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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 595 records · Page 33
All sourcesNASA NTRSDOE OSTI
NASA NTRS1988

A preliminary design study of supersonic through-flow fan inlets

From Mach 3.20 cruise propulsion systems, preliminary design studies for two supersonic through-flow fan primary inlets and a single core inlet were undertaken. Method of characteristics and one dimensional performance techniques were applied to assess the potential improvements supersonic through-flow fan technology has over more conventional systems. A fixed geometry supersonic through-flow fan primary inlet was found to have better performance than a conventional inlet design on the basis of total pressure recovery, air flow, aerodynamic drag and size and weight.

Barnhart, Paul J.↗
NASA NTRS1988

CFD application to supersonic/hypersonic inlet airframe integration

Supersonic external compression inlets are introduced, and the computational fluid dynamics (CFD) codes and tests needed to study flow associated with these inlets are outlined. Normal shock wave turbulent boundary layer interaction is discussed. Boundary layer control is considered. Glancing sidewall shock interaction is treated. The CFD validation of hypersonic inlet configurations is explained. Scramjet inlet modules are shown.

Benson, Thomas J.↗
NASA NTRS1989

Three dimensional viscous analysis of a hypersonic inlet

The flow fields in supersonic/hypersonic inlets are currently being studied at NASA Lewis Research Center using 2- and 3-D full Navier-Stokes and Parabolized Navier-Stokes solvers. These tools have been used to analyze the flow through the McDonnell Douglas Option 2 inlet which has been tested at Calspan in support of the National Aerospace Plane Program. Comparisons between the computational and experimental results are presented. These comparisons lead to better overall understanding of the complex flows present in this class of inlets. The aspects of the flow field emphasized in this work are the 3-D effects, the transition from laminar to turbulent flow, and the strong nonuniformities generated within the inlet.

Reddy, D. R.↗
NASA NTRS1989

Three dimensional viscous analysis of a hypersonic inlet

The flow fields in supersonic/hypersonic inlets are currently being studied at NASA Lewis Research Center using 2- and 3-d full Navier-Stokes and parabolized Navier-Stokes solvers. These tools have been used to analyze the flow through the McDonnell Douglas Option 2 inlet which has been tested at Calspan in support of the National Aerospace Plane Program. Comparisons between the computational and experimental results are presented. These comparisons lead to better overall understanding of the complex flows present in this class of inlets. The aspects of the flow field emphasized in this work are the 3-D effects, the transition from laminar to turbulent flow, and the strong nonuniformities generated within the inlet.

Reddy, D. R.↗
NASA NTRS1988

Patched-grid computations of high-speed inlet flows

Two implicit algorithms for the compressible Navier-Stokes equations are applied to the computation of high-speed inlet flows typical of those which are encountered on propulsion systems of advanced hypersonic vehicles. A patched grid formulation is used to treat accurately the blunt cowl interaction with the forebody compression field and efficiently treat variations in inlet geometry downstream of the inlet entrance. Results from the two algorithms, each based on an upwind-biased spatial differencing approach for the convective and pressure terms within a flux-difference-splitting framework, are compared with experimental results for both inlets, and reasonable agreement is found.

Thomas, J. L.↗
NASA NTRS1951

Theoretical Investigation of Submerged Inlets at Low Speeds

The general characteristics of the flow field in a submerged air inlet are investigated by theoretical, wind-tunnel, and visual-flow studies. Equations are developed for calculating the laminar and turbulent boundary-layer growth along the ramp floor for parallel, divergent, and convergent ramp walls, and a general equation is derived relating the boundary-layer pressure losses to the boundary-layer thickness. It is demonstrated that the growth of the boundary layer on the floor of the divergent-ramp inlet is retarded and that a vortex pair is generated in such an inlet. Functional relationships are established between the pressure losses in the vortices and the geometry of the inlet. A general discussion of the boundary layer and vortex formations is included, in which variations of the various losses and of the incremental external drag with mass-flow ratio are considered. Effects of compressibility are also discussed.

Sacks, Alvin H.↗
NASA NTRS1990

Comparison of 3-D viscous flow computations of Mach 5 inlet with experimental data

A time marching 3-D full Navier-Stokes code, called PARC3D, is validated for an experimental Mach 5 inlet configuration using the data obtained in the 10 x 10 ft supersonic wind tunnel at the NASA Lewis Research Center. For the first time, a solution is obtained for this configuration with the actual geometry, the tunnel conditions, and all the bleed zones modeled in the computation. Pitot pressure profiles and static pressures at various locations in the inlet are compared with the corresponding experimental data. The effect of bleed zones, located in different places on the inlet walls, in eliminating the low energy vortical flow generated from the 3-D shock-boundary layer interaction is simulated very well even though some approximations are used in applying the bleed boundary conditions and in the turbulence model. A further detailed study of the effect of individual bleed ports is needed to understand fully the actual mechanism of efficiently eliminating the vortical flow from the inlet. A better turbulence model would help to improve the accuracy even further in predicting the corner flow boundary layer profiles.

Reddy, D. R.↗
NASA NTRS1991

Vortex generator design for aircraft inlet distortion as a numerical optimization problem

Aerodynamic compatibility of aircraft/inlet/engine systems is a difficult design problem for aircraft that must operate in many different flight regimes. Takeoff, subsonic cruise, supersonic cruise, transonic maneuvering, and high altitude loiter each place different constraints on inlet design. Vortex generators, small wing like sections mounted on the inside surfaces of the inlet duct, are used to control flow separation and engine face distortion. The design of vortex generator installations in an inlet is defined as a problem addressable by numerical optimization techniques. A performance parameter is suggested to account for both inlet distortion and total pressure loss at a series of design flight conditions. The resulting optimization problem is difficult since some of the design parameters take on integer values. If numerical procedures could be used to reduce multimillion dollar development test programs to a small set of verification tests, numerical optimization could have a significant impact on both cost and elapsed time to design new aircraft.

Anderson, Bernhard H.↗
NASA NTRS1992

Internal shock interactions in propulsion/airframe integrated three-dimensional sidewall compression scramjet inlets

The advantages and design requirements of propulsion/airframe integration for high Mach number flight are studied in terms of the 3D sidewall compression scramjet inlet. The present work addresses in a parametric fashion the inviscid effects of leading edge sweep, sidewall compression, and inflow Mach number on the internal shock structure in terms of inlet compression and mass capture. The source of the Mach number invariance with leading edge sweep for a constant sidewall compression class of inlet is identified, and a previously undocumented spillage phenomenon in a constant effective wedge angle class of inlets is discussed.

Holland, Scott D.↗
NASA NTRS1992

Computational analysis of ramjet engine inlet interaction

A computational analysis of a ramjet engine at Mach 3.5 has been conducted and compared to results obtained experimentally. This study focuses on the behavior of the inlet both with and without combustor backpressure. Increased backpressure results in separation of the body side boundary layer and a resultant static pressure rise in the inlet throat region. The computational results compare well with the experimental data for static pressure distribution through the engine, inlet throat flow profiles, and mass capture. The computational analysis slightly underpredicts the thickness of the engine body surface boundary layer and the extent of the interaction caused by backpressure; however, the interaction is observed at approximately the same level of backpressure both experimentally and computationally. This study demonstrates the ability of two different Navier-Stokes codes, namely RPLUS and PARC, to calculate the flow features of this ramjet engine and to provide more detailed information on the process of inlet interaction and unstart.

Duncan, Beverly↗
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 NTRS1957

Investigation of an Underslung Normal-wedge Inlet at Free-stream Mach Numbers from 1.50 to 1.99

Two normal -wedge-inlet configurations (straight and sweptback splitter plates) were investigated in the Lewis 8- by 6-foot supersonic wind tunnel at Mach numbers of 1.50 to 1.99. Negligible differences occurred in the performances of the two normal-wedge inlets. In comparison with a previously tested scoop inlet, the normal-wedge configuration showed superior thrust-minus-drag characteristics at Mach 1.99. However the scoop inlet was better at Mach numbers of 1.80 and 1.50.

Vargo, Donald J↗
NASA NTRS1956

Use of Subsonic Diffuser Mach Number as a Supersonic-Inlet Control Parameter

A Mach number measured in the subsonic diffuser was used experimentally as the inlet control parameter of a bypass control system for an axisymmetric supersonic inlet operated in combination with a J34 turbo-jet engine at Mach numbers from 1.6 to 2.0. The control maintained the inlet in either critical or supercritical operation, and, when set for critical diffuser operation, the control recovered from disturbances that placed the inlet in both subcritical buzz and supercritical operation. A slotted-rake orifice gave a more representative value of subsonic diffuser Mach number than the single total-pressure probe used as a control input.

Whalen, Paul P↗
NASA NTRS1951

Analysis of thrust augmentation of turbojet engines by water injection at compressor inlet including charts for calculating compression processes with water injection

A psychrometric chart having total pressure (sum of partial pressures of air and water vapor) as a variable, a Mollier diagram for air saturated with water vapor, and charts showing the thermodynamic properties of various air-water vapor and exhaust gas-water vapor mixtures are presented as aids in calculating the thrust augmentation of a turbojet engine resulting from the injection of water at the compressor inlet. Curves are presented that show the theoretical performance of the augmentation method for various amounts of water injected and the effects of varying flight Mach number, altitude, ambient-air temperature, ambient relative humidity, compressor pressure ratio, and inlet-diffuser efficiency. Numerical examples, illustrating the use of the psychrometric chart and the Mollier diagram in calculating both compressor-inlet and compressor-outlet conditions when water is injected at the compressor inlet, are presented.

Wilcox, E Clinton↗
NASA NTRS1946

Development of Wing Inlets

Lift, drag, internal flow, and pressure distribution measurements were made on a low-drag airfoil incorporating various air inlet designs. Two leading-edge air inlets are developed which feature higher lift coefficients and critical Mach than the basic airfoil. Higher lift coefficients and critical speeds are obtained for leading half of these inlet sections but because of high suction pressures near exist, slightly lower critical speeds are obtained for the entire inlet section than the basic airfoil.

AIRFOILS - DOUGLAS G-17 - CHAR , MOMENTS & ORDINAT↗
NASA NTRS1945

The Effect of Inlet Temperature and Pressure on the Efficiency of a Single-stage Impulse Turbine Having a 13.2-inch Pitch-line Diameter Wheel

Efficiency tests have been conducted on a single-stage impulse turbine having a 13.2-inch pitch-line diameter wheel and a cast nozzle diaphram over a range of turbine speeds from 3000 to 17,000 rpm, pressure ratios from 1.5 to 5.0, inlet total temperatures from 1200 deg to 2000 deg R, and inlet total pressures from 18 to 59 inches of mercury absolute. The effect of inlet temperature and pressure on turbine efficiency for constant pressure ration and blade-to-jet speed ration is correlated against a factor derived from the equation for Reynolds number. The degree of correlation indicates that the change in turbine efficiency with inlet temperature and [ressure for constant pressure ration and blade-to-jet speed ration is principally a Reynolds number effect.

TURBINES - WHEELS↗
NASA NTRS1945

The Effect of Inlet Pressure and Temperature on the Efficiency of a Single-stage Impulse Turbine Having an 11.0-inch Pitch-line Diameter Wheel

Efficiency tests have been conducted on a single-stage impulse engine having an 11-inch pitch-line diameter wheel with inserted buckets and a fabricated nozzle diaphragm. The tests were made to determine the effect of inlet pressure, Inlet temperature, speed, and pressure ratio on the turbine efficiency. An analysis is presented that relates the effect of inlet pressure and temperature to the Reynolds number of the flow. The agreement between the analysis and the experimental data indicates that the changes in turbine efficiency with Inlet pressure and temperature may be principally a Reynolds number effect.

TURBINES, EXHAUST - EFFICIENCY↗
NASA NTRS1946

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 oof 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.

Ferri, Antonio↗