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At least 739 records · Page 41
All sourcesNASA NTRSDOE OSTI
NASA NTRS2008

Models of Lift and Drag Coefficients of Stalled and Unstalled Airfoils in Wind Turbines and Wind Tunnels

Equations are developed with which to calculate lift and drag coefficients along the spans of torsionally-stiff rotating airfoils of the type used in wind turbine rotors and wind tunnel fans, at angles of attack in both the unstalled and stalled aerodynamic regimes. Explicit adjustments are made for the effects of aspect ratio (length to chord width) and airfoil thickness ratio. Calculated lift and drag parameters are compared to measured parameters for 55 airfoil data sets including 585 test points. Mean deviation was found to be -0.4 percent and standard deviation was 4.8 percent. When the proposed equations were applied to the calculation of power from a stall-controlled wind turbine tested in a NASA wind tunnel, mean deviation from 54 data points was -1.3 percent and standard deviation was 4.0 percent. Pressure-rise calculations for a large wind tunnel fan deviated by 2.7 percent (mean) and 4.4 percent (standard). The assumption that a single set of lift and drag coefficient equations can represent the stalled aerodynamic behavior of a wide variety of airfoils was found to be satisfactory.

Spera, David A.↗
NASA NTRS1984

A wind tunnel database using RIM

Engineering data base development which has become increasingly widespread to industry with the availability of data management systems is examined. A large data base was developed for wind tunnel data and related model test information, using RIM as the data base manager. The arrangement of the wind tunnel data into the proper schema for the most efficient database utilization is discussed. The FORTRAN interface program of RIM is used extensively in the loading phases of the data base and by the users. Several examples to illustrate how the Wind Tunnel Data base might be searched for specific data items and test information using RIM are presented.

Wray, W. O., Jr.↗
NASA NTRS1979

Wind-tunnel fan noise reduction including effects of turning vanes on noise propagation

Fan silencers with low-pressure drop have been designed and installed in the NASA Ames 7- by 10-Foot Wind Tunnel No. 1. The silencers are composed of an absorbent lining flush with the wind-tunnel walls upstream of the fan and an absorbent splitter in the duct downstream of the fan. An acoustical insertion loss of 4 to 12 dB was measured between 100 and 1000 Hz. High-frequency performance of the silencers was reduced by test-section noise and by corner vanes which turned the sound waves so they followed the duct axis. Sound below 1.2 kHz diffracted around the 305-mm-long vanes and impinged on acoustically absorbent corner walls. Wind-tunnel flow losses are discussed.

Soderman, P. T.↗
NASA NTRS2015

The Acoustic Environment of the NASA Glenn 9- by 15-foot Low-Speed Wind Tunnel

The 9- by 15-Foot Low Speed Wind Tunnel is an acoustic testing facility with a long history of aircraft propulsion noise research. Due to interest in renovating the facility to support future testing of advanced quiet engine designs, a study was conducted to document the background noise level in the facility and investigate the sources of contaminating noise. The anechoic quality of the facility was also investigated using an interrupted noise method. The present report discusses these aspects of the noise environment in this facility.

Background Noise↗
NASA NTRSDate not supplied

Dynamic Mode Decomposition of Unsteady Pressure-Sensitive Paint Measurements for the NASA Unitary Plan Wind Tunnel Tests

This paper discusses the Dynamic Mode Decomposition (DMD) of the Unsteady Pressure-Sensitive Paint (uPSP) measurements, which were collected with four Phantom high-speed cameras at a constant sample frequency in the Ascent Transient Aerodynamics Test (ATAT) of the Space Launch System (SLS) Block 1 cargo vehicle with the Unitary Plan Wind Tunnel (UPWT) 11-by-11-foot Transonic Wind Tunnel in September 2019 at NASA Ames Research Center. The conventional DMD algorithm is based on the Singular Value Decomposition (SVD). For the data with zero mean, the DMD is equivalent to the Discrete Fourier Transform (DFT). Since the uPSP is mainly used to determine the unsteady property of the aerodynamic flow, the DMD of the uPSP measurements is implemented in two steps: (1) subtract the mean value from the uPSP measurement; (2) apply the Fast Fourier Transform (FFT) on the resulting data with zero mean. The DMD of the uPSP measurements with FFT has two advantages: (1) the FFT algorithm is well known for its computational efficiency, therefore, compared to the SVD-based DMD algorithm, the DMD with FFT reduces the computation time; (2) the DMD with FFT can be easily implemented in parallel processing. The DMD outputs were generated with the execution in parallel of a code in C, with libraries of FFTW for FFT and MPI/OpenMP for parallel processing, on the NASA Pleiades supercomputer. In this paper, the results of DMD of the uPSP measurements in the tests of Mach sweep runs of the SLS ATAT are presented, and the effectiveness of the DMD of the uPSP measurements in the diagnosis of the unsteady, aerodynamic phenomena is demonstrated. The work described in this paper is a part of NASA’s development of a new state-of-the-art uPSP capability in production wind tunnels. Funding for this research was provided by the NASA Aeroscience Evaluation and Test Capabilities Project.

Pressure-Sensitive Paint↗
NASA NTRS1987

Portable Test And Monitoring System For Wind-Tunnel Models

Portable system developed to test and monitor instrumentation used in wind-tunnel models. Self-contained and moves easily to model, either before or after model installed in wind tunnel. System is 44 1/2 in. high, 22 in. wide, and 17 in. deep and weighs 100 lb. Primary benefits realized with portable test and monitoring system associated with saving of time.

Poupard, Charles A.↗
NASA NTRS2001

Flight, Wind-Tunnel, and Computational Fluid Dynamics Comparison for Cranked Arrow Wing (F-16XL-1) at Subsonic and Transonic Speeds

Geometrical, flight, computational fluid dynamics (CFD), and wind-tunnel studies for the F-16XL-1 airplane are summarized over a wide range of test conditions. Details are as follows: (1) For geometry, the upper surface of the airplane and the numerical surface description compare reasonably well. (2) For flight, CFD, and wind-tunnel surface pressures, the comparisons are generally good at low angles of attack at both subsonic and transonic speeds, however, local differences are present. In addition, the shock location at transonic speeds from wind-tunnel pressure contours is near the aileron hinge line and generally is in correlative agreement with flight results. (3) For boundary layers, flight profiles were predicted reasonably well for attached flow and underneath the primary vortex but not for the secondary vortex. Flight data indicate the presence of an interaction of the secondary vortex system and the boundary layer and the boundary-layer measurements show the secondary vortex located more outboard than predicted. (4) Predicted and measured skin friction distributions showed qualitative agreement for a two vortex system. (5) Web-based data-extraction and computational-graphical tools have proven useful in expediting the preceding comparisons. (6) Data fusion has produced insightful results for a variety of visualization-based data sets.

John E Lamar↗
NASA NTRS1994

Dynamic response of NASA Rotor Test Apparatus and Sikorsky S-76 hub mounted in the 80- by 120-Foot Wind Tunnel

A shake test was conducted in the 80- by 120-Foot Wind Tunnel at NASA Ames Research Center, using the NASA Ames Rotor Test Apparatus (RTA) and the Sikorsky S-76 rotor hub. The primary objective of this shake test was to determine the modal properties of the RTA, the S-76 rotor hub, and the model support system installed in the wind tunnel. Random excitation was applied at the rotor hub, and vibration responses were measured using accelerometers mounted at various critical locations on the model and the model support system. Transfer functions were computed using the load cell data and the accelerometer responses. The transfer function data were used to compute the system modal parameters with the aid of modal analysis software.

Peterson, Randall L.↗
NASA NTRS1977

The effect of forward speed on J85 engine noise from suppressor nozzles as measured in the NASA-Ames 40- by 80-foot wind tunnel

An investigation to determine the effect of forward speed on the exhaust noise from a conical ejector nozzle and three suppressor nozzles mounted behind a J85 engine was performed in a 40- by 80-foot wind tunnel. The nozzles were tested at three engine power settings and at wind tunnel forward speeds up to 91 m/sec (300 ft/sec). In addition, outdoor static tests were conducted to determine (1) the differences between near field and far field measurements, (2) the effect of an airframe on the far field directivity of each nozzle, and (3) the relative suppression of each nozzle with respect to the baseline conical ejector nozzle. It was found that corrections to near field data are necessary to extrapolate to far field data and that the presence of the airframe changed the far field directivity as measured statically. The results show that the effect of forward speed was to reduce the noise from each nozzle more in the area of peak noise, but the change in forward quadrant noise was small or negligible. A comparison of wind tunnel data with available flight test data shows good agreement.

Atencio, A., Jr.↗
NASA NTRS1981

Operating manual holographic interferometry system for 2 x 2 foot transonic wind tunnel

A holographic interferometer system was installed in a 2X2 foot transonic wind tunnel. The system incorporates a modern, 10 pps, Nd:YAG pulsed laser which provides reliable operation and is easy to align. The spatial filtering requirements of the unstable resonator beam are described as well as the integration of the system into the existing Schieren system. A two plate holographic interferometer is used to reconstruct flow field data. For static wind tunnel models the single exposure holograms are recorded in the usual manner; however, for dynamic models such as oscillating airfoils, synchronous laser hologram recording is used.

Craig, J. E.↗
NASA NTRS1978

Wind-tunnel testing of VTOL and STOL aircraft

The basic concepts of wind-tunnel boundary interference are discussed and the development of the theory for VTOL-STOL aircraft is described. Features affecting the wall interference, such as wake roll-up, configuration differences, recirculation limits, and interference nonuniformity, are discussed. The effects of the level of correction on allowable model size are shown to be amenable to generalized presentation. Finally, experimental confirmation of wind-tunnel interference theory is presented for jet-flap, rotor, and fan-in-wing models.

Heyson, H. H.↗
NASA NTRS2004

Assessing Videogrammetry for Static Aeroelastic Testing of a Wind-Tunnel Model

The Videogrammetric Model Deformation (VMD) technique, developed at NASA Langley Research Center, was recently used to measure displacements and local surface angle changes on a static aeroelastic wind-tunnel model. The results were assessed for consistency, accuracy and usefulness. Vertical displacement measurements and surface angular deflections (derived from vertical displacements) taken at no-wind/no-load conditions were analyzed. For accuracy assessment, angular measurements were compared to those from a highly accurate accelerometer. Shewhart's Variables Control Charts were used in the assessment of consistency and uncertainty. Some bad data points were discovered, and it is shown that the measurement results at certain targets were more consistent than at other targets. Physical explanations for this lack of consistency have not been determined. However, overall the measurements were sufficiently accurate to be very useful in monitoring wind-tunnel model aeroelastic deformation and determining flexible stability and control derivatives. After a structural model component failed during a highly loaded condition, analysis of VMD data clearly indicated progressive structural deterioration as the wind-tunnel condition where failure occurred was approached. As a result, subsequent testing successfully incorporated near- real-time monitoring of VMD data in order to ensure structural integrity. The potential for higher levels of consistency and accuracy through the use of statistical quality control practices are discussed and recommended for future applications.

Spain, Charles V.↗
NASA NTRS1979

Wind-tunnel investigation of highly maneuverable supersonic V/STOL fighter

Results from the initial wind-tunnel test of a large-scale, highly maneuverable supersonic V/STOL fighter model in the Ames 40- by 80-foot wind tunnel are summarized. The STOL configuration which was tested combined upper surface blowing and spanwise blowing to improve the lift characteristics over a wide angle-of-attack range. A close-coupled canard was added to this configuration to create a highly maneuverable STOL aircraft. The 7.28 m (24 ft) span model was powered by two J-97 turbojet engines, each producing 9340 N (2200 lb) thrust at a pressure ratio of 2. With the nozzle flap and aileron set at 30 deg, the model produced lift coefficients greater than 4. The model was longitudinally unstable because of the forward canard position and because of the large body area of fuselage, strake, and nacelles forward of the center of gravity.

Falarski, M. D.↗
NASA NTRS1984

Correlations of supersonic boundary-layer transition on cones including effects of large axial variations in wind-tunnel noise

Transition data on sharp tip cones in two pilot low disturbance wind tunnels at Mach numbers of 3.5 and 5 were correlated in terms of noise parameters with data from several conventional wind tunnels and with data from supersonic flight tests on a transition cone. The noise parameters were developed to account for the large axial variations of the free stream noise and the very high frequency noise spectra that occurred in the low disturbance tunnels for some test conditions. The noise could be varied in these tunnels from high levels, approaching those in conventional tunnels, to extremely low levels. The correlations indicated that transition in the low disturbance tunnels was dominated by the local stream noise that was incident on the cone boundary layer unstream of the neutral stability point. The correlation results also suggested that high frequency components of the low disturbance tunnel noise spectra had significant effects on transition when the noise was incident on the boundary layer both upstream and downstream of the neutral stability point.

Chen, F. J.↗
NASA NTRSDate not supplied

Development and Assessment of a New Particle Image Velocimetry System in the NASA GRC 225 cm2 Wind Tunnel

As part of an ongoing campaign to provide detailed and thorough measurements in a two-dimensional impinging shock-wave/boundary-layer interaction (SWBLI), a custom implementation of the Particle Image Velocimetry (PIV) measurement technique in the NASA Glenn 225 cm2 Wind Tunnel was constructed. The flow seeding apparatus was devised to locally seed the region of interest in the overall flowfield to avoid coating the viewing window with seed material. The ability of the PIV system to provide small particles which faithfully track the underlying gas-phase flowfield was of key importance for the success of this system to make quality measurements in the SWBLI flowfield. In order to optimize and assess the performance of the particle seeding system, PIV measurements were made across a planar oblique shock wave and particle response was analyzed to quantify the particle lag effects due to the finite inertia of the seed droplets. Video recordings from a camera placed in the wind tunnel plenum tank were also used to discern the near-field behavior of the seeded plume and guide the design evolution of the seeding apparatus. The resulting optimized seeding configuration was applied to make PIV measurements in the wind tunnel test section boundary layer to assess the seed particle spatial coverage and concentration. This sample boundary layer dataset provides an analog to the eventual boundary layer measurements to be made in the axisymmetric test section under much more challenged viewing conditions.

PIV↗
NASA NTRSDate not supplied

Development and Assessment of a New Particle Image Velocimetry System in the NASA GRC 225 cm2 Wind Tunnel

As part of an ongoing campaign to provide detailed and thorough measurements in a two-dimensional impinging shock-wave/boundary-layer interaction (SWBLI), a custom implementation of the Particle Image Velocimetry (PIV) measurement technique in the NASA Glenn 225 cm2 Wind Tunnel was constructed. The flow seeding apparatus was devised to locally seed the region of interest in the overall flowfield to avoid coating the viewing window with seed material. The ability of the PIV system to provide small particles which faithfully track the underlying gas-phase flowfield was of key importance for the success of this system to make quality measurements in the SWBLI flowfield. In order to optimize and assess the performance of the particle seeding system, PIV measurements were made across a planar oblique shock wave and particle response was analyzed to quantify the particle lag effects due to the finite inertia of the seed droplets. Video recordings from a camera placed in the wind tunnel plenum tank were also used to discern the near-field behavior of the seeded plume and guide the design evolution of the seeding apparatus. The resulting optimized seeding configuration was applied to make PIV measurements in the wind tunnel test section boundary layer to assess the seed particle spatial coverage and concentration. This sample boundary layer dataset provides an analog to the eventual boundary layer measurements to be made in the axisymmetric test section under much more challenged viewing conditions.

PIV↗