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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 919 records · Page 51
All sourcesNASA NTRSDOE OSTI
NASA NTRS1974

Rational functional representation of flap noise spectra including correction for reflection effects

A rational function is presented for the acoustic spectra generated by deflection of engine exhaust jets for under-the-wing and over-the-wing versions of externally blown flaps. The functional representation is intended to provide a means for compact storage of data and for data analysis. The expressions are based on Fourier transform functions for the Strouhal normalized pressure spectral density, and on a correction for reflection effects based on Thomas' (1969) N-independent-source model extended by use of a reflected ray transfer function. Curve fit comparisons are presented for blown-flap data taken from turbofan engine tests and from large-scale cold-flow model tests. Application of the rational function to scrubbing noise theory is also indicated.

Miles, J. H.↗
NASA NTRS1976

Inlet spillage drag tests and numerical flow-field analysis at subsonic and transonic speeds of a 1/8-scale, two-dimensional, external-compression, variable-geometry, supersonic inlet configuration

Accurate spillage drag and pressure data are presented for a realistic supersonic inlet configuration. Results are compared with predictions from a finite-differencing, inviscid analysis computer procedure. The analytical technique shows good promise for the evaluation of inlet drag, but necessary refinements were identified. A detailed description of the analytical procedure is contained in the Appendix.

Hawkins, J. E.↗
NASA NTRS1985

Simulating Pressure Effects of High-Flow Volumes

Dynamic test stresses realized without high-volume pumps. Assembled in Sections in gas-flow passage, contoured mandrel restricts flow rate to valve convenient for testing and spatially varies pressure on passage walls to simulate operating-pressure profile. Realistic test pressures thereby achieved without extremely high flow volumes.

Kaufman, M.↗
NASA NTRS1952

Design and Test of Mixed-flow Impellers II : Experimental Results, Impeller Model MFI-1A

An investigation was conducted to determine the performance characteristics of a mixed-flow impeller which was designed with special emphasis on the reduction or elimination of flow decelerations along wetted surfaces. The performance was investigated over a range of equivalent impeller speeds from 700 to 1600 feet per second and over a range of flow rates from maximum adiabatic temperature-rise efficiency, measured at 1-1/2 impeller diameters in a vaneless diffuser, at the design speed of 1400 feet per second were 4.00 and 0.83, respectively.

THEORY AND EXPERIMENT↗
NASA NTRS1953

Low-speed Aileron Effectiveness as Determined by Force Tests and Visual-flow Observations on a 52 Degree Sweptback Wing with and Without Chord-extensions

The results of an investigation of the effects of leading-edge chord-extensions and extensible leading-edge flaps on the low-speed aileron characteristics of a 52 degrees sweptback wing with circular-arc airfoil sections are presented. A description and illustrations of the vortex flow as it passes over the wing are included. The investigation was conducted at Reynolds numbers of 5.5 times 10(exp 6) and 1.3 times 10(exp 6) and corresponding Mach numbers of 0.11 and 0.065 for an angle-of-attack range of 0 degrees through stall.

CONTROLS, FLAP-TYPE - COMPLETE WINGS↗
NASA NTRS1954

Design and Test of Mixed-flow Impellers IV : Experimental Results for Impeller Models MFI-1 and MFI-2 with Changes in Blade Height

Modifications A and B of impeller model MFI-1 and A, B, and C of impeller model MFI-2 were investigated experimentally in an attempt to determine what allowance in blade height should be made for boundary layer and viscous losses in an impeller designed for isentropic compressible flow. A gradual increase in blade height was arbitrarily made from inlet to outlet in anticipation of a gradual build-up of boundary layer. Apparently there was a rapid build-up of boundary layer near the inlet in the experimental case rather than a gradual one. Therefore, the proper allowance for boundary layer cannot be described from the data obtained. Decreasing the pressure gradient along the shroud by reducing the blade height allowance apparently did little to increase the overall efficiency. At the design speed of 1400 feet per second, the overall adiabatic efficiency was increased from 0.83 for the MFI-1A to 0.85 for the MFI-1B with reduction in height; however, it is indicated from the theoretical velocity distribution and outlet surveys that the increase was due to a change from decelerating to accelerating flow along the hub rather than from any change along the shroud. It is further indicated that the consequences of a thickened or separated boundary layer depend not only on the design velocity gradients but also on the shape of the passage.

Hamrick, Joseph T↗
NASA NTRS1998

The Advanced Reusable Technologies Project

Currently, NASA and its industry partners are performing ground testing of hydrogen-fueled rocket based combined-cycle flowpaths. Successful ramjet and scramjet testing at Mach 6 and scramjet testing at Mach 8 have been performed. Cold flow mixing tests have also been successfully performed as have inlet operability tests. Additional testing of RBCC flowpaths at air augmented rocket and rocket only modes is underway.

Turner, James E.↗
NASA NTRS2009

Transition Analysis for the HIFiRE-5 Vehicle

The Hypersonic International Flight Research and Experimentation (HIFiRE) 5 flight experiment by Air Force Research Laboratories and Australian Defense Science and Technology Organization is designed to provide in-flight boundary-layer transition data for a canonical 3D configuration at hypersonic Mach numbers. This paper outlines the progress, to date, on boundary layer stability analysis for the HIFiRE-5 flight configuration, as well as for selected test conditions from the wind tunnel experiments supporting the flight test. At flow conditions corresponding to the end of the test window, rather large values of linear amplification factor are predicted for both second mode (N>40) and crossflow (N>20) instabilities, strongly supporting the feasibility of first in-flight measurements of natural transition on a fully three-dimensional hypersonic configuration. Additional results highlight the rich mixture of instability mechanisms relevant to a large segment of the flight trajectory, as well as the effects of angle of attack and yaw angle on the predicted transition fronts for ground facility experiments at Mach 6.

Choudhari, Meelan M.↗
NASA NTRS2006

Rayleigh Scattering for Measuring Flow in a Nozzle Testing Facility

A molecular Rayleigh-scattering-based air-density measurement system was built in a large nozzle-and-engine-component test facility for surveying supersonic plumes from jet-engine exhaust. A molecular Rayleigh-scattering-based air-density measurement system was built in a large nozzle-and-enginecomponent test facility for surveying supersonic plumes from jet-engine exhaust

Gomez, Carlos R.↗