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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 NTRS2008

Wind Tunnel Testing of Powered Lift, All-Wing STOL Model

Short take-off and landing (STOL) systems can offer significant capabilities to warfighters and, for civil operators thriving on maximizing efficiencies they can improve airspace use while containing noise within airport environments. In order to provide data for next generation systems, a wind tunnel test of an all-wing cruise efficient, short take-off and landing (CE STOL) configuration was conducted in the National Aeronautics and Space Administration (NASA) Langley Research Center (LaRC) 14- by 22-foot Subsonic Wind Tunnel. The test s purpose was to mature the aerodynamic aspects of an integrated powered lift system within an advanced mobility configuration capable of CE STOL. The full-span model made use of steady flap blowing and a lifting centerbody to achieve high lift coefficients. The test occurred during April through June of 2007 and included objectives for advancing the state-of-the-art of powered lift testing through gathering force and moment data, on-body pressure data, and off-body flow field measurements during automatically controlled blowing conditions. Data were obtained for variations in model configuration, angles of attack and sideslip, blowing coefficient, and height above ground. The database produced by this effort is being used to advance design techniques and computational tools for developing systems with integrated powered lift technologies.

Collins, Scott W.↗
NASA NTRS2015

Reducing the Effect of Transducer Mount Induced Noise on Aeroacoustic Wind Tunnel Testing Data with a New Transducer Mount Design

Flight vehicle aeroacoustic environments induced during transonic and supersonic flight are usually predicted by subscale wind tunnel testing utilizing high frequency miniature pressure transducers. In order to minimize noise induced by the measurement itself, transducer flush mounting with the model surface is very important. The National Aeronautics and Space Administration (NASA) has accomplished flushness in recent testing campaigns via use of a transducer holder that can be machined and sanded. A single hole in the holder allows the flow medium to interact with the transducer diaphragm. Noise is induced by the resulting cavity however, and is a challenge to remove in post-processing. A new holder design has been developed that minimizes the effects of this transducer mount induced noise (XMIN) by reducing the resonance amplitude or increasing its resonance frequency beyond the range of interest. This paper describes a test conducted at the NASA/George C. Marshall Space Flight Center Trisonic Wind Tunnel intended to verify the effectiveness of this design. The results from this test show that this new transducer holder design does significantly reduce the influence of XMIN on measured fluctuating pressure levels without degrading a transducer's ability to accurately measure the noise external to the model.

Herron, Andrew J.↗
NASA NTRS1975

Low Speed Wind Tunnel Tests on a One-Seventh Scale Model of the H.126 Jet Flap Aircraft

Low speed wind tunnel tests were performed on a one-seventh scale model of the British H.126 jet flap research aircraft over a range of jet momentum coefficients. The primary objective was to compare model aerodynamic characteristics with those of the aircraft, with the intent to provide preliminary data needed towards establishing small-to-full scale correlating techniques on jet flap V/STOL aircraft configurations. Lift and drag coefficients from the model and aircraft tests were found to be in reasonable agreement. The pitching moment coefficient and trim condition correlation was poor. A secondary objective was to evaluate a modified thrust nozzle having thrust reversal capability. The results showed there was a considerable loss of lift in the reverse thrust operational mode because of increased nozzle-wing flow interference. A comparison between the model simulated H.126 wing jet efflux and the model uniform pressure distribution wing jet efflux indicated no more than 5% loss in weight flow rate.

Laub, G. H.↗
NASA NTRS1936

Wind-Tunnel Tests of a 10-foot-diameter Gyroplane Rotor

This report presents the results of wind-tunnel tests on a model gyroplane rotor 10 feet in diameter. The rotor blades had zero sweepback and zero offset; the hub contained a feathering mechanism that provided control of the rotor rolling moment, but not of the pitching moment. The rotor was tested with 4 blades and with 2 blades. The entire useful range of pitch settings and tip-speed ratios was investigated including the phase of operation in which the rotor turned very slowly, or idled.

Wheatley, John B↗
NASA NTRS2012

Forced Oscillation Wind Tunnel Testing for FASER Flight Research Aircraft

As unmanned air vehicles (UAVs) continue to expand their flight envelopes into areas of high angular rate and high angle of attack, modeling the complex unsteady aerodynamics for simulation in these regimes has become more difficult using traditional methods. The goal of this experiment was to improve the current six degree-of-freedom aerodynamic model of a small UAV by replacing the analytically derived damping derivatives with experimentally derived values. The UAV is named the Free-flying Aircraft for Sub-scale Experimental Research, FASER, and was tested in the NASA Langley Research Center 12- Foot Low-Speed Tunnel. The forced oscillation wind tunnel test technique was used to measure damping in the roll and yaw axes. By imparting a variety of sinusoidal motions, the effects of non-dimensional angular rate and reduced frequency were examined over a large range of angle of attack and side-slip combinations. Tests were performed at angles of attack from -5 to 40 degrees, sideslip angles of -30 to 30 degrees, oscillation amplitudes from 5 to 30 degrees, and reduced frequencies from 0.010 to 0.133. Additionally, the effect of aileron or elevator deflection on the damping coefficients was examined. Comparisons are made of two different data reduction methods used to obtain the damping derivatives. The results show that the damping derivatives are mainly a function of angle of attack and have dependence on the non-dimensional rate and reduced frequency only in the stall/post-stall regime

Hoe, Garrison↗
NASA NTRS2017

High Lift Common Research Model for Wind Tunnel Testing: An Active Flow Control Perspective

This paper provides an overview of a research and development effort sponsored by the NASA Advanced Air Transport Technology Project to achieve the required high-lift performance using active flow control (AFC) on simple hinged flaps while reducing the cruise drag associated with the external mechanisms on slotted flaps of a generic modern transport aircraft. The removal of the external fairings for the Fowler flap mechanism could help to reduce drag by 3.3 counts. The main challenge is to develop an AFC system that can provide the necessary lift recovery on a simple hinged flap high-lift system while using the limited pneumatic power available on the aircraft. Innovative low-power AFC concepts will be investigated in the flap shoulder region. The AFC concepts being explored include steady blowing and unsteady blowing operating in the spatial and/or temporal domain. Both conventional and AFC-enabled high-lift configurations were designed for the current effort. The high-lift configurations share the cruise geometry that is based on the NASA Common Research Model, and therefore, are also open geometries. A 10%-scale High Lift Common Research Model (HL-CRM) is being designed for testing at the NASA Langley Research Center 14- by 22-Foot Subsonic Tunnel during fiscal year 2018. The overall project plan, status, HL-CRM configurations, and AFC objectives for the wind tunnel test are described.

Lin, John C.↗
NASA NTRS1973

The results of a high-speed wind tunnel test to investigate the effects of the NASA refan JT8D engine nacelles on the stability and control characteristics of the Boeing 727 airplane

A high speed wind tunnel test was conducted to investigate the effects of the NASA Refan JT8D engine nacelles on the stability and control characteristics of the Boeing 727 airplane. The test was performed at the Calspan Corporation 8x8 ft. (2.44x2.44 m.) transonic wind tunnel. Both the 727-200 and -100 models were tested. A small nose-down pitching moment increment and a slight increase in longitudinal stability were noted due to the Refan nacelles. The directional stability of the 727-200 airplane increased up to 10 percent. A smaller improvement was observed on the 727-100 model. In general, the high speed stability and control characteristics of the basic airplane are not significantly altered by the Refan nacelle installation.

Kupcis, E. A.↗
NASA NTRS1985

Icing tunnel tests of Electro-Impulse De-Icing of an engine inlet and high-speed wings

A brief review is given of four earlier tests in the NASA Lewis Icing Research Tunnel and of flight tests in NASA's Icing Research Aircraft and in a Cessna 206 airplane. Details are given of recent icing tunnel tests of thicker-skinned wings, a Gates Learjet, a composite leading edge, and a Boeing 767, and of a Falcon Fanjet engine inlet. These were tested at speeds from 87 to 220 knots, air temperatures from -2 to -15 C, LWC values of 0.6 to 2.4 grams/cu meter, and median droplet diameters from 12 to 20 microns. Energy requirements are reported, as well as conclusions from comparisons of several Electro-Impulse De-Icing coil system designs. Fundamental studies of the structural dynamics and ice shedding of a 12.7 cm (5 inch) diameter semicylinder are described. Some potential problem areas are discussed: fatigue of skin and coil mountings, system weight and cost, electro-magnetic interference and noise.

Zumwalt, G. W.↗
NASA NTRS1972

Wind-tunnel tests of a single stage to orbit space shuttle at Mach numbers of 2.60, 3.85, and 4.64

Wind tunnel tests were conducted to determine the static aerodynamic characteristics of a single-stage-to-orbit space shuttle model at Mach numbers of 2.60, 3.85, and 4.64. Test parameters included various payloads for the launch configurations and various corner radii for the entry configurations. The test results for the launch configurations generally indicated that a decrease in payload size resulted in a marked increase in axial force. Decreasing the corner radii of the entry configuration led to large increases in axial force. The results also indicated that the presence of an engine door on the entry configuration caused a measurable positive increment in pitching moment. Changing the entry configuration afterbody geometry from a biconic shape to a conic shape increased the normal-force and stability level.

Graves, E. B.↗
NASA NTRS1982

Wind tunnel tests of high-lift systems for advanced transports using high-aspect-ratio supercritical wings

The wind tunnel testing of an advanced technology high lift system for a wide body and a narrow body transport incorporating high aspect ratio supercritical wings is described. This testing has added to the very limited low speed high Reynolds number data base for this class or aircraft. The experimental results include the effects on low speed aerodynamic characteristics of various leading and trailing edge devices, nacelles and pylons, ailerons, and spoilers, and the effects of Mach and Reynolds numbers.

Allen, J. B.↗