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At least 577 records · Page 32
All sourcesNASA NTRSDOE OSTI
NASA NTRS1994

Application of fuzzy logic to the control of wind tunnel settling chamber temperature

The application of Fuzzy Logic Controllers (FLC's) to the control of nonlinear processes, typically controlled by a human operator, is a topic of much study. Recent application of a microprocessor-based FLC to the control of temperature processes in several wind tunnels has proven to be very successful. The control of temperature processes in the wind tunnels requires the ability to monitor temperature feedback from several points and to accommodate varying operating conditions in the wind tunnels. The FLC has an intuitive and easily configurable structure which incorporates the flexibility required to have such an ability. The design and implementation of the FLC is presented along with process data from the wind tunnels under automatic control.

Gwaltney, David A.↗
NASA NTRS1984

Cryogenic wind tunnel technology. A way to measurement at higher Reynolds numbers

The goals, design, problems, and value of cryogenic transonic wind tunnels being developed in Europe are discussed. The disadvantages inherent in low-Reynolds-number (Re) wind tunnel simulations of aircraft flight at high Re are reviewed, and the cryogenic tunnel is shown to be the most practical method to achieve high Re. The design proposed for the European Transonic Wind tunnel (ETW) is presented: parameters include cross section. DISPLAY 83A46484/2 = 4 sq m, operating pressure = 5 bar, temperature = 110-120 K, maximum Re = 40 x 10 to the 6th, liquid N2 consumption = 40,000 metric tons/year, and power = 39,5 MW. The smaller Cologne subsonic tunnel being adapted to cryogenic use for preliminary studies is described. Problems of configuration, materials, and liquid N2 evaporation and handling and the research underway to solve them are outlined. The benefits to be gained by the construction of these costly installations are seen more in applied aerodynamics than in basic research in fluid physics. The need for parallel development of both high Re tunnels and computers capable of performing high-Re numerical analysis is stressed.

Beck, J. W.↗
NASA NTRSDate not supplied

Wind Tunnel Testing Techniques for a Tandem Tilt-Wing, Distributed Electric Propulsion VTOL Aircraft

Urban Air Mobility (UAM) vertical takeoff and landing (VTOL) aircraftdesigns frequently include multiple distributed propulsors, complex wing-propulsor aerodynamics, significant airframe configuration changes during normal flight operations, and no historical database regarding the best ways to transition between vertical and horizontal flight. This paper describes the methodology used for wind tunnel testing of the Langley Aerodrome No. 8 (LA-8) in the NASA Langley 12-Foot Low-SpeedTunnel during multiple test entries in 2019 and2020. The LA-8 is a tandem tilt-wing aircraft with 4 motor-propellerunits and 4 control surfaces distributed across each wing, plus an inverted V-tail with 2 ruddervators on the fuselage. An initial tunnel entryused one-factor-at-a-time (OFAT) testing to (1) define candidate trimmed transition corridors between vertical and horizontal flight, (2) assess whether there was adequate control authority, and (3) define appropriatetest factor ranges for subsequent design of experiment (DOE) wind tunnel testing. The total number of independent variables for these wind tunnel tests (23) made DOE testing an efficient option for assessing the large number of potential interactions associated with the LA-8. The general advantages and disadvantages of OFAT and DOE wind tunnel testing techniques are also discussed –along with the benefits of a combined approach

wind tunnel research techniques↗
NASA NTRS1997

Static Aeroelastic Analysis of Transonic Wind Tunnel Models Using Finite Element Methods

A computational method for accurately predicting the static aeroelastic deformations of typical transonic transport wind tunnel models is described. The method utilizes a finite element method (FEM) for predicting the deformations. Extensive calibration/validation of this method was carried out using a novel wind-off wind tunnel model static loading experiment and wind-on optical wing twist measurements obtained during a recent wind tunnel test in the National Transonic Facility (NTF) at NASA LaRC. Further validations were carried out using a Navier-Stokes computational fluid dynamics (CFD) flow solver to calculate wing pressure distributions about several aeroelastically deformed wings and comparing these predictions with NTF experimental data. Results from this aeroelastic deformation method are in good overall agreement with experimentally measured values. Including the predicted deformations significantly improves the correlation between CFD predicted and experimentally measured wing & pressures.

Hooker, John R.↗
NASA NTRS1997

Comparison of Full-Scale XV-15 Wind Tunnel and In-Flight Blade-Vortex Interaction Noise

An isolated full-scale XV-15 rotor was tested in helicopter mode in the NASA Ames 80 by 120-Foot Wind Tunnel. Extensive acoustic data were obtained to define the rotor operating condition for maximum blade-vortex interaction (BVI) noise. Additional data were obtained at operating conditions simulating flight up to 80 knots. An XV-15 aircraft was also tested under operating conditions corresponding to landing approaches for which BVI is expected to be a maximum. In-flight acoustic data were obtained using the YO-3A acoustic research aircraft. An attempt was made to closely match wind tunnel and flight test operating conditions. Details of the two tests are described and some representative acoustic results are presented. Comparisons are shown between the wind tunnel data and corresponding flight test data. Preliminary results indicate very good correlation of the BVI-related features. However, some differences between flight test and wind tunnel results exist away from the BVI event, thought to arise from differences in the two flow environments.

Kitaplioglu, Cahit↗
NASA NTRS1995

Modal Correction Method For Dynamically Induced Errors In Wind-Tunnel Model Attitude Measurements

This paper describes a method for correcting the dynamically induced bias errors in wind tunnel model attitude measurements using measured modal properties of the model system. At NASA Langley Research Center, the predominant instrumentation used to measure model attitude is a servo-accelerometer device that senses the model attitude with respect to the local vertical. Under smooth wind tunnel operating conditions, this inertial device can measure the model attitude with an accuracy of 0.01 degree. During wind tunnel tests when the model is responding at high dynamic amplitudes, the inertial device also senses the centrifugal acceleration associated with model vibration. This centrifugal acceleration results in a bias error in the model attitude measurement. A study of the response of a cantilevered model system to a simulated dynamic environment shows significant bias error in the model attitude measurement can occur and is vibration mode and amplitude dependent. For each vibration mode contributing to the bias error, the error is estimated from the measured modal properties and tangential accelerations at the model attitude device. Linear superposition is used to combine the bias estimates for individual modes to determine the overall bias error as a function of time. The modal correction model predicts the bias error to a high degree of accuracy for the vibration modes characterized in the simulated dynamic environment.

Buehrle, R. D.↗
NASA NTRS1990

Wind-tunnel investigation of a flush airdata system at Mach numbers from 0.7 to 1.4

Flush pressure orifices installed on the nose section of a 1/7-scale model of the F-14 airplane were evaluated for use as a flush airdata system (FADS). Wing-tunnel tests were conducted in the 11- by 11-ft Unitary Wind Tunnel at NASA Ames Research Center. A full-scale FADS of the same configuration was previously tested using an F-14 aircraft at the Dryden Flight Research Facility of NASA Ames Research Center (Ames-Dryden). These tests, which were published, are part of a NASA program to assess accuracies of FADS for use on aircraft. The test program also provides data to validate algorithms for the shuttle entry airdata system developed at the NASA Langley Research Center. The wind-tunnel test Mach numbers were 0.73, 0.90, 1.05, 1.20, and 1.39. Angles of attack were varied in 2 deg increments from -4 deg to 20 deg. Sideslip angles were varied in 4 deg increments from -8 deg to 8 deg. Airdata parameters were evaluated for determination of free-stream values of stagnation pressure, static pressure, angle of attack, angle of sideslip, and Mach number. These parameters are, in most cases, the same as the parameters investigated in the flight test program. The basic FADS wind-tunnel data are presented in tabular form. A discussion of the more accurate parameters is included.

Larson, Terry J.↗
NASA NTRSDate not supplied

SLD Instrumentation in Icing Wind Tunnels – Investigation Overview

A collaborative effort to better understand cloud characterization probes in Supercooled Large Drop (SLD) conditions, as well the ability to simulate these conditions in several icing wind tunnels, was undertaken by NASA, NRCC, CIRA, ECCC, FAA and Met Analytics, Inc. Both drop sizing and liquid water content, LWC, were measured with various probes using current to emerging technologies. To ensure the best possible data quality from the newest probes, the probe manufacturers, SEA, Inc. and Artium, Inc. were invited to support testing and data analysis efforts. A common set of probes was identified to test in each of the three participating facilities: NRCC’s Altitude Icing Wind Tunnel, NASA’s Icing Research Tunnel and CIRA’s Icing Wind Tunnel. From the common set of probes, a subset were identified to use for comparison across the three facilities. These were the CDP-2 and 2D-S for drop sizing, and the Multi-wire for LWC. The LWC value was also checked by measuring the ice accretion thickness under hard rime conditions on a NACA-0012 airfoil. A common test matrix with sweeps in both LWC and median volume diameter, MVD, was developed. Each facility achieved these conditions as determined by their own calibration. The MVD ranged from 20 to at least 200 um, and LWC ranged from 0.5 to 3 g/m3. The comparison probes tested at common conditions in each facility were intended to allow for a direct comparison, and check of potential facility bias.

Icing Wind Tunnel↗
NASA NTRS1982

Jet V/STOL wind-tunnel simulation and groundplane effects

The pretest preparation necessary to define the objectives of an appropriate investigation into the jet V/STOL wind tunnel simulation and ground plane effects were examined. Low speed wind tunnel testing of V/STOL aircraft concepts to determine the aerodynamic propulsion interaction effects during the transition between hover and wingborne flight is a necessary step in the development cycle of this type of aircraft. Powered models are normally used to determine the aerodynamic performance characteristics. Several factors which influence the selection of the model concept and the engine simulator are discussed. Some of the test techniques important for this class of aircraft model are examined. Wind tunnel wall effects important to this aircraft testing with special emphasis on groundplane effects are reviewed.

Margason, R. J.↗
NASA NTRS1974

Wind tunnel model and method

The design and development of a wind tunnel model equipped with pressure measuring devices are discussed. The pressure measuring orifices are integrally constructed in the wind tunnel model and do not contribute to distortions of the aerodynamic surface. The construction of a typical model is described and a drawing of the device is included.

Jackson, C. M., Jr.↗
NASA NTRS1994

The Unitary Plan Wind Tunnels

The Unitary Plan Facility is the most heavily used wind tunnel in all of NASA. Every major commercial transport and almost every fighter built in the United States over the last 30 years has been tested in this tunnel. Also tested in this tunnel complex were models of the Space Shuttle, as well as the Mercury, Gemini, and Apollo capsules. The wind tunnel represents a unique national asset of vital importance to the nation's defense and its competitive position in the world aerospace market. In 1985, the Unitary Plan Facility was named a National Historic Landmark by the National Park Service because of 'its significant associations with the development of the American Space Program.'

Wedgworth, Kevin↗
NASA NTRS2006

Wind Tunnel Results of the B-52B with the X-43A Stack

A low-speed wind-tunnel test was performed with a three-percent-scale model of a booster rocket mated to an X-43A research vehicle, a combination referred to as the Hyper-X launch vehicle. The test was conducted both in free-stream air and in the presence of a partial model of the B-52B airplane. The objectives of the test were to obtain force and moment data to generate structural loads affecting the pylon of the B-52B airplane and to determine the aerodynamic influence of the B-52B airplane on the Hyper-X launch vehicle to evaluate launch separation characteristics. The wind-tunnel test was conducted at a low-speed wind tunnel in Hampton, Virginia. All moments and forces reported are based either on the aerodynamic influence of the B-52B airplane or are for the Hyper-X launch vehicle in free-stream air. Overall, the test showed that the B-52B airplane imparts a strong downwash onto the Hyper-X launch vehicle, reducing the net lift of the Hyper-X launch vehicle. Also, pitching and rolling moments are imparted onto the booster and are a strong function of the launch-drop angle of attack.

Davis, Mark C.↗
NASA NTRS1976

A low speed self streamlining wind tunnel

A two dimensional test section in a low speed wind tunnel is producing flow conditions free from wall interference. The test section has flexible top and bottom walls, and rigid sidewalls from which the models are mounted spanning the tunnel. All walls are unperforated, and the flexible walls are positioned by screw jacks. To eliminate wall interference, the wind tunnel itself supplies the information required in this streamlining process, when run with the model present. Measurements taken at the flexible walls are used by the tunnel computer to check wall contours. When the static pressure distribution in the test section along a contoured flexible wall matches that computed for an imaginary flow field passing over the outside of the same contour, the wall is a streamline in an infinite flow field and the test section flow is free from wall interference. A series of iterations brings the walls from straight to streamlines. Illustrative aerodynamic data is presented, taken on a bluff body and a lifting wing.

Goodyer, M. J.↗
NASA NTRS1996

Wind Tunnel Tests of an Inflatable Airplane

In this video a wind tunnel investigation of aerodynamic and structural deflection characteristics of an inflatable airplane is shown. The film includes scenarios during wind tunnel tests of an inflatable airplane in the Langley Full Scale Tunnel with the main objective of obtaining load factors prior to wing buckle of 4.5 to 5.0 g. The inflation pressure during the test was indicated to be 7.0 psi.

Source record↗
NASA NTRS1984

A comparison of computer-generated lift and drag polars for a Wortmann airfoil to flight and wind tunnel results

Computations of drag polars for a low-speed Wortmann sailplane airfoil are compared to both wind tunnel and flight results. Excellent correlation is shown to exist between computations and flight results except when separated flow regimes were encountered. Wind tunnel transition locations are shown to agree with computed predictions. Smoothness of the input coordinates to the PROFILE airfoil analysis computer program was found to be essential to obtain accurate comparisons of drag polars or transition location to either the flight or wind tunnel results.

Bowers, A. H.↗
NASA NTRS1975

Sound propagation from a simple source in a wind tunnel

The nature of the acoustic field of a simple source in a wind tunnel under flow conditions was examined theoretically and experimentally. The motivation of the study was to establish aspects of the theoretical framework for interpreting acoustic data taken (in wind) tunnels using in wind microphones. Three distinct investigations were performed and are described in detail.

Cole, J. E., III↗
NASA NTRS1981

Wind tunnel supplementary Mach number minimum section insert

A device is described which changes the Mach number capability of a wind tunnel without permanently altering the existing nozzle of the tunnel. An insert is removably attached to the wall of the existing nozzle expansion area thereby creating a second minimum section upstream of the model test section. The added insert may be removed without complicated and expensive changes to the basic wind tunnel. In one embodiment, a removable insert is disposed within wind tunnel nozzle walls with a portion of the flow boundary layer being bled off from the tunnel via passageway and tunnel exit to reduce the extent of separated flow normally occuring upstream of the insert contraction section.

Couch, L. M.↗
NASA NTRS1975

The self streamlining wind tunnel

A two dimensional test section in a low speed wind tunnel capable of producing flow conditions free from wall interference is presented. Flexible top and bottom walls, and rigid sidewalls from which models were mounted spanning the tunnel are shown. All walls were unperforated, and the flexible walls were positioned by screw jacks. To eliminate wall interference, the wind tunnel itself supplied the information required in the streamlining process, when run with the model present. Measurements taken at the flexible walls were used by the tunnels computer check wall contours. Suitable adjustments based on streamlining criteria were then suggested by the computer. The streamlining criterion adopted when generating infinite flowfield conditions was a matching of static pressures in the test section at a wall with pressures computed for an imaginary inviscid flowfield passing over the outside of the same wall. Aerodynamic data taken on a cylindrical model operating under high blockage conditions are presented to illustrate the operation of the tunnel in its various modes.

Goodyer, M. J.↗