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At least 1,315 records · Page 73
All sourcesNASA NTRSDOE OSTI
NASA NTRS1971

Flight investigation of airframe installation effects on an auxiliary inlet ejector nozzle on an underwing engine nacelle

The local flow field approaching an installed nozzle may vary from isolated test conditions, thereby affecting exhausting nozzle performance. An installation of general interest is a podded engine mounted near the aft lower surface of the wing. The effect of this installation on the performance of an auxiliary inlet ejector nozzle was investigated over a Mach number range of 0.7 to 1.3 by using a modified F-106B aircraft. Both floating and fixed-open door configurations were examined. The ejector nozzle trailing-edge flaps were simulated in the closed position with rigid structure which provided a boattail angle of 15 deg. Primary nozzle area was varied as exhaust gas temperature was varied between 982.2 and 2003.3 K.

Burley, R. R.↗
NASA NTRS1972

High-capacity compact vortex valve for increasing stability of supersonic mixed-compression inlets

A special vortex valve was developed for increasing the stability margin of supersonic jet-engine inlets without the use of moving mechanical parts. A compact valve with high flow capacity and with high gain was obtained by using a unique outer wall in the vortex chamber instead of the usual cylindrical shape. The physical features of the valve are described, and the steady pressure-flow characteristics measured on scale models are given. Procedures are provided for adjusting the operating pressure, flow capacity, and gain through systematic changes in geometry.

Gebben, V. D.↗
NASA NTRS1972

Dynamic flow angle measurements at the diffuser exit of a Mach 2.5 mixed compression inlet

An exploratory investigation was made to determine the magnitude of the dynamic flow angle at a single point in the diffuser exit of a Mach 2.50 inlet. Results indicate that the maximum flow angle variation was + or - 3 degrees except when vortex generators were installed in the subsonic diffuser. The amplitude of these flow angle fluctuations was directly related to the amplitude of the total pressure fluctuations at the probe survey point. Flow angle fluctuations as high as + or - 11 degrees were recorded when vortex generators were installed near the probe survey point.

Prati, W. M.↗
NASA NTRS1972

Performance of a bicone inlet designed for Mach 2.5 with internal distributed compression and 40 percent internal contraction

The inlet was designed to have the minimum internal contraction consistent with high total-pressure recovery and low cowl drag. Without a bypass system, the peak pressure recoveries increased from 0.890 to 0.936 when the supercritical bleed mass flow ratio was varied from 0.035 to 0.060. With an operating bypass system and installed centerbody vortex generators, a slight increase in peak pressure recovery was obtained. The values of steady-state distortion and dynamic distortion were below 0.10 and 0.02, respectively, near critical operation. Simulation of a turbofan engine with concentric pipes showed no effect on compressor face flow profiles with varying bypass flow ratio.

Wasserbauer, J. F.↗
NASA NTRS1972

Identification of boiler inlet transfer functions and estimation of system parameters

An iterative computer method is described for identifying boiler transfer functions using frequency response data. An objective penalized performance measure and a nonlinear minimization technique are used to cause the locus of points generated by a transfer function to resemble the locus of points obtained from frequency response measurements. Different transfer functions can be tried until a satisfactory empirical transfer function of the system is found. To illustrate the method, some examples and some results from a study of a set of data consisting of measurements of the inlet impedance of a single tube forced flow boiler with inserts are given.

Miles, J. H.↗
NASA NTRS1972

Predicted upwash angles at engine inlets for STOL aircraft

Upwash angles were predicted for a STOL lifting system by using a two-dimensional potential flow analysis. Upwash angles are presented for distances ahead of the wing leading edge of 50, 75, and 100 percent of wing chord. The upwash angle was determined to be insensitive to the vertical location of the engine inlet. For a wide range of takeoff and landing conditions, the upwash angle was found to be a function of the total two-dimensional lift coefficient. Upwash angles, along with typical flow fields, are presented for a range of total two-dimensional lift coefficients from 2 to 12. Three-dimensional effects were considered in estimating upwash angles for an unswept-wing externally blown flap aircraft. For this STOL configuration, effective upwash angles during takeoff, approach, and waveoff conditions were found to be 22 deg, 26 deg, and 36 deg, respectively.

Albers, J. A.↗
NASA NTRS1972

Performance estimates for a supersonic axisymmetric inlet system.

Results of recent tests of a large-scale axisymmetric mixed-compression inlet system designed for Mach number 2.65 indicate that the performance of the supersonic diffuser and required boundary-layer bleed system can be accurately estimated with relatively new analytic methods. Thus, many wind tunnel testing hours can be saved by avoiding much of the usual ?cut and try' wind tunnel development.

Sorensen, N. E.↗
NASA NTRS1972

Flight investigation of an underwing nacelle installation of an auxiliary-inlet ejector nozzle with a clamshell flow diverter from Mach 0.6 to 1.3

A nozzle installation of general interest is a podded engine mounted near the aft lower surface of the wing. The effect of this installation on the performance of an auxiliary-inlet ejector nozzle with a clamshell flow diverter was investigated over a Mach number range of 0.6 to 1.3 by using a modified F-106B aircraft. The clamshell flow diverter was tested in a 17 deg position with double-hinged synchronized floating doors. The ejector nozzle trailing-edge flaps were simulated in the closed position with a rigid structure which provided a boattail angle of 10 deg. Primary nozzle area was varied as exhaust gas temperature was varied between 975 and 1561 K. With the nozzle in a subsonic cruise position, the nozzle gross thrust coefficient was 0.918 at a flight Mach number of 0.9.

Head, V. L.↗
NASA NTRS1973

Design of a bleed system for a Mach 3.5 inlet

An analytical bleed system design procedure is applied to a Mach 3.5, axisymmetric, translating centerbody inlet. The design features a traveling centerbody bleed system with a ducting arrangement separating low-, medium-, and high-pressure bleed. Three forward plenums and nine throat plenums are installed on the centerbody to meet the bleed requirements in the started Mach range between 1.6 and 3.5. The cowl contains four stationary plenums, three for forward and one for throat bleed. To achieve maximum bleed plenum pressure and thereby minimize bleed drag all bleed holes are inclined 20 deg to the surface except in the cowl throat region. Here the requirement of high stability margin with minimum total pressure recovery penalty resulted in 90 deg bleed holes. The bleed hole diameter varies from bleed plenum to bleed plenum to achieve the most efficient boundary layer control, while the bleed exits are sized to operate at the highest possible plenum pressure without unchoking the bleed holes. Bleed flow rates, bleed plenum pressures, and boundary layer development along the cowl and centerbody are predicted over the entire started Mach range.

Syberg, J.↗
NASA NTRS1973

The effect of nozzle inlet shape, lip thickness, and exit shape and size on subsonic jet noise

Far field noise data were taken for convergent nozzles of various shapes and sizes at subsonic velocities exceeding 400 feet per second. For a circular nozzle, the nozzle inlet shape and lip thickness had no effect on the noise level, directivity, or spectra when compared at the same nozzle exit diameter and peak exhaust velocity. A sharp edged orifice was one exception to this statement. Coannular nozzles can produce additional high frequency noise. Blunt ended centerbodies, where there is significant base drag, also generate significant additional noise. The total noise power generation was essentially the same for circular, slot, and plug nozzles of good aerodynamic shape. The noise radiation patterns were essentially the same for these nozzle shapes except near the nozzle exhaust axis.

Olsen, W. A.↗