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At least 541 records · Page 30
All sourcesNASA NTRSDOE OSTI
NASA NTRS2000

Photogrammetry Applied to Wind Tunnel Testing

In image-based measurements, quantitative image data must be mapped to three-dimensional object space. Analytical photogrammetric methods, which may be used to accomplish this task, are discussed from the viewpoint of experimental fluid dynamicists. The Direct Linear Transformation (DLT) for camera calibration, used in pressure sensitive paint, is summarized. An optimization method for camera calibration is developed that can be used to determine the camera calibration parameters, including those describing lens distortion, from a single image. Combined with the DLT method, this method allows a rapid and comprehensive in-situ camera calibration and therefore is particularly useful for quantitative flow visualization and other measurements such as model attitude and deformation in production wind tunnels. The paper also includes a brief description of typical photogrammetric applications to temperature- and pressure-sensitive paint measurements and model deformation measurements in wind tunnels.

Liu, Tian-Shu↗
NASA NTRS1996

Wind tunnel test results of a 1/8-scale fan-in-wing model

A 1/8-scale model of a fan-in-wing concept considered for development by Grumman Aerospace Corporation for the U.S. Army was tested in the Langley 14- by 22-Foot Subsonic Tunnel. Hover testing, which included height above a pressure-instrumented ground plane, angle of pitch, and angle of roll for a range of fan thrust, was conducted in a model preparation area near the tunnel. The air loads and surface pressures on the model were measured for several configurations in the model preparation area and in the tunnel. The major hover configuration change was varying the angles of the vanes attached to the exit of the fans for producing propulsive force. As the model height above the ground was decreased, there was a significant variation of thrust-removed normal force with constant fan speed. The greatest variation was generally for the height-to-fan exit diameter ratio of less than 2.5; the variation was reduced by deflecting fan exit flow outboard with the vanes. In the tunnel angles of pitch and sideslip, height above the tunnel floor, and wind speed were varied for a range of fan thrust and different vane angle configurations. Other configuration features such as flap deflections and tail incidence were evaluated as well. Though the V-tail empennage provided an increase in static longitudinal stability, the total model configuration remained unstable.

Wilson, John C.↗
NASA NTRS1972

Transonic wind tunnel tests of a .015 scale space shuttle booster model

Details of the test program, an outline of the test procedure, and results of the investigation in tabular form are presented. The base drag reduction potential was determined at cruise conditions (M = .40) of a modified base region, with and without base bleed flow. Transonic effects and cruise control effectiveness of the new base region were also obtained. The tests were performed at Mach numbers from 0.40 to 1.10, at angles of attack from -4 to +20 degrees and at angles of sideslip from -6 to +6 degrees, all at a Reynolds number of 6.56 million per meter (2.0 million per foot).

Dennis, R. J.↗
NASA NTRS1973

Wind tunnel tests of a 20 inch diameter 1.15 pressure ratio fan engine model

Aerodynamic and acoustic measurements at a typical STOL aircraft takeoff and landing velocity demonstrated that a 1.35 inlet lip area contraction ratio was superior to a 1.26 ratio at high nacelle incidence angles. Reverse thrust, obtained with a variable pitch rotor, was lower at the landing velocity, and the noise level higher, than at the static condition. High speed tests showed that, for the design cruise Mach number of 0.75, internal losses and external drag were 27 percent of the ideal fan net thrust, and propulsive efficiency was estimated to be 59 percent for an 85 percent efficient fan stage.

Wesoky, H. L.↗
NASA NTRS1974

Low speed wind tunnel tests of a 1/9-scale model of a variable-sweep advanced supersonic transport

Tests have been conducted in the Langley full-scale tunnel to determine the aerodynamic characteristics of a 1/9-scale variable-sweep advanced supersonic transport configuration. The model configurations investigated were the basic unflapped arrangement, and a takeoff and landing flap arrangement with several strake leading edge flow control devices. The tests were conducted for an angle-of-attack range from about minus 5 deg to 36 deg and a sideslip range from minus 5 deg to 10 deg. The tests were conducted for a range of Reynolds number from 3.92 million to 5.95 million corresponding to test velocities of about 54.5 knots and 81.7 knots, respectively.

Mclemore, H. C.↗
NASA NTRS1974

Data correlation and analysis of arc tunnel and wind tunnel tests of RSI joints and gaps. Volume 1: Technical report

Heat transfer data measured in gaps typical of those under consideration for joints in space shuttle reusable surface insulation protection systems have been assimilated, analyzed and correlated. The data were obtained in four NASA facilities. Several types of gaps were investigated with emphasis on simple butt joints. Gap widths ranged from 0.07 to 0.7 cm and depths ranged from 1 to 6 cm. Laminar, transitional and turbulent boundary layer flows over the gap opening were investigated. Three-dimensional heating variations were observed within gaps in the absence of external flow pressure gradients. Heat transfer correlation equations were obtained for several of the tests. Thermal protection system performance with and without gaps was compared for a representative shuttle entry trajectory.

Christensen, H. E.↗
NASA NTRS1974

Wind Tunnel Test of Low Boom Equivalent Body at Mach 4

A body of revolution, representing the equivalent area distribution of a low boom aircraft design cruising at 24,384 meters at a Mach number of 4, was tested to determine whether its theoretical sonic boom characteristics could be experimentally verified. Results indicate that the essential features of the ground signature are well predicted.

Howard, F. G.↗
NASA NTRS1973

Wind tunnel tests of a 20 inch diameter 1.15 pressure ratio fan engine model

Aerodynamic and acoustic measurements at a typical STOL aircraft takeoff and landing velocity demonstrated that a 1.35 inlet lip area contraction ratio was superior to a 1.26 ratio at high nacelle incidence angles. Reverse thrust, obtained with a variable pitch rotor, was lower at the landing velocity, and the noise level higher, than at the static condition. High speed tests showed that, for the design cruise Mach number of 0.75, internal losses and external drag were 27 per cent of the ideal fan net thrust, and propulsive efficiency was estimated to be 59 per cent for an 85 per cent efficient fan stage. For comparison, a similar 1.55 pressure ratio fan system would have a propulsive efficiency of 62 per cent.

Wesoky, H. L.↗