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At least 775 records · Page 43
All sourcesNASA NTRSDOE OSTI
NASA NTRS1972

A Wind Tunnel Flight Correlation of Apollo 15 Sonic Boom

A correlation of sonic boom pressure signatures recorded during reentry of the Apollo 15 command module with wind-tunnel signatures extrapolated to flight distances has been made for Mach numbers of 1.16 and 4.57. The flight pressure signatures were recorded by pressure sensors located onboard ships positioned near the ground track while the wind-tunnel signatures were measured during tests of a 0.016-scale model of the command module. The agreement between estimates based on wind-tunnel data and flight measurements was better at Mach 4.57 than at Mach 1.16.

Hicks, R. M.↗
NASA NTRS2008

Aeroservoelastic Testing of a Sidewall Mounted Free Flying Wind-Tunnel Model

A team comprised of the Air Force Research Laboratory (AFRL), Northrop Grumman, Lockheed Martin, and the NASA Langley Research Center conducted three j wind-tunnel tests in the Transonic Dynamics Tunnel to demonstrate active control technologies relevant to large, exible vehicles. In the rst of these three tests, a semispan, aeroelastically scaled, wind-tunnel model of a ying wing SensorCraft vehi- cle was mounted to a force balance to demonstrate gust load alleviation. In the second and third tests, the same wing was mated to a new, multi-degree-of-freedom, sidewall mount. This mount allowed the half-span model to translate vertically and pitch at the wing root, allowing better simulation of the full span vehicle's rigid-body modes. Gust Load Alleviation (GLA) and Body Freedom Flutter (BFF) suppression were successfully demonstrated. The rigid body degrees-of-freedom required that the model be own in the wind tunnel using an active control system. This risky mode of testing necessitated that a model arrestment system be integrated into the new mount. The safe and successful completion of these free-flying tests required the development and integration of custom hardware and software. This paper describes the many systems, software, and procedures that were developed as part of this effort.

Scott, Robert C.↗
NASA NTRS1986

Surface roughness studies for wind tunnel models used in high Reynolds number testing

This paper focuses on stylus and optical techniques for the measurement of surface roughness in wind tunnel models. The stylus instruments provide detailed information, such as surface profiles and area maps, that may then be used either to calculate statistical properties (i.e., the rms surface roughness) or to study individual surface peaks or other features. By contrast, certain optical techniques yield area-averaged statistical properties of the surface roughness directly. Two instruments that use the technique of optical angular scattering are compared. One is a research instrument that has been developed to study the basic scattering phenomena by testing the optical theories and surface models used in inverse calculations of statistical roughness parameters. The second instrument is more compact and is under development as a hand held, on-line device to be used during manufacture of wind tunnel models for the National Transonic Facility at NASA Langley Research Center. The scattering geometries for the two instruments are compared and results from these instruments and the stylus technique are shown for roughness specimens that are typical of the surface finish of wind tunnel models.

Vorburger, T. V.↗
NASA NTRS2002

Design and Development of a Deep Acoustic Lining for the 40-by 80-Foot Wind Tunnel Test Section

The work described in this report has made effective use of design teams to build a state-of-the-art anechoic wind-tunnel facility. Many potential design solutions were evaluated using engineering analysis, and computational tools. Design alternatives were then evaluated using specially developed testing techniques, Large-scale coupon testing was then performed to develop confidence that the preferred design would meet the acoustic, aerodynamic, and structural objectives of the project. Finally, designs were frozen and the final product was installed in the wind tunnel. The result of this technically ambitious project has been the creation of a unique acoustic wind tunnel. Its large test section (39 ft x 79 ft x SO ft), potentially near-anechoic environment, and medium subsonic speed capability (M = 0.45) will support a full range of aeroacoustic testing-from rotorcraft and other vertical takeoff and landing aircraft to the take-off/landing configurations of both subsonic and supersonic transports.

Soderman, Paul T.↗
NASA NTRS2006

Estimation of Longitudinal Unsteady Aerodynamics of a Wing-Tail Combination From Wind Tunnel Data

This paper presents an initial step toward model identification from wind tunnel data for an airliner configuration. Two approaches to modeling a transport configuration are considered and applied to both steady and large-amplitude forced-oscillation wind tunnel data taken over a wide range of angles of attack. Only limited conclusions could be drawn from this initial data set. Although model estimated time histories of normal force and pitching moment agree reasonably well with the corresponding measured values, model damping parameters did not, for some cases, have values consistent with small amplitude oscillatory data. In addition, large parameter standard errors implied poor information content for model structure determination and parameter estimation. Further investigation of the modeling problem for more general aerodynamic models is recommended with close attention to experiment design for obtaining parameters with high accuracy.

Murphy, Patrick C.↗
NASA NTRS2011

Axisymmetric Calculations of a Low-Boom Inlet in a Supersonic Wind Tunnel

This paper describes axisymmetric CFD predictions made of a supersonic low-boom inlet with a facility diffuser, cold pipe, and mass flow plug within wind tunnel walls, and compares the CFD calculations with the experimental data. The inlet was designed for use on a small supersonic aircraft that would cruise at Mach 1.6, with a Mach number over the wing of 1.7. The inlet was tested in the 8-ft by 6-ft Supersonic Wind Tunnel at NASA Glenn Research Center in the fall of 2010 to demonstrate the performance and stability of a practical flight design that included a novel bypass duct. The inlet design is discussed here briefly. Prior to the test, CFD calculations were made to predict the performance of the inlet and its associated wind tunnel hardware, and to estimate flow areas needed to throttle the inlet. The calculations were done with the Wind-US CFD code and are described in detail. After the test, comparisons were made between computed and measured shock patterns, total pressure recoveries, and centerline pressures. The results showed that the dual-stream inlet had excellent performance, with capture ratios near one, a peak core total pressure recovery of 96 percent, and a large stable operating range. Predicted core recovery agreed well with the experiment but predicted bypass recovery and maximum capture ratio were high. Calculations of offdesign performance of the inlet along a flight profile agreed well with measurements and previous calculations.

Chima, Rodrick V.↗
NASA NTRS1981

A review of flight-to-wind tunnel drag correlation

Comparisons are made of wind-tunnel-model and flight drag data for various configurations representing aircraft from the mid-1940s to the 1970s. Discrepancies between model and flight data such as Reynolds number effects, wall interference, and aeroelastic problems are discussed. String support effects and the inability of models to simulate surface deflections for longitudinal trim are also studied. A wind tunnel-to-flight correlation of turbulent friction drag confirms the incompressible Karman-Schoenherr variation of turbulent skin friction with Reynolds number and the T' method for accounting compressibility effects. NASA tested 10 deg cone research indicates that model tests which are affected by tunnel noise may require the lower disturbance level environment available in flight, and it is concluded that new cryogenic facilities will improve the fidelity of model simulations of full-scale flight flow phenomena.

Saltzman, E. J.↗
NASA NTRS1974

Results of investigations on an 0.015 scale 140A/B configuration space shuttle vehicle orbiter model (49-0) in the LTV 4 by 4-foot high speed wind tunnel (0A84)

Data obtained from a wind tunnel test of an 0.015-scale 140 A/B configuration space shuttle vehicle orbiter model (49-0) are presented. Runs were conducted at Mach numbers 0.6, 0.9, 1.2, 1.6, and 2.0 and Reynolds numbers of 8.4, 9.0, 9.0, 10.9, and 13.6 million per foot respectively. Various control surface settings were investigated from angles of attack of minus 4 degrees to plus 30 degrees at fixed angles of sideslip of zero and minus 5 degrees and through angles of sideslip from minus 2 to plus 8 degrees at fixed angles of attack of zero, plus 10, plus 15, and plus 20 degrees. The purpose of the test was to define the longitudinal and lateral-directional stability and control characteristics for the updated SSV configuration.

Esparza, V.↗
NASA NTRS1975

Data report for tests on the heat transfer effects of the 0.0175 scale Rockwell International Space Shuttle Vehicle model 22-OT in the AEDC 50 inch B wind tunnel (0H4B), volume 1

Results of wind tunnel heat transfer tests of 0.0175-scale Rockwell International Space Shuttle Vehicle configurations for orbiter alone, tank alone, and orbiter plus external tank are presented. Body flap shielding of SSME's during simulated entry was investigated. The tests were conducted at Mach 8 for thirteen Reynolds number.

Foster, T. F.↗
NASA NTRS1975

Data report for tests on the heat transfer effects of the 0.0175-scale Rockwell International Space Shuttle Vehicle model 22-OT in the AEDC 50-inch B wind tunnel (OH4B), volume 3

Results of wind tunnel heat transfer tests of 0.0175-scale Rockwell International Space Shuttle Vehicle configurations for orbiter alone, tank alone, and orbiter plus external tank are presented. Body flap shielding of SSME's during simulated entry was also investigated. The tests were conducted at Mach 8 for thirteen Reynolds number per foot values ranging from 0.5 million to 3.72 million.

Foster, T. F.↗
NASA NTRS1975

Results of investigations on an 0.015-scale model (49-0) of the Rockwell International Space Shuttle orbiter in the NASA-Ames Research Center 3.5-foot hypersonic wind tunnel (0A98)

The results of a wind tunnel test are presented; the model used for this test was 0.015-scale 140 A/B hybrid configuration of the space shuttle orbiter. The primary test objectives were to obtain incremental data on the effects of a sting mount on base pressures and force and moment data. The increments obtained included the addition of MPS nozzles as well as the deletion of the simulated sting mount. Six-component aerodynamic force and moment data were recorded over an angle of attack range from 12 to 42 degrees at 0 and 5 degrees angles of sideslip. The testing was accomplished at Mach 5.3 and Mach 10.3. The effects of various elevon, body flap, and speed brake settings were investigated, and static pressures were measured at the fuselage base for use in force-data reduction.

Milam, M. D.↗
NASA NTRS1976

Terminal area energy management regime investigations utilizing an 0.030-scale model (47-0) of the space shuttle vehicle orbiter configuration 140A/B/C/R in the Ames Research Center 11 x 11 foot transonic wind tunnel (OH/48)

Data obtained in a wind tunnel test were examined to: (1) obtain pressure distributions, forces and moments over the vehicle 5 Orbiter in the terminal area energy management (TAEM) and approach phases of flight; (2) obtain elevon and rudder hinge moments in the TAEM and approach phases of flight; (3) obtain body flap and elevon loads for verification of loads balancing with integrated pressure distributions; and (4) obtain pressure distributions near the short OMS pods in the high subsonic, transonic and low supersonic Mach number regimes. Testing was conducted over a Mach number range from 0.6 to 1.4 with Reynolds number variations from 7.57 x 1 million to 2.74 x 1 million per foot. Model angle of attack was varied from -4 to 16 degrees and angles of sideslip ranged from -8 to 8 degrees.

Hawthorne, P. J.↗
NASA NTRS1976

Aeroelastic stability and control of an oblique wing - Wind tunnel experiments

Results are presented for wind tunnel tests of an elastic wing model to verify the theoretical predictions for the aeroelastic instability of an oblique wing. The model wing has an elliptic planform of 10 to 1 axis ratio and a symmetrical airfoil section of 7-1/2% thickness/chord ratio. The wing is of wood and as may be seen in the photographs presented, slack wires are used to limit the amplitude of unstable motions. The fuselage is mounted on bearings permitting freedom of roll, but provision is made to clamp the fuselage for some of the tests. It is found that freedom in roll increases the dynamic pressure at which aeroelastic instability first appears. With the model free in roll, the effectiveness of the ailerons in maintaining trim is not noticeably affected by passage through the speed at which the wing would become unstable if clamped.

Jones, R. T.↗
NASA NTRS1976

Results of the Low Speed Aeroelastic Buffet Test with a 0.046-scale Model (747-ax1322-d-3/orbiter 8-0) of the 747 Cam/orbiter in the University of Washington Wind Tunnel (CS 3)

A series of wind tunnel studies designed to assess the potential buffet problems resulting from orbiter wake characteristics with its tailcone removed are presented to provide design loads and acceleration environments, and to develop data on buffet sensitivity to various aerodynamic configurations and flight parameters. Data are intended to support subsequent analyses of structural fatigue life, crew efficiency, and equipment vibrations.

Gillins, R. L.↗
NASA NTRS1980

Comparison of calculated and measured blade loads on a full-scale tilting proprotor in a wind tunnel

The loads measured in a wind tunnel on a full-scale tilting proprotor are compared with calculated results. The data consists primarily of oscillatory beamwise bending moments at 35% radial station, oscillatory spindle chord bending moments, and oscillatory pitch link loads. The measured and calculated results as a function of thrust are compared over a range of nacelle angles from 0 to 75 deg, and a range of speeds from 80 to 185 knots.

Johnson, W.↗
NASA NTRS1985

Optimization studies for the development of adaptive wall wind tunnel

The development of adaptive wall wind tunnels makes it possible to establish interference-free boundaries in the region of the model. It is, therefore, feasible to simulate free air conditions. However, the amount of total testing time involved in establishing the required conditions and the complexity of the needed instrumentation prevent currently the use of such methods for routine production testing. Optimization studies have been proposed with the objective to find approaches which will make a more extended application of adaptive wall techniques possible. A new 25cm x 11cm adaptive wall test section was designed and installed to carry out some of these studies. Investigations envisaged are related to the development of a one step convergence method to reduce total testing time, and to measurement procedures which will permit a simplification of the instrumentation. Attention is given to the proposed concept, the test section, instrumentation, and the influence coefficients.

Bodapati, S.↗
NASA NTRS1988

Propfan test assessment testbed aircraft stability and control/performance 1/9-scale wind tunnel tests

One-ninth scale wind tunnel model tests of the Propfan Test Assessment (PTA) aircraft were performed in three different NASA facilities. Wing and propfan nacelle static pressures, model forces and moments, and flow field at the propfan plane were measured in these tests. Tests started in June 1985 and were completed in January 1987. These data were needed to assure PTA safety of flight, predict PTA performance, and validate analytical codes that will be used to predict flow fields in which the propfan will operate.

Little, B. H., Jr.↗