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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 325 records · Page 18
All sourcesNASA NTRSDOE OSTI
NASA NTRS1985

Performance and surge limits of a TF30-P-3 turbofan engine/axisymmetric mixed-compression inlet propulsion system at Mach 2.5

Steady-state performance and inlet-engine compatibility were investigated with a low-bleed inlet. The inlet had minimum internal contraction, consistent with high total pressure recovery and low cowl drag. The inlet-engine combination displayed good performance with only about 2% of inlet performance bleed. The inlet-engine combination had 5.58 deg angle-of-attack capability with 6% bleed.

Wasserbauer, J. F.↗
NASA NTRS1986

A new design concept for indraft wind-tunnel inlets with application to the National Full-Scale Aerodynamics Complex

The present inlet design concept for an indraft wind tunnel, which is especially suited to applications for which a specific test section flwo quality is required with minimum inlet size, employs a cascade or 'vaneset' to control flow at the inlet plane, so that test section total pressure variation is minimized. Potential flow panel methods, together with empirical pressure loss predictions, are used to predict inlet cascade performance. This concept has been used to develop an alternative inlet design for the 80 x 120-ft wind tunnel at NASA Ames Research Center. Experimental results show that a short length/diameter ratio wind tunnel inlet furnishing atmospheric wind isolation and uniform test section flow can be designed.

Ross, J. C.↗
NASA NTRS1986

A new design concept for indraft wind-tunnel inlets with application to the national full-scale aerodynamic complex

The present inlet design concept for an indraft wind tunnel, which is especially suited to applications for which a specific test section flow quality is required with minimum inlet size, employs a cascade or vaneset to control flow at the inlet plane, so that test section total pressure variation is minimized. Potential flow panel methods, together with empirical pressure loss predictions, are used to predict inlet cascade performance. This concept has been used to develop an alternative inlet design for the 80 x 120-ft wind tunnel at NASA Ames Research Center. Experimental results show that a short length/diameter ratio wind tunnel inlet furnishing atmospheric wind isolation and uniform test section flow can be designed.

Ross, James C.↗
NASA NTRS1987

High-speed inlet research program and supporting analyses

A Mach 5 cruise aircraft was studied in a joint program effort. The propulsion system chosen for this aircraft was an over-under turbojet/ramjet system. The ramjet portion of the inlet is to be tested in NASA Lewis' 10 x 10 SWT. Goals of the test program are to obtain performance data and bleed requirements, and also to obtain analysis code validation data. Supporting analysis of the inlet using a three-dimensional Navier-Stokes code (PEPSIS) indicates that sidewall shock/boundary layer interactions cause large separated regions in the corners underneath the cowl. Such separations generally lead to inlet unstart, and are thus a major concern. As a result of the analysis, additional bleed regions were added to the inlet model sidewalls and cowl to control separations in the corners. A two-dimensional analysis incorporating bleed on the ramp is also presented. Supporting experiments for the Mach 5 programs were conducted in the Lewis' 1 x 1 SWT. A small-scale model representing the inlet geometry up to the ramp shoulder and cowl lip was tested to verify the accelerator plate test technique and to obtain data on flow migration in the ramp and sidewall boundary layers. Another study explored several ramp bleed configurations to control boundary layer separations in that region. Design of a two-dimensional Mach 5 cruise inlet represents several major challenges including multimode operation and dual flow, high temperatures, and three-dimensional airflow effects.

Coltrin, Robert E.↗
NASA NTRS1990

Sea level static calibration of a compact multimission aircraft propulsion simulator with inlet flow distortion

Wind tunnel tests of propulsion-integrated aircraft models have identified inlet flow distortion as a major source of compressor airflow measurement error in turbine-powered propulsion simulators. Consequently, two Compact Multimission Aircraft Propulsion Simulator (CMAPS) units were statically tested at sea level ambient conditions to establish simulator operating performance characteristics and to calibrate the compressor airflow against an accurate bellmouth flowmeter in the presence of inlet flow distortions. The distortions were generated using various-shaped wire mesh screens placed upstream of the compressor. CMAPS operating maps and performance envelopes were obtained for inlet total pressure distortions (ratio of the difference between the maximum and minimum total pressures to the average total pressure) up to 35 percent, and were compared to baseline simulator operating characteristics for a uniform inlet. Deviations from CMAPS baseline performance were attributed to the coupled variation of both compressor inlet-flow distortion and Reynolds number index throughout the simulator operating envelope for each screen configuration. Four independent methods were used to determine CMAPS compressor airflow; direct compressor inlet and discharge measurements, an entering/exiting flow-balance relationships, and a correlation between the mixer pressure and the corrected compressor airflow. Of the four methods, the last yielded the least scatter in the compressor flow coefficient, approximately + or - 3 percent over the range of flow distortions.

Won, Mark J.↗
NASA NTRS1991

Zonal analysis of two high-speed inlets

Using a zonal technique, thin layer Navier-Stokes solutions for two high speed inlet geometries are presented and compared with experimental data. The first configuration consists of a 3-D inlet preceded by a sharp flat plate. Results with two different grids demonstrate the importance of adequate grid refinement in high speed internal flow computations. The fine grid solution has reasonably good agreement with experimental heat transfer and pressure values inside the inlet. The other configuration consists of a 3-D inlet mounted on a research hypersonic forebody. Numerical results for this configuration have good agreement with experimental pressure data along the forebody, but not inside the inlet. A more refined grid calculation is currently being done to better predict the flowfield in the inlet.

Dilley, A. D.↗
NASA NTRS1991

Results from computational analysis of a mixed compression supersonic inlet

A numerical study was performed to simulate the critical flow through a supersonic inlet. This flowfield has many phenomena such as shock waves, strong viscous effects, turbulent boundary layer development, boundary layer separations and mass flow suction through the walls (bleed). The computational tools used in this study were two full Navier-Stokes (FNS) codes. The supersonic inlet that was analyzed in this study is the Variable Diameter Centerbody (VDC), inlet. This inlet is a candidate concept for the next generation supersonic transport. Application of the code to the inlet geometry involved effort in generating an efficient grid geometry and specifying boundary conditions, particularly in the bleed region and at the outflow boundary. Results for a critical inlet operation compare favorably to Method of Characteristics predictions and experimental data.

Saunders, J. D.↗
NASA NTRS1993

An experimental study of the effects of bodyside compression on forward swept sidewall compression inlets ingesting a turbulent boundary layer

Forward swept sidewall compression inlets have been tested in the Mach 4 Blowdown Facility at the NASA Langley Research Center to study the effects of bodyside compression surfaces on inlet performance in the presence of an incoming turbulent boundary layer. The measurements include mass flow capture and mean surface pressure distributions obtained during simulated combustion pressure increases downstream of the inlet. The kerosene-lampblack surface tracer technique has been used to obtain patterns of the local wall shear stress direction. Inlet performance is evaluated using starting and unstarting characteristics, mass capture, mean surface pressure distributions and permissible back pressure limits. The results indicate that inlet performance can be improved with selected bodyside compression surfaces placed between the inlet sidewalls.

Rodi, Patrick E.↗
NASA NTRS1993

Installed F/A-18 inlet flow calculations at 60 deg angle-of-attack and 10 deg side slip

This paper presents the results of PARC3D numerical calculations on a 19.78 percent scale forebody/inlet model of the F/A-18 at a Mach number of 0.20, an angle-of-attack of 60 deg, and a side-slip angle of 10 deg. The main purpose of these calculations is to support an upcoming wind-tunnel test program in the prediction of engine inlet compressor face total pressure recovery and flow distortion. The GRIDGEN system was used to generate a grid which includes the inlet and lip, and other aircraft components which are considered to be important to inlet performance, such as the ramp/splitter plate, the diverter and slot, and the deflected leading edge flap. PARC3D shows complex flow patterns on the fuselage surfaces below the leading edge extensions, on the ramp/splitter plate, inlet lip, and inside the inlet. PARC3D tends to underpredict total pressure recovery and overpredict the flow distortion at the inlet compressor face.

Podleski, S. D.↗
NASA NTRS1993

Computational effects of inlet representation on powered hypersonic, airbreathing models

Computational results are presented to illustrate the powered aftbody effects of representing the scramjet inlet on a generic hypersonic vehicle with a fairing, to divert the external flow, as compared to an operating flow-through scramjet inlet. This study is pertinent to the ground testing of hypersonic, airbreathing models employing scramjet exhaust flow simulation in typical small-scale hypersonic wind tunnels. The comparison of aftbody effects due to inlet representation is well-suited for computational study, since small model size typically precludes the ability to ingest flow into the inlet and perform exhaust simulation at the same time. Two-dimensional analysis indicates that, although flowfield differences exist for the two types of inlet representations, little, if any, difference in surface aftbody characteristics is caused by fairing over the inlet.

Huebner, Lawrence D.↗
NASA NTRS1953

Theoretical Performance Characteristics of Sharp Lip Inlets at Subsonic Speeds

A method is presented for the estimation of the subsonic-flight-speed characteristics of sharp-lip inlets applicable to supersonic aircraft. The analysis, based on a simple momentum balance consideration, permits the computation of inlet pressure recovery - mass-flow relations and additive-drag coefficients for forward velocities from zero to the speed of sound. The penalties for operation of a sharp-lip inlet at velocity ratios other than 1.0 may be severe; at lower velocity ratios an additive drag is incurred that is not cancelled by lip suction, while at higher velocity ratios, unavoidable losses in inlet total pressure will result. In particular, at the take-off condition, the total pressure and the mass flow for a choked inlet are only 79 percent of the values ideally attainable with a rounded lip. Experimental data obtained at zero speed with a sharp-lip supersonic inlet model were in substantial agreement with the theoretical results.

Fradenburgh, Evan A↗
NASA NTRS1944

Dynamics of the inlet system of a four-stroke engine

Tests were run on a single-cylinder and a multicylinder four-stroke engine in order to determine the effect of the dynamics of the inlet system upon indicated mean effective pressure. Tests on the single-cylinder engine were made at various speeds, inlet valve timings, and inlet pipe lengths. These tests indicated that the indicated mean effective pressure could be raised considerably at any one speed by the use of a suitably long inlet pipe. Tests at other speeds with this length of pipe showed higher indicated mean effective pressure than with a very short pipe, although not so high as could be obtained with the pipe length adjusted for each speed. A general relation was discovered between optimum time of inlet valve closing and pipe length; namely, that longer pipes require later inlet valve closing in order to be fully effective. Tests were also made on three cylinders connected to a single pipe. With this arrangement, increased volumetric efficiency at low speed was obtainable by using a long pipe, but only with a sacrifice of volumetric efficiency at high speed. Volumetric efficiency at high speed was progressively lower as the pipe length was increased.

ENGINES-CHRYSLER↗
NASA NTRS1954

Theoretical performance characteristics of sharp-lip inlets at subsonic speeds

A method is presented for the estimation of the subsonic-flight-speed characteristics of sharp-lip inlets applicable to supersonic aircraft. The analysis, based on a simple momentum balance consideration, permits the computation of inlet-pressure-recovery mass-flow relations and additive-drag coefficients for forward velocities from zero to the speed of sound. The penalties for operation of a sharp-lip inlet at velocity ratios other than 1.0 may be severe; at lower velocity ratios an additive drag is incurred that is not cancelled by lip suction, while at higher velocity ratios, unavoidable losses in inlet total pressure will result. In particular, at the take-off condition, the total pressure and the mass flow for a choked inlet are only 79 percent of the values ideally attainable with a rounded lip. The test specimens were polished and unnotched. The manufacturer of the material, the Aluminum Company of America, has made axial-load tests on 24S-T4 and 75S-T6 rod material. The test techniques used at the three laboratories are described in detail; the test results are presented and are compared with each other and with results obtained on unpolished sheet by the National Bureau of Standards. Experimental data obtained at zero speed with a sharp-lip supersonic inlet model were in substantial agreement with the theoretical results. (author)

Fradenburgh, Evan A↗
NASA NTRS1948

Analysis of Effects of Inlet Pressure Losses on Performance of Axial-Flow Type Turbojet Engine

The experimentally determined performance characteristics of an axial-flow turbojet engine have been used to estimate the effects of inlet total-pressure losses on net thrust and specific fuel consumption at a constant engine speed. At low altitudes and flight Mach numbers, inlet pressure losses cause an increase in engine discharge temperature and it is possible that the maximum allowable turbine temperature maybe exceeded. An inlet absolute total-pressure loss of 10 percent will result in a thrust loss of 14 percent and a 15-percent increase in specific fuel consumption based on net thrust. At high altitudes and flight Mach numbers, choking conditions exist in the exhaust nozzle and the inlet pressure losses do not affect the discharge temperatures. Under these conditions, a 10-percent loss in inlet absolute total pressure produces a 22-percent loss in net thrust and a 16-percent increase in specific fuel consumption. If the exhaust-nozzle-outlet area is adjusted to compensate for the effect of inlet losses on discharge temperature in the nonchoking cases (low altitude and Mach numbers), the thrust and fuel consumption will be changed in a manner similar to the results obtained in the choking cases.

Sanders, Newell D↗
NASA NTRS1956

Zero-angle-of-attack Performance of Two-dimensional Inlets near Mach Number 3

An extensive program was undertaken to investigate the effect of several geometric variables on the performance of two-dimensional inlets. This investigation included inlets having single-wedge, double-wedge, and isentropic compression ramps with various side-plate configurations and subsonic diffusers. The tests were conducted over a range of Reynolds number based on inlet height from 0.50 to 2.67(sup x)10(sup 6). Generally, the performance levels of the two-dimensional inlets were somewhat below those obtained previously with comparable axisymmetric models. At Mach number 3.05 the optimum total-pressure recovery was obtained with an isentropic inlet which compressed the external flow to a Mach number of 1.88. Rectangular side plates and a long high-exit-Mach-number subsonic diffuser with filleted corners were used with this inlet. A critical total-pressure recovery of 0.71 was realized with a corresponding mass-flow ratio of 0.965. Subcritical stability to a mass-flow ratio of 0.60 was obtained.

Woollett, Richard R↗
NASA NTRS1994

An Interactive, Design and Educational Tool for Supersonic External-Compression Inlets

A workstation-based interactive design tool called VU-INLET was developed for the inviscid flow in rectangular, supersonic, external-compression inlets. VU-INLET solves for the flow conditions from free stream, through the supersonic compression ramps, across the terminal normal shock region and the subsonic diffuser to the engine face. It calculates the shock locations, the capture streamtube, and the additive drag of the inlet. The inlet geometry can be modified using a graphical user interface and the new flow conditions recalculated interactively. Free stream conditions and engine airflow can also be interactively varied and off-design performance evaluated. Flow results from VU-INLET can be saved to a file for a permanent record, and a series of help screens make the simulator easy to learn and use. This paper will detail the underlying assumptions of the models and the numerical methods used in the simulator.

Benson, Thomas J.↗
NASA NTRS1992

Status of an inlet configuration trade study for the Douglas HSCT

An inlet concept integration trade study for an HSCT is being conducted under contract to NASA LeRC. The HSCT mission has a supersonic cruise Mach number of 2.4 and a subsonic cruise Mach number of 0.95. The engine selected for this study is the GE VCE (variable cycle engine) with FLADE (fan on blade). Six inlet configurations will be defined. Inlet configurations will be axisymmetric and rectangular mixed-compression inlets in single-engine nacelles. Airplane performance for each inlet configuration will be estimated and then compared. The most appropriate inlet configuration for this airplane/engine combination will be determined by Sep. 1991.

Jones, Jay R.↗
NASA NTRS1995

Prediction of Wind Tunnel Effects on the Installed F/A-18A Inlet Flow Field at High Angles-of-attack

NASA Lewis is currently engaged in a research effort as a team member of the High Alpha Technology Program (HATP) within NASA. This program utilizes a specially equipped F/A-18A, the High Alpha Research Vehicle (HARV), in an ambitious effort to improve the maneuverability of high performance military aircraft at low subsonic speed, high angle of attack conditions. The overall objective of the Lewis effort is to develop inlet technology that will ensure efficient airflow delivery to the engine during these maneuvers. One part of the Lewis approach utilizes computational fluid dynamics codes to predict the installed performance of inlets for these highly maneuverable aircraft. Wind tunnel tests were a major component of the Lewis program. Since the available wind tunnel was small (9 x 15 ft) as compared to the scale of the model of the F/A-18A (19.78 percent), there were questions about the capability to obtain useful inlet performance data. The blockage effects were expected to be very large. This report represents the results of an analysis to determine how the wind tunnel walls effect inlet performance at several angles of attack. The predictions for the external particle traces along the fuselage indicate the influence of the wind tunnel side wall under the model is greater at 30 deg angle of attack than at 50 deg angle of attack on the under Leading Edge Extension (LEX) vortex trajectory. The side wall above the model appears to have negligible influence on the under LEX vortex. This may be due to the LEX acting as 'shield' to the upper wall effects. As expected, the wind tunnel has a significant influence on the external forces. The lift and drag coefficients increase significantly for the wind tunnel model as compared to free stream conditions. The wind tunnel had a small effect on the inlet recovery and on inlet total pressure distortion patterns. The predicted recoveries for the wind tunnel model are within one percentage point of the model recoveries in free stream conditions.

Smith, Crawford F.↗