigation of air-launched ram-jet engines has been extended to include a study of models with a nominal design free-stream Mach number of 2.40. These models require auxiliary thrust in order to attain a flight speed at which the ram jet becomes self-accelerating. A rocket-boosting technique for providing this auxiliary thrust is described and time histories of two rocket-boosted ram-jet flights are presented. In one flight, the model attained a maximum Mach number of 2.20 before a fuel system failure resulted in the destruction of the engine. Performance data for this model are presented in terms of thrust and drag coefficients, diffuser pressure recovery, mass-flow ratio, combustion efficiency, specific fuel consumption, and over-all engine efficiency.

AIR INLETS - NOSE, ANNULAR↗
NASA NTRS1993

Method of reducing drag in aerodynamic systems

In the present method, boundary layer thickening is combined with laminar flow control to reduce drag. An aerodynamic body is accelerated enabling a ram turbine on the body to receive air at velocity V sub 0. The discharge air is directed over an aft portion of the aerodynamic body producing boundary layer thickening. The ram turbine also drives a compressor by applying torque to a shaft connected between the ram turbine and the compressor. The compressor sucks in lower boundary layer air through inlets in the shell of the aircraft producing laminar flow control and reducing drag. The discharge from the compressor is expanded in a nozzle to produce thrust.

Hrach, Frank J.↗