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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 253 records · Page 14
All sourcesNASA NTRSDOE OSTI
NASA NTRS1986

Wind tunnel tests of rotor blade sections with replications of ice formations accreted in hover

Full scale reproductions of ice accretions molded during the documentation of a hover test program were fabricated by means of epoxy castings and used for a wind tunnel test program. Surface static pressure distributions were recorded and used to evaluate lift and pitching moment increments while drag was determined by wake surveys. Through the range of the tests, corresponding to those conditions encountered in hover and in flat pitch, integration of the pressure distributions showed negligible changes in lift and in pitching moment, but the drag was significantly increased.

Lee, J. D.↗
NASA NTRS1961

Large-Scale Wind-Tunnel Tests of Exhaust Ingestion Due to Thrust Reversal on a Four-Engine Jet Transport during Ground Roll

Wind-tunnel tests have been conducted on a large-scale model of a swept-wing jet transport type airplane to study the factors affecting exhaust gas ingestion into the engine inlets when thrust reversal is used during ground roll. The model was equipped with four small jet engines mounted in nacelles beneath the wing. The tests included studies of both cascade and target type reversers. The data obtained included the free-stream velocity at the occurrence of exhaust gas ingestion in the outboard engine and the increment of drag due to thrust reversal for various modifications of thrust reverser configuration. Motion picture films of smoke flow studies were also obtained to supplement the data. The results show that the free-stream velocity at which ingestion occurred in the outboard engines could be reduced considerably, by simple modifications to the reversers, without reducing the effective drag due to reversed thrust.

Tolhurst, William H., Jr.↗
NASA NTRS2011

Correlating CFD Simulation with Wind Tunnel Test for the Full-Scale UH-60A Airloads Rotor

Data from the recent UH-60A Airloads Test in the National Full-Scale Aerodynamics Complex 40- by 80- Foot Wind Tunnel at NASA Ames Research Center are presented and compared to predictions computed by a loosely coupled Computational Fluid Dynamics (CFD)/Comprehensive analysis. Primary calculations model the rotor in free-air, but initial calculations are presented including a model of the tunnel test section. The conditions studied include a speed sweep at constant lift up to an advance ratio of 0.4 and a thrust sweep at constant speed into deep stall. Predictions show reasonable agreement with measurement for integrated performance indicators such as power and propulsive but occasionally deviate significantly. Detailed analysis of sectional airloads reveals good correlation in overall trends for normal force and pitching moment but pitching moment mean often differs. Chord force is frequently plagued by mean shifts and an overprediction of drag on the advancing side. Locations of significant aerodynamic phenomena are predicted accurately although the magnitude of individual events is often missed.

Romandr, Ethan↗
NASA NTRS2014

Low-Speed Dynamic Wind Tunnel Test Analysis of a Generic 53 Degree Swept UCAV Configuration With Controls

Several static and dynamic forced-motion wind tunnel tests have been conducted on a generic unmanned combat air vehicle (UCAV) configuration with a 53deg swept leading edge. These tests are part of an international research effort to assess and advance the state-of-art of computational fluid dynamics (CFD) methods to predict the static and dynamic stability and control characteristics for this type of configuration. This paper describes the dynamic forced motion data collected from two different models of this UCAV configuration as well as analysis of the control surface deflections on the dynamic forces and moments.

Vicroy, Dan D.↗
NASA NTRS2014

Experimental Aeroheating Study of Mid-L/D Entry Vehicle Geometries: NASA LaRC 20-Inch Mach 6 Air Tunnel Test 6966

Aeroheating data on mid lift-to-drag ratio entry vehicle configurations has been obtained through hypersonic wind tunnel testing. Vehicles of this class have been proposed for high-mass Mars missions, such as sample return and crewed exploration, for which the conventional sphere-cone entry vehicle geometries of previous Mars missions are insufficient. Several configurations were investigated, including elliptically-blunted cylinders with both circular and elliptical cross sections, biconic geometries based on launch vehicle dual-use shrouds, and parametrically-optimized analytic geometries. Testing was conducted at Mach 6 over a range of Reynolds numbers sufficient to generate laminar, transitional, and turbulent flow. Global aeroheating data were obtained using phosphor thermography. Both stream-wise and cross-flow transition occured on different configurations. Comparisons were made with laminar and turbulent computational predictions generated with an algebraic turbulence model. Predictions were generally in good agreement in regions of laminar or fully-turbulent flow; however for transitional cases, the lack of a transition onset prediction capability produced less accurate comparisons. The data obtained in this study are intended to be used for prelimary mission design studies and the development and validation of computational methods.

Hollis, Brian R.↗
NASA NTRS1973

The results of a low speed wind tunnel test to investigate the effects of installing refan JT8D engines on the McDonnell Douglas DC-9-30

A low speed wind tunnel test was conducted to assess the effects of the larger JT8D refan nacelles on the stability and control characteristics of the DC-9-30, with emphasis on the deep stall regime. Deep stall pitching moment and elevator hinge moment data, and low angle of attack tail-on and tail-off pitching moment data are presented. The refan nacelle was tested in conjunction with various pylons of reduced span relative to the production DC-9-30 pylon. Also, a horizontal tail that was larger than the production tail was tested. The data show that the refan installation has a small detrimental effect on the DC-9-30 deep stall recovery capability, that recovery characteristics are essentially independent of pylon span, and that the larger horizontal tail significantly increases recovery margins. The deep stall characteristics with the refan installation, within the range of pylon spans tested, are acceptable with no additional design changes anticipated.

Chrisenberry, H. E.↗
NASA NTRS1982

Configuration design studies and wind tunnel tests of an energy efficient transport with a high-aspect-ratio supercritical wing

The results of design studies and wind tunnel tests of high aspect ratio supercritical wings suitable for a medium range, narrow body transport aircraft flying near M=0.80 were presented. The basic characteristics of the wing design were derived from system studies of advanced transport aircraft where detailed structural and aerodynamic tradeoffs were used to determine the most optimum design from the standpoint of fuel usage and direct operating cost. These basic characteristics included wing area, aspect ratio, average thickness, and sweep. The detailed wing design was accomplished through application of previous test results and advanced computational transonic flow procedures. In addition to the basic wing/body development, considerable attention was directed to nacelle/plyon location effects, horizontal tail effects, and boundary layer transition effects. Results of these tests showed that the basic cruise performance objectives were met or exceeded.

Henne, P. A.↗
NASA NTRS2015

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

Characterization of flight vehicle unsteady aerodynamics is often studied via large scale wind tunnel testing. Boundary layer noise is measured by miniature pressure transducers installed in a model. Noise levels (2-5 dB ref. 20 μPa) can be induced when transducer is mounted out of flush with model outer surface. This effect must be minimized to accurately determine aerodynamically induced acoustic environments.

Herron, Andrew J.↗
NASA NTRS1971

On the use of freon-12 for increasing Reynolds number in wind-tunnel testing of three dimensional aircraft models at subcritical and supercritical Mach numbers

The aerodynamic suitability of Freon-12 for general wind-tunnel testing was investigated at low and high subsonic speeds. Static aerodynamic characteristics of two transport airplane models were determined from strain gage balance measurements in both air and Freon-12 at several Reynolds numbers. A low-speed high-lift configuration was evaluated at Mach number 0.25, and a high-speed cruise wing-fuselage combination was tested at Mach numbers up to 0.825. The data obtained in air and in Freon-12 agree well, even in stalled flow, until compressibility effects evidently become significant in air and in Freon-12 agree well, even in stalled flow, until compressibility effects evidently become significant in air.

Treon, S. L.↗
NASA NTRS1938

Wind-tunnel Tests of a 2-engine Airplane Model as a Preliminary Study of Flight Conditions Arising on the Failure of the Engine

Wind tunnel tests of a 15-foot-span model of a two-engine low wing transport airplane were made as a preliminary study of the emergency arising from the failure of one engine in flight. Two methods of reducing the initial yawing moment resulting from the failure of one engine were investigated and the equilibrium conditions were explored for two basic modes on one engine, one with zero angle of sideslip and the other with several degrees of sideslip. The added drag resulting from the unsymmetrical attitudes required for flight on one engine was determined for the model airplane. The effects of the application of power upon the stability, controllability, lift, and drag of the model airplane were measured. A dynamic pressure survey of the propeller slipstream was made in the neighborhood of the tail surfaces at three angles of attack. The added parasite drag of the model airplane resulting from the unfavorable conditions of flight on one engine was estimated. From 35 to 50 percent of this added drag was due to the drag of the dead engine propeller and the other 50 to 65 percent was due to the unsymmetrical attitude of the airplane. The mode of flight on one engine in which the angle of sideslip was zero was found to require less power than the mode in which the angle of sideslip was several degrees.

Hartman, Edwin P↗
NASA NTRS1934

Wind-tunnel tests on combinations of a wing with fixed auxiliary airfoils having various chords and profiles

This report presents the results of wind tunnel tests on various auxiliary airfoils having three different airfoil sections and several different chord lengths in combination with a Clark y model wing in a sufficient number of relative positions to determine the optimum with regard to certain criterions of aerodynamic performance. The airfoil sections included a symmetrical profile, one of medium camber, and a highly cambered one. The chord sizes of the auxiliary airfoils ranged from 7.5 to 25 percent of the chord of the main wing, and the span was equal to that of the main wing.

Weick, Fred E↗
NASA NTRS1949

Wind-Tunnel Tests of a 0.16-Scale Model of the Douglas MX-656 Airplane at High Subsonic Speeds: Stability and Control Characteristics - 1

Wind tunnel tests of the 0.16-scale Douglas MX-656 model were made at low and high subsonic Mach numbers to investigate the static longitudinal- and lateral stability characteristics. The tests shows that undesirable changes in longitudinal stability at the stall were apparently caused by an altered downwash pattern at the tail. The jettisonable nose fins were highly destabilizing. Compressibility effects for the test Mach numbers were not detrimental to the longitudinal- or lateral-stability characteristics.

Hamilton, William T.↗
NASA NTRS2003

Wind Tunnel Testing of Various Disk-Gap-Band Parachutes

Two Disk-Gap-Band model parachute designs were tested in the NASA Langley Transonic Dynamics Tunnel. The purposes of these tests were to determine the drag and static stability coefficients of these two model parachutes at various subsonic Mach numbers in support of the Mars Exploration Rover mission. The two model parachute designs were designated 1.6 Viking and MPF. These model parachute designs were chosen to investigate the tradeoff between drag and static stability. Each of the parachute designs was tested with models fabricated from MIL-C-7020 Type III or F-111 fabric. The reason for testing model parachutes fabricated with different fabrics was to evaluate the effect of fabric permeability on the drag and static stability coefficients. Several improvements over the Viking-era wind tunnel tests were implemented in the testing procedures and data analyses. Among these improvements were corrections for test fixture drag interference and blockage effects, and use of an improved test fixture for measuring static stability coefficients. The 1.6 Viking model parachutes had drag coefficients from 0.440 to 0.539, while the MPF model parachutes had drag coefficients from 0.363 to 0.428. The 1.6 Viking model parachutes had drag coefficients 18 to 22 percent higher than the MPF model parachute for equivalent fabric materials and test conditions. Model parachutes of the same design tested at the same conditions had drag coefficients approximately 11 to 15 percent higher when manufactured from F-111 fabric as compared to those fabricated from MIL-C-7020 Type III fabric. The lower fabric permeability of the F-111 fabric was the source of this difference. The MPF model parachutes had smaller absolute statically stable trim angles of attack as compared to the 1.6 Viking model parachutes for equivalent fabric materials and test conditions. This was attributed to the MPF model parachutes larger band height to nominal diameter ratio. For both designs, model parachutes fabricated from F-111 fabric had significantly greater statically stable absolute trim angles of attack at equivalent test conditions as compared to those fabricated from MILC-7020 Type III fabric. This reduction in static stability exhibited by model parachutes fabricated from F-111 fabric was attributed to the lower permeability of the F-111 fabric. The drag and static stability coefficient results were interpolated to obtain their values at Mars flight conditions using total porosity as the interpolating parameter.

Cruz, Juan R.↗
NASA NTRS2001

Flow-Visualization Techniques Used at High Speed by Configuration Aerodynamics Wind-Tunnel-Test Team

This paper summarizes a variety of optically based flow-visualization techniques used for high-speed research by the Configuration Aerodynamics Wind-Tunnel Test Team of the High-Speed Research Program during its tenure. The work of other national experts is included for completeness. Details of each technique with applications and status in various national wind tunnels are given.

Lamar, John E.↗
NASA NTRS1974

RSRA sixth scale wind tunnel test. Tabulated balance data, volume 2

Summaries are presented of all the force and moment data acquired during the RSRA Sixth Scale Wind Tunnel Test. These data include and supplement the data presented in curve form in previous reports. Each summary includes the model configuration, wing and empennage incidences and deflections, and recorded balance data. The first group of data in each summary presents the force and moment data in full scale parametric form, the dynamic pressure and velocity in the test section, and the powered nacelle fan speed. The second and third groups of data are the balance data in nondimensional coefficient form. The wind axis coefficient data corresponds to the parametric data divided by the wing area for forces and divided by the product of the wing area and wing span or mean aerodynamic chord for moments. The stability axis data resolves the wind axis data with respect to the angle of yaw.

Ruddell, A.↗
NASA NTRS1928

Wind Tunnel Tests on Autorotation and the "Flat Spin."

This report deals with the autorotational characteristics of certain differing wing systems as determined from wind tunnel tests made at the Langley Memorial Aeronautical Laboratory. The investigation was confined to autorotation about a fixed axis in the plane of symmetry and parallel to the wind direction. Analysis of the tests leads to the following conclusions: autorotation below 30 degree angle of attack is governed chiefly by wing profile, and above that angle by wing arrangement. The strip method of autorotation analysis gives uncertain results between maximum C subscript L and 35 degrees. The polar curve of a wing system, and to a lower degree of accuracy the polar of a complete airplane model are sufficient for direct determination of the limits of rotary instability, subject to strip method limitations. The results of the investigation indicate that in free flight a monoplane is incapable of flat spinning, whereas an unstaggered biplane has inherent flat-spinning tendencies. The difficulty of maintaining equilibrium in stalled flight is due primarily to rotary instability, a rapid change from stability to instability occurring as the angle of maximum lift is exceeded. (author)

Knight, Montgomery↗