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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 217 records · Page 12
All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

Study of Trailing Edge Mini Flaps Interactions on a Wind Tunnel Scale Aspect Ratio 13.5 Common Research Model

This paper presents a computation study of an aspect ratio 13.5 wind tunnel scale Common Research Model (CRM) with trailing edge mini-flaps. A vortex-lattice model of the CRM is developed with transonic small disturbance and integral boundary-layer corrections coupled with NASTRAN equivalent beam model for the rapid CRM aeroelastic simulation. The simulations are performed with individual trailing edge mini-flap deflected and a series of groups of two adjacent trailing edge mini-flaps deflected at the same time. A surrogate model is developed to estimate the aerodynamic interactions between the flaps. To validate the surrogate model, static aeroelastic simulations are performed using FUN3D CFD solver.

Interaction↗
NASA NTRS1978

Performance of single-stage axial-flow transonic compressor with rotor and stator aspect ratios of 1.19 and 1.26, respectively, and with design pressure ratio of 1.82

The overall and blade-element performance of a low-aspect-ratio transonic compressor stage is presented over the stable operating flow range at 70, 90, and 100 percent design speeds. At design speed the rotor and stage achieved peak efficiencies of 0.872 and 0.845 at pressure ratios of 1.875 and 1.842, respectively. The stage stall margin at design speed was 21.8 percent.

Reid, L.↗
NASA NTRS1980

Performance of single-stage axial-flow transonic compressor with rotor and stator aspect ratios of 1.19 and 1.26 respectively, and with design pressure ratio of 2.05

The overall and blade-element performances of a low-aspect-ratio transonic compressor stage are presented over the stable operating flow range for speeds from 50 to 100 percent of design. At design speed the rotor and stage achieved peak efficiencies of 0.876 and 0.840 at pressure ratios of 2.056 and 2.000, respectively. The stage stall margin at design speed was 10 percent.

Moore, R. D.↗
NASA NTRS1980

Transonic steady- and unsteady-pressure measurements on a high-aspect-ratio supercritical-wing model with oscillating control surfaces

A supercritical wing with an aspect ratio of 10.76 and with two trailing-edge oscillating control surfaces is described. The semispan wing is instrumented with 252 static orifices and 164 in situ dynamic-pressure gages for studying the effects of control-surface position and motion on steady- and unsteady-pressures at transonic speeds. Results from initial tests conducted in the Langley Transonic Dynamics Tunnel at two Reynolds numbers are presented in tabular form.

Sandford, M. C.↗
NASA NTRS1990

Effects of spoiler surfaces on the aeroelastic behavior of a low-aspect-ratio rectangular wing

The relationship between spoiler surfaces and flutter for a low-aspect ratio, rectangular wing is examined. The model design is a rigid wing mounted to the wing tunnel wall using a flexible, rectangular beam. The spoiler surface is composed of thin, rectangular aluminum plates that are vertically mounted to the wing surface. The effects of changes in spoiler geometry and location on wing flutter are studied using wind tunnel testing. It is determined that increases in spoiler height and weight cause an increase in the flutter dynamic pressure; however, variations in the location of the spoiler surfaces have little effect on flutter. Diagrams of the wing model are provided.

Cole, Stanley R.↗
NASA NTRS1976

Effect of coolant flow ejection on aerodynamic performance of low-aspect-ratio vanes. 1: Performance with coolant ejection holes plugged

The aerodynamic performance of a low aspect ratio turbine vane designed with coolant flow ejection holes on the vane surfaces was experimentally determined in a full-annular cascade with the coolant ejection holes plugged. The purpose was to establish a baseline for comparison with tests where flow is ejected from the vane surfaces. The vanes were tested over a mean-section ideal critical velocity ratio range of 0.64 to 0.98. This ideal critical velocity ratio corresponds to the vane inlet total to vane aftermixed static pressure ratio at the mean section. The variations in vane efficiency and aftermixed flow conditions with circumferential and radial position were obtained.

Haas, J. E.↗
NASA NTRS1981

Langley high-lift research on a high-aspect-ratio supercritical wing configuration

To determine the low speed performance characteristics of a representative high aspect ratio supercritical wing, two low speed jet transport models were fabricated. A 12-ft. span model was used for low Reynolds number tests in the Langley 4- by 7-Meter Tunnel and the second, a 7.5-ft. span model, was used for high Reynolds number tests in the Ames 12-foot Pressure Tunnel. A brief summary of the results of the tests of these two models is presented and comparisons are made between the data obtained on these two models and other similar models. Follow-on two and three dimensional research efforts related to the EET high-lift configurations are also presented and discussed.

Morgan, H. L., Jr.↗
NASA NTRS1985

The buffeting pressure field of a high-aspect-ratio swept wing

An experimental study of the fluctuating pressure field of a high-aspect-ratio swept transport-type wing model in transonic buffeting flow is examined. A high degree of similarity between the buffet-related features of the wing flowfield and 2-D transonic flowfields is indicated. At several spanwise stations, unsteady lift produced by the pressure fluctuations is evaluated and found to increase in intensity as local flow separation develops. Results show that although the wing does vibrate, the measured lift fluctuations are driving, and not driven by wing vibrations. The influence of tunnel-flow unsteadiness on the data is considered.

Roos, F. W.↗
NASA NTRS1989

Three dimensional flow field measurements of a 4:1 aspect ratio subsonic jet

Flow field measurements for a subsonic rectangular cold air jet with an aspect ratio of 4:1 (12.7 x 50.8 mm) at a Mach number of 0.09 and Re of 100,000 have been carried out using a three-dimensional laser Doppler anemometer system. Mean velocity measurements show that the jet width spreads more rapidly along the minor axis than along the major axis. The outward velocities, however, are not significantly different for the two axes, indicating the presence of enhanced mixing along the minor axis. The jet slowly changes from a rectangular jet to a circular jet as the flow progresses downstream.

Morrison, G. L.↗
NASA NTRS1991

The effects of winglets on low aspect ratio wings at supersonic Mach numbers

A computational study was conducted on two wings, of aspect ratios 1.244 and 1.865, each having 65 degree leading edge sweep angles, to determine the effects of nonplanar winglets at supersonic Mach numbers. A Mach number of 1.62 was selected as the design value. The winglets studied were parametrically varied in alignment, length, sweep, camber, thickness, and dihedral angle to determine which geometry had the best predicted performance. For the computational analysis, an available Euler marching technique was used. The results indicated that the possibility existed for wing-winglet geometries to equal the performance of wing-alone bodies in supersonic flows with both bodies having the same semispan. The first wing with winglet used NACA 1402 airfoils for the base wing and was shown to have lift-to-pressure drag ratios within 0.136 percent to 0.360 percent of the NACA 1402 wing-alone. The other base wing was a natural flow wing which was previously designed specifically for a Mach number of 1.62. The results obtained showed that the natural wing-alone had a slightly higher lift-to-pressure drag than the natural wing with winglets.

Keenan, James A.↗
NASA NTRS2017

Jet-Surface Interaction Noise from High-Aspect Ratio Nozzles: Test Summary

Noise and flow data have been acquired for a 16:1 aspect ratio rectangular nozzle exhausting near a simple surface at the NASA Glenn Research Center as part of an ongoing effort to understand, model, and predict the noise produced by current and future concept aircraft employing a tightly integrated engine airframe designs. The particular concept under consideration in this experiment is a blended-wing-body airframe powered by a series of electric fans exhausting through slot nozzle over an aft deck. The exhaust Mach number and surface length were parametrically varied during the test. Far-field noise data were acquired for all nozzle surface geometries and exhaust flow conditions. Phased-array noise source localization data and in-flow pressure data were also acquired for a subset of the isolated (no surface) and surface configurations; these measurements provide data that have proven useful for modeling the jet-surface interaction noise source and the surface effect on the jet-mixing noise in round jets. A summary of the nozzle surface geometry, flow conditions tested, and data collected are presented.

Propulsion Noise↗
NASA NTRS1979

Performance of two-stage fan having low-aspect-ratio first-stage rotor blading

The NASA two stage fan was tested with a low aspect ratio first stage rotor having no midspan dampers. At design speed the fan achieved an adiabatic design efficiency of 0.846, and peak efficiencies for the first stage and rotor of 0.870 and 0.906, respectively. Peak efficiency occurred very close to the stall line. In an attempt to improve stall margin, the fan was retested with circumferentially grooved casing treatment and with a series of stator blade resets. Results showed no improvement in stall margin with casing treatment but increased to 8 percent with stator blade reset.

Urasek, D. C.↗
NASA NTRS1983

Aileron effectiveness for a subsonic transport model with a high-aspect-ratio supercritical wing

Aileron effectiveness for a subsonc energy efficient transport (EET) model with a high aspect ratio supercritical wing was determined in the 8-foot transonic pressure tunnel. Data are presented for ailerons located at three positions along the wing span. The ailerons were designed as a preliminary active control concept with gust load alleviation, maneuver load alleviation, and flutter suppression systems. A linear variation of rolling moment coefficient with angle of attack for individual and multiple aileron deflections at Mach numbers up to 0.81 is indicated. For Mach numbers greater than 0.81, the rolling moment coefficient data become nonlinear with increasing angle of attack. At Mach numbers near the design value increased aileron effectiveness resulted from aft transition locations, which produced relatively thin boundary layers and greater effective aileron deflections. Individual aileron deflections on the right wing panel produced only small effects on yawing moment and side force coefficients.

Jacobs, P. F.↗