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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 973 records · Page 54
All sourcesNASA NTRSDOE OSTI
NASA NTRS1980

YF-12 inlet suppression of compressor noise - First results

An aeroacoustic test program was performed with a YF-12 aircraft at ground-static conditions. The objective was to collect acoustic and aerodynamic data that could determine the cause of inlet noise suppression observed earlier. The first results of the test program are presented here. There was no indication that the flow was close to choking. The data indicated significant reduction in sound pressure level (SPL) across the strut and bypass region at frequencies near the blade passing. Far-field data showed that the maximum sound pressure level near the blade-passing frequency was at zero degrees from the inlet centerline.

Bangert, L. H.↗
NASA NTRS1981

Ultra-lean combustion at high inlet temperatures

Combustion at inlet air temperatures of 1100 to 1250 K was studied for application to advanced automotive gas turbine engines. Combustion was initiated by the hot environment, and therefore no external ignition source was used. Combustion was stabilized without a flameholder. The tests were performed in a 12 cm diameter test section at a pressure of 2.5 x 10 to the 5th power Pa, with reference velocities of 32 to 60 m/sec and at maximum combustion temperatures of 1350 to 1850 K. Number 2 diesel fuel was injected by means of a multiple source fuel injector. Unburned hydrocarbons emissions were negligible for all test conditions. Nitrogen oxides emissions were less than 1.9 g NO2/kg fuel for combustion temperatures below 1680 K. Carbon monoxide emissions were less than 16 g CO/kg fuel for combustion temperatures greater than 1600 K, inlet air temperatures higher than 1150 K, and residence times greater than 4.3 microseconds.

Anderson, D. N.↗
NASA NTRS1981

Some aspects of calculating flows about three-dimensional subsonic inlets

Based on the potential flow model, computations were carried out for various three-dimensional inlet models. Some of these calculated results are presented in the forms of surface static pressure, flow angularity, surface flow pattern, and inlet flow field. Comparisons with experimental data are also made.

Kao, H. C.↗
NASA NTRS1980

Seasat detection of waves, currents and inlet discharge

A new era of remote sensing for coastal and oceanographic monitoring was born on June 26, 1978 with the launch of Seasat. Duck-X was a 2 month experiment conducted during August to October 1978 off the east coast of the U.S.A. for the validation of the Seasat synthetic aperture radar (SAR). During this field experiment, various oceanographic phenomena were monitored. Ground truth observations of these phenomena have been correlated with Seasat SAR imagery. The ground truth sensors included airborne photographic and radar imagery, meteorological satellite imagery, land based radars, and conventional wave gauges. The direction and length of the principal ocean wave trains are compared for the periods of Seasat overflight of the Duck-X area. During these overflights significant wave heights were 1.5 m and less and the maximum wave period was 15 sec. The current correlations concentrate on the western boundary of the Gulf Stream and its associated eddy structure. Inlet outflow is shown for inlets on the east coast of the U.S.A. This ground truth study has indicated that the SAR imagery contains an unanticipated abundance of information on a variety of oceanographic and coastal phenomena.

Mattie, M. G.↗
NASA NTRS1981

Ultra-lean combustion at high inlet temperatures

Combustion at inlet-air temperatures of 1100 to 1250 K was studied for application to advanced automotive gas turbine engines. Combustion was initiated by the hot environment, and therefore no external ignition source was used. Combustion was stabilized without a flameholder. The tests were performed in a 12-cm-diameter test section at a pressure of 250,000 Pa, with reference velocities of 32 to 60 m/s and at maximum combustion temperatures of 1350 to 1850 K. Number 2 diesel fuel was injected by means of a multiple-source fuel injector. Unburned hydrocarbon emissions were negligible for all test conditions. Nitrogen oxide emissions were less than 1.9 g NO2/kg fuel for combustion temperatures below 1680 K. Carbon monoxide emissions were less than 16 g CO/kg fuel for combustion temperatures greater than 1600 K, inlet-air temperatures higher than 1150 K, and residence times greater than 4.3 ms.

Anderson, D. N.↗
NASA NTRS1981

Effect of a part span variable inlet guide vane on TF34 fan performance

Experimental aerodynamic and performance data were obtained from a TF34 engine. Part span variable inlet guide vanes mounted in front of the fan on the TF34 engine were tested to demonstrate the feasibility of modulating air flow and thrust for vertical takeoff aircraft systems. The fan was mapped to stall for a range of speeds and variable inlet guide were settings. Modulated fan tip performance and unmodulated hub performance were evaluated with a without an extended fan bypass splitter. The effect of a crosswind distortion screen on performance was also evaluated.

Alvarez, J.↗
NASA NTRS1982

Experimental and analytical results of tangential blowing applied to a subsonic V/STOL inlet

Engine inlets for subsonic V/STOL aircraft must operate over a wide range of conditions without internal flow separation. Experimental and analytical investigations were conducted to evaluate the effectiveness of tangential blowing to maintain attached flow to high angles of attack. The inlet had a relatively thin lip with a blowing slot located either on the lip or in the diffuser. The height and width of these slots was varied. Experimentally determined flow separation boundaries showed that lip blowing achieved higher angle of attack capability than diffuser blowing. This capability was achieved with the largest slot circumferential extent and either of the two slot heights. Predicted (analytical) separation boundaries showed good agreement except at the highest angles of attack.

Burley, R. R.↗
NASA NTRS1981

Some aspects of calculating flows about three-dimensional subsonic inlets

Various three dimensional inlet models were calculated based on the potential flow model. Results are presented in the forms of surface static pressure, flow angularity, surface flow pattern, and inlet flow field. It is indicated that the extension of the lower lip can reduce the adverse pressure gradient and increase the flow separation bound.

Kao, H. C.↗
NASA NTRS1982

Three-dimensional flow calculations including boundary layer effects for supersonic inlets at angle of attack

An analysis is presented for calculating the steady three-dimensional flow field in supersonic mixed-compression inlets at incidence. A zonal modeling approach is employed to obtain the solution. The supersonic core flow is computed using a second-order pentahedral bicharacteristic algorithm. The bow shock wave and the reflected internal shock train are determined using a three-dimensional discrete shock fitting procedure. The boundary layer flow adjacent to both the centerbody and the cowl is computed using a second-order implicit finite difference method. The flow in a shock wave-boundary layer interaction region is computed using an integral formulation. The culmination of the present research effort is the development of a production-type computer program capable of analyzing flow in a variety of mixed-compression aircraft inlets. Numerical results and experimental correlations are presented to illustrate application of the analysis.

Vadyak, J.↗
NASA NTRS1982

Tangential blowing for control of strong normal shock - Boundary layer interactions on inlet ramps

The use of tangential blowing from a row of holes in an aft facing step is found to provide good control of the ramp boundary layer, normal shock interaction on a fixed geometry inlet over a wide range of inlet mass flow ratios. Ramp Mach numbers of 1.36 and 1.96 are investigated. The blowing geometry is found to have a significant effect on system performance at the highest Mach number. The use of high-temperature air in the blowing system, however, has only a slight effect on performance. The required blowing rates are significantly high for the most severe test conditions. In addition, the required blowing coefficient is found to be proportional to the normal shock pressure rise.

Schwendemann, M. F.↗
NASA NTRS1982

Utilizing numerical techniques in turbofan inlet acoustic suppressor design

Numerical theories in conjunction with previously published analytical results are used to augment current analytical theories in the acoustic design of a turbofan inlet nacelle. In particular, a finite element-integral theory is used to study the effect of the inlet lip radius on the far field radiation pattern and to determine the optimum impedance in an actual engine environment. For some single mode JT15D data, the numerical theory and experiment are found to be in a good agreement.

Baumeister, K. J.↗
NASA NTRS1981

Experimental investigation of tangential blowing for control of the strong shock boundary layer interaction on inlet ramps

A 0.165-scale isolated inlet model was tested in the NASA Lewis Research Center 8-ft by 6-ft Supersonic Wind Tunnel. Ramp boundary layer control was provided by tangential blowing from a row of holes in an aft-facing step set into the ramp surface. Testing was performed at Mach numbers from 1.36 to 1.96 using both cold and heated air in the blowing system. Stable inlet flow was achieved at all Mach numbers. Blowing hole geometry was found to be significant at 1.96M. Blowing air temperature was found to have only a small effect on system performance. High blowing levels were required at the most severe test conditions.

Schwendemann, M. F.↗
NASA NTRS1983

Validation of a three-dimensional viscous analysis of axisymmetric supersonic inlet flow fields

A three-dimensional viscous marching analysis for supersonic inlets was developed. To verify this analysis several benchmark axisymmetric test configurations were studied and are compared to experimental data. Detailed two-dimensional results for shock-boundary layer interactions are presented for flows with and without boundary layer bleed. Three dimensional calculations of a cone at angle of attack and a full inlet at attack are also discussed and evaluated. Results of the calculations demonstrate the code's ability to predict complex flow fields and establish guidelines for future calculations using similar codes.

Benson, T. J.↗
NASA NTRS1983

Numerical analysis of a scramjet inlet flow field using the three-dimensional Navier-Stokes equations

A computer program has been developed to analyze supersonic combustion ramjet (scramjet) inlet flow fields. The program solves the three-dimensional Euler or Navier-Stokes equations in full conservation form by a well-known explicit, predictor-corrector technique. Turbulence is modeled by an algebraic eddy-viscosity model. Detailed laminar and turbulent flow results are presented for a symmetric wedge corner and a comparison is made with the available experimental results to allow assessment of the program. Results are then presented for an actual scramjet inlet configuration.

Kumar, A.↗
NASA NTRS1983

Flow fields and aerodynamic characteristics for hypersonic missiles with mid-fuselage inlets

A study was made to quantify forebody flow fields and to evaluate aerodynamic performance trends on a matrix of fuselage shapes for the mid-inlet/bolt-on-engine class of hypersonic airbreathing missiles for the Navy's vertical box launcher. The study indicated that inlet mass flow and pressure recovery can be increased by cambering the nose and increasing the width of the fuselage at both Mach 4 acceleration and Mach 6 cruise conditions. Aerodynamic trim predictions show that the drag at zero lift at Mach 4 decreases while the L/D max at Mach 6 increases with the nose camber, although these tendencies reverse with increasing width of maximum fuselage cross section.

Hunt, J. L.↗
NASA NTRS1983

The design and flight test of an engine inlet bulk acoustic liner

This paper summarizes the design, fabrication and flight evaluation of a Kevlar acoustic liner configuration for a JT15D turbofan engine. The liner was designed to suppress, by a measurable amount, a dominant (13,0) BPF tone. This tone or spinning mode was produced for research purposes by installing 41 circumferentially distributed small diameter rods upstream of the 28 fan blades. Duct liner attenuations calculated by a finite element procedure were compared to far field power (insertion) losses deduced from flight data. The finite element program modeled the variable geometry of the JT15D inlet and used a uniform flow with a boundary layer roll-off to model the inlet flow field. Calculated liner losses were generally conservative. That is, measured far field power losses were generally greater than attenuations calculated by the finite element computer program.

Lester, H. C.↗
NASA NTRS1983

Calculation of a simulated 3-D high speed inlet using the Navier-Stokes equations

A hybrid numerical algorithm, developed to solve the full three-dimensional Navier-Stokes equations, is applied to the computation of the flowfield in a simulated three-dimensional high speed aircraft inlet at a Mach number of 2.5 and Reynolds number of 1.4 x 10 to the 7th based on inlet length. The numerical algorithm incorporates a coordinate transformation in order to handle general flow geometries, and utilizes the algebraic turbulent eddy viscosity model of Baldwin and Lomax. The hybrid algorithm has been vectorized on the CDC CYBER 203 computer using the SL/1 vector programming language developed at NASA Langley. The computed results are compared with experimental measurements of the ramp and cowl static pressures, and boundary layer pitot profiles. The results are also compared with a previous two-dimensional Navier-Stokes computation of the same configuration. The agreement with the experimental data is generally good; however, additional improvements in turbulence modeling are needed.

Knight, D. D.↗
NASA NTRS1982

Experimental evaluation of inlet turbulence, wall boundary layer, surface finish, and fillet radius on small axial turbine state performance

The results of an experimental investigation are presented which establish the effects of surface finish, fillet radius, inlet boundary layer thickness, and free-stream inlet turbulence level on the aerodynamic performance of a small axial flow turbine stator. The principal objective was to help understand why large turbine efficiency is not maintained when a large turbine is scaled to a smaller size and to provide the turbine designer with the performance compromises expected for a small scale design. A comprehensive test matrix was used to gain an understanding of the effects of each variable over the full range of all the other variables.

Kozak, A. A.↗