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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 199 records · Page 11
All sourcesNASA NTRSDOE OSTI
NASA NTRS1985

Flight testing the fixed-wing configuration of the Rotor Systems Research Aircraft (RSRA)

The Rotor Systems Research Aircraft (RSRA) is a unique research aircraft designed to flight test advanced helicopter rotor system. Its principal flight test configuration is as a compound helicopter. The fixed wing configuration of the RSRA was primarily considered an energy fly-home mode in the event it became necessary to sever an unstable rotor system in flight. While it had always been planned to flight test the fixed wing configuration, the selection of the RSRA as the flight test bed for the X-wing rotor accelerated this schedule. This paper discusses the build-up to, and the test of, the RSRA fixed wing configuration. It is written primarily from the test pilot's perspective.

Hall, G. W.↗
NASA NTRS1987

Performance and loads data from a hover test of a 0.658-scale V-22 rotor and wing

A hover test of a 0.658-scale model of a V-22 rotor and wing was conducted at the Outdoor Aerodynamic Research Facility at Ames Research Center. The primary objectives of the test were to obtain accurate measurements of the hover performance of the rotor system, and to measure the aerodynamic interactions between the rotor and wing. Data were acquired for rotor tip Mach numbers ranging from 0.1 to 0.73. This report presents data on rotor performance, rotor-wake downwash velocities, rotor system loads, wing forces and moments, and wing surface pressures.

Felker, Fort F.↗
NASA NTRS2005

Flight Test of the F/A-18 Active Aeroelastic Wing Airplane

Successful flight-testing of the Active Aeroelastic Wing airplane was completed in March 2005. This program, which started in 1996, was a joint activity sponsored by NASA, Air Force Research Laboratory, and industry contractors. The test program contained two flight test phases conducted in early 2003 and early 2005. During the first phase of flight test, aerodynamic models and load models of the wing control surfaces and wing structure were developed. Design teams built new research control laws for the Active Aeroelastic Wing airplane using these flight-validated models; and throughout the final phase of flight test, these new control laws were demonstrated. The control laws were designed to optimize strategies for moving the wing control surfaces to maximize roll rates in the transonic and supersonic flight regimes. Control surface hinge moments and wing loads were constrained to remain within hydraulic and load limits. This paper describes briefly the flight control system architecture as well as the design approach used by Active Aeroelastic Wing project engineers to develop flight control system gains. Additionally, this paper presents flight test techniques and comparison between flight test results and predictions.

Clarke, Robert↗
NASA NTRS2005

Flight Test of the F/A-18 Active Aeroelastic Wing Airplane

Successful flight-testing of the Active Aeroelastic Wing airplane was completed in March 2005. This program, which started in 1996, was a joint activity sponsored by NASA, Air Force Research Laboratory, and industry contractors. The test program contained two flight test phases conducted in early 2003 and early 2005. During the first phase of flight test, aerodynamic models and load models of the wing control surfaces and wing structure were developed. Design teams built new research control laws for the Active Aeroelastic Wing airplane using these flight-validated models; and throughout the final phase of flight test, these new control laws were demonstrated. The control laws were designed to optimize strategies for moving the wing control surfaces to maximize roll rates in the transonic and supersonic flight regimes. Control surface hinge moments and wing loads were constrained to remain within hydraulic and load limits. This paper describes briefly the flight control system architecture as well as the design approach used by Active Aeroelastic Wing project engineers to develop flight control system gains. Additionally, this paper presents flight test techniques and comparison between flight test results and predictions.

Clarke, Robert↗
NASA NTRS1962

Summary of Flight-Test Results of the VZ-2 Tilt-Wing Aircraft

Flight-test information gained from a tilt-wing research aircraft tested at the Langley Research Center has shown that design problems exist in such fields as low-speed stability and control, handling qualities, and flow separation during transition. The control power in the near-hovering configuration was considered by the pilots to be inadequate in yaw, marginal in pitch, and excessive in roll. Solutions for some of the design problems are indicated; for example, the addition of a leading-edge droop to the wing in an attempt to delay flow separation resulted in such significantly improved handling qualities in the transition range that an additional descent capability of 1,100 feet per minute was obtained.

Pegg, Robert J.↗
NASA NTRS2001

DARPA/AFRL/NASA Smart Wing Second Wind Tunnel Test Results

To quantify the benefits of smart materials and structures adaptive wing technology, Northrop Grumman Corp. (NGC) built and tested two 16% scale wind tunnel models (a conventional and a "smart" model) of a fighter/attack aircraft under the DARPA/AFRL/NASA Smart Materials and Structures Development - Smart Wing Phase 1. Performance gains quantified included increased pitching moment (C(sub M)), increased rolling moment (C(subl)) and improved pressure distribution. The benefits were obtained for hingeless, contoured trailing edge control surfaces with embedded shape memory alloy (SMA) wires and spanwise wing twist effected by SMA torque tube mechanisms, compared to conventional hinged control surfaces. This paper presents an overview of the results from the second wind tunnel test performed at the NASA Langley Research Center s (LaRC) 16ft Transonic Dynamic Tunnel (TDT) in June 1998. Successful results obtained were: 1) 5 degrees of spanwise twist and 8-12% increase in rolling moment utilizing a single SMA torque tube, 2) 12 degrees of deflection, and 10% increase in rolling moment due to hingeless, contoured aileron, and 3) demonstration of optical techniques for measuring spanwise twist and deflected shape.

Scherer, L. B.↗
NASA NTRS1941

Tests of Round and Flat Spoilers on a Tapered Wing in the NACA 19-Foot Pressure Wind Tunnel

Several arrangements of round and flat spanwise spoilers attached to the upper surface of a tapered wing were tested in the NACA 19-foot pressure wind tunnel to determine the most effective type, location, and size of spoiler necessary to reduce greatly the lift on the wings of large flying boats when moored. The effect of the various spoilers on the lift, the drag, and the pitching-moment characteristics of the tapered wing was measured over a range of angles of attack from zero to maximum lift. The most effective type of spoiler was found to be the flat type with no space between it and the wing surface. The chordwise location of such a spoiler was not critical within the range investigated, from 5 to 20 percent of the wing chord from the leading edge.

Wenzinger, Carl J↗
NASA NTRS1933

Wind-tunnel tests of a Clark Y wing with a narrow auxiliary airfoil in different positions

Aerodynamic force tests were made on a combination of a Clark Y wing and a narrow auxiliary airfoil to find the best location of the auxiliary airfoil with respect to the main wing. The auxiliary was a highly cambered airfoil of medium thickness having a chord 14.5 per cent that of the main wing. It was tested in 141 different positions ahead of, above, and behind the nose portion of the main wing, the range of the test points being extended until the best aerodynamic conditions were covered. A range of positions was found in which the combination of main wing and auxiliary gave substantially greater aerodynamic efficiency and higher maximum lift coefficients (based on total area) than the main Clark Y wing alone. In the optimum position tested, considering both the maximum lift and the speed-range ratio, the combination of main wing and auxiliary gave an increase in the maximum lift coefficient of 32 per cent together with an increase in the ratio of 21 per cent of the respective values for the main Clark Y wing alone.

Weick, Fred E↗
NASA NTRS1929

Full Scale Investigation of the Drag of a Wing Radiator

Tests were made on the 1927 Williams racer in order to determine the effect of the wing radiator on the airfoil characteristics. It was found that the radiator doubled the minimum drag of the portion of the wing it covered, and also reduced the lift somewhat.

Weick, Fred E↗
NASA NTRS1938

Wind-tunnel tests of a Clark Y wing having split flaps with gaps

Tests were made of the Clark Y wing having split flaps with a gap between the flap and the lower surface of the wing. Lift, drag, and pitching moments were measured for the wing with three different sizes of flap. It was found that any gap between the flap and the wing reduced the lift, the drag, and the pitching moments, but that the center-of-pressure movement and the ratio of lift to drag were little affected.

Wenzinger, Carl J↗
NASA NTRS1974

Influence of configuration details on the subsonic characteristics of a space shuttle orbiter design

An investigation was conducted in the Langley low-turbulence pressure tunnel of a model of a space shuttle orbiter design in order to determine the influence of minor configuration geometric details on the aerodynamic characteristics at subsonic speeds. A plane wing was tested with a small planform fillet; a twisted wing was tested with both a small and a large planform fillet. Tailored attitude-control propulsion-system wing-tip and body pods, trisegmented elevons, and canopy effects were also investigated. The tests were conducted at angles of attack from -3 deg to 24 deg for sideslip angles of 0 deg and 6 deg and at a Mach number of 0.25.

Decker, J. P.↗
NASA NTRS1973

Design and fabrication of an augmentor wing model for acoustic tests

The design and fabrication of a full-scale section of an augmentor wing to be used for acoustic testing at the Lewis Research Center are discussed. This hardware will be used primarily to investigate scaling effects of acoustic data obtained during the Boeing-run model tests. Typical model test data is shown in the report, together with predictions on both performance and acoustics that can be expected from the full-scale section to be built. Areas covered include: the aerodynamic and acoustic criteria of the flap system and nozzles, detailed discussion of the hardware, test system operation procedure, and stress analysis of the entire test system.

Jackson, J.↗
NASA NTRS1979

Wind-tunnel investigation of the thrust augmentor performance of a large-scale swept wing model

Tests were made in the Ames 40- by 80-foot wind tunnel to determine the forward speed effects on wing-mounted thrust augmentors. The large-scale model was powered by the compressor output of J-85 driven viper compressors. The flap settings used were 15 deg and 30 deg with 0 deg, 15 deg, and 30 deg aileron settings. The maximum duct pressure, and wind tunnel dynamic pressure were 66 cmHg (26 in Hg) and 1190 N/sq m (25 lb/sq ft), respectively. All tests were made at zero sideslip. Test results are presented without analysis.

Koenig, D. G.↗
NASA NTRS1982

Effect of nacelles on aerodynamic characteristics of an executive-jet model with simulated, partial-chord, laminar-flow-control wing glove

Tests were conducted in the Langley High-Speed 7- by 10-Foot Tunnel using a 1/10-scale model of an executive jet to examine the effects of the nacelles on the wing pressures and model longitudinal aerodynamic characteristics. For the present investigation, each wing panel was modified with a simulated, partial-chord, laminar-flow-control glove. Horizontal-tail effects were also briefly examined. The tests covered a range of Mach numbers from 0.40 to 0.82 and lift coefficients from 0.20 to 0.55. Oil-flow photographs of the wing at selected conditions are included.

Campbell, R. L.↗
NASA NTRS1997

Flight Wing Surface Pressure and Boundary-Layer Data Report from the F-111 Smooth Variable-Camber Supercritical Mission Adaptive Wing

Flight tests were conducted using the advanced fighter technology integration F-111 (AFTI/F-111) aircraft modified with a variable-sweep supercritical mission adaptive wing (MAW). The MAW leading- and trailing-edge variable-camber surfaces were deflected in flight to provide a near-ideal wing camber shape for the flight condition. The MAW features smooth, flexible upper surfaces and fully enclosed lower surfaces, which distinguishes it from conventional flaps that have discontinuous surfaces and exposed or semi-exposed mechanisms. Upper and lower surface wing pressure distributions were measured along four streamwise rows on the right wing for cruise, maneuvering, and landing configurations. Boundary-layer measurements were obtained near the trailing edge for one of the rows. Cruise and maneuvering wing leading-edge sweeps were 26 deg for Mach numbers less than 1 and 45 deg or 58 deg for Mach numbers greater than 1. The landing wing sweep was 9 deg or 16 deg. Mach numbers ranged from 0.27 to 1.41, angles of attack from 2 deg to 13 deg, and Reynolds number per unit foot from 1.4 x 10(exp 6) to 6.5 x 10(exp 6). Leading-edge cambers ranged from O deg to 20 deg down, and trailing-edge cambers ranged from 1 deg up to 19 deg down. Wing deflection data for a Mach number of 0.85 are shown for three cambers. Wing pressure and boundary-layer data are given. Selected data comparisons are shown. Measured wing coordinates are given for three streamwise semispan locations for cruise camber and one spanwise location for maneuver camber.

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