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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 1,441 records · Page 80
All sourcesNASA NTRSDOE OSTI
NASA NTRS1975

Hypersonic aeroheating test of space shuttle vehicle configuration 3 (model 22-OTS) in the NASA-Ames 3.5-foot hypersonic wind tunnel (IH20), volume 1

The results of hypersonic wind tunnel testing of an 0.0175 scale version of the vehicle 3 space shuttle configuration are presented. Temperature measurements were made on the launch configuration, orbiter plus tank, orbiter alone, tank alone, and solid rocket booster alone to provide heat transfer data. The test was conducted at free-stream Mach numbers of 5.3 and 7.3 and at free-stream Reynolds numbers of 1.5 million, 3.7 million, 5.0 million, and 7.0 million per foot. The model was tested at angles of attack from -5 deg to 20 deg and side slip angles of -5 deg and 0 deg.

Kingsland, R. B.↗
NASA NTRS1982

Shadowgraph techniques in transonic tests with powered nacelles

Shadowgraph photography techniques in transonic wind tunnel tests are described. The system was used in tests with powered nacelle rings to visualize the shock wave pattern in the exhaust flow. The technique is also used for installed nacelle tests on a floor mounted half model with a turbine powered fan engine simulator.

Hutton, P. G.↗
NASA NTRS2015

Development of Background-Oriented Schlieren for NASA Langley Research Center Ground Test Facilities

This paper provides an overview of recent wind tunnel tests performed at the NASA Langley Research Center where the Background-Oriented Schlieren (BOS) technique was used to provide information pertaining to flow-field density disturbances. The facilities in which the BOS technique was applied included the National Transonic Facility (NTF), Transonic Dynamics Tunnel (TDT), 31-Inch Mach 10 Air Tunnel, 15-Inch Mach 6 High-Temperature Air Tunnel, Rotor Test Cell at the 14 by 22 Subsonic Tunnel, and a 13-Inch Low-Speed Tunnel.

Bathel, Brett F.↗
DOE OSTI2026

CFD-Enhanced Calibration of a Multihole Probe for Small Uncrewed Aircraft Systems

A robust calibration method with a 9-hole probe (9HP) for inertial wind vector measurements from a small uncrewed aircraft system (sUAS) is presented. Calibration accuracy is improved by using computational fluid dynamics (CFD) to estimate corrections, such as those for test section blockage, that are generally applied in wind tunnel testing. A method for estimating experimental bias is presented to account for flow variations over the pressure taps as the 9HP is repositioned in the test section. Installation effects from flow over the airframe and upwash produced by the wings are estimated from CFD simulations of the entire airframe at the expected cruise speed. An initial CFD analysis has demonstrated a quantifiable linear relationship between the measured and actual angle of attack as measured from the relative wind frame. In conclusion, the objective of these additional steps is to increase the accuracy of 9HP calibration.

54 ENVIRONMENTAL SCIENCES↗
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 NTRS1995

Control effectiveness and tip-fin dihedral effects for the HL-20 lifting-body configuration at Mach numbers from 1.6 to 4.5

Wind tunnel tests were made with a scale model of the HL-20 in the Langley Unitary Plan Wind Tunnel. Pitch control was investigated by deflecting the elevon surfaces on the outboard fins and body flaps on the fuselage. Yaw control tests were made with the all movable center fin deflected 5 deg. Almost full negative body flap deflection (-30 deg) was required to trim the HL-20 (moment reference center at 0.54-percent body length from nose) to positive values of life in the Mach number range from 1.6 to 2.5. Elevons were twice as effective as body flaps as a longitudinal trim device. The elevons were effective as a roll control, but because of tip-fin dihedral angle, produced about as much adverse yawing moment as rolling moment. The body flaps were less effective in producing rolling moment, but produced little adverse yawing moment. The yaw effectiveness of the all movable center fin was essentially constant over the angle-of-attack range at each Mach number. The value of yawing moment, however, was small. Center-fin deflection produced almost no rolling moments. The model was directionally unstable over most of the Mach number range with tip-fin dihedral angles less than the baseline value of 50 deg.

Cruz, Christopher I.↗
NASA NTRSDate not supplied

Accounting for the Influence of Decorrelation in Microphone Phased Array Deconvolution Methods

Microphone phased arrays are a common tool for use in aeroacoustic wind tunnel testing. The analysis of acquired array data is known to suffer from decorrelation effects, where the coherence of an acoustic wave measured by a pair of microphones is degraded as the wave passes through a turbulent free shear layer or boundary layer. This paper describes, in detail, how to mitigate the influence of decorrelation effects when processing array data with deconvolution methods. This is done using the DAMAS algorithm as an example, applied to recent airframe noise test data acquired in the NASA Langley 14- by 22-Foot Subsonic Tunnel. Two ways of handling the turbulent propagation modeling, both assuming plane wave propagation, are described. Results show that while turbulence model fit parameters may differ, both methods output extremely similar deconvolution results. Further improvements likely require more accurate mean shear layer data prior to developing more involved turbulence models.

microphone array↗
NASA NTRS1980

Test section configuration for aerodynamic testing in shock tubes

This paper presents results of a study of the test section configuration required to minimize or alleviate interference effects on model flow produced by the presence of test section walls in the aerodynamic testing of two dimensional transonic airfoils in a shock tube. Tests at a nominal Mach number of 0.85 and a chord Reynolds number of 2,000,000 were carried out by means of schlieren photography and pressure measurements for several symmetric airfoil profiles using shock tube test sections with unmodified straight walls, contoured walls, and slotted walls with adjacent chambers. Results were compared with corresponding results from conventional wind tunnel tests of the airfoils. Results for the straight wall tests show major airfoil flow distortions. Results from contoured wall tests and those performed using a slotted wall test section developed in this study exhibit essential agreement with wind tunnel results. The collective results show that test sections for aerodynamic testing can be designed for shock tubes that will alleviate wall interference effects.

Cook, W. J.↗
NASA NTRS1947

Experimental and Calculated Characteristics of Several NACA 44-series Wings with Aspect Ratios of 8, 10, and 12 and Taper Ratios of 2.5 and 3.5

The aerodynamic characteristics of seven unswept tapered wings were determined by calculation from two-dimensional data and by wind-tunnel tests in order to demonstrate the accuracy of the calculations and to show some of the effects of aspect ratio, taper ratio, and root thickness-chord ratio. The characteristics were calculated by the usual application of the lifting-line theory which assumes linear section lift curves and also by an application of the theory which allows the use of nonlinear lift curves. A correction to the lift for the effect of chord was made by using the Jones edge-velocity factor. The wings had aspect ratios of 8, 10, and 12, taper ratios of 2.5 and 3.5, and NACA 44-series airfoils.

WINGS, TAPERED - CHARACTERISTICS↗
NASA NTRSDate not supplied

Analysis of Transonic Unsteady Aerodynamic Environments using Unsteady Pressure Sensitive Paint for the Space Launch System Block 1 Cargo Launch Vehicle

Predicting launch vehicle unsteady aerodynamic loads due to buffet remains a significant challenge. Current practices for modeling buffet environments involve the development of buffet forcing functions using unsteady pressure measurements acquired during wind-tunnel tests. These practices often result in significant uncertainty in buffet environments for coupled loads analyses due to the complex spatio-temporal nature of the unsteady pressure field and challenge of its estimation using discrete sensors. Unsteady pressure sensitive paint, on the other hand, can provide unsteady pressure data at a comparatively high-spatial-density and may overcome the challenge of unsteady pressure field estimation with discrete sensors and lead to improvements in the development of buffet forcing functions. In this paper, comparisons of the fluctuating pressure field are made for the Space Launch System Block 1 cargo launch vehicle measured using unsteady pressure sensitive paint and pressure transducers.

buffet↗
NASA NTRSDate not supplied

Analysis of Transonic Unsteady Aerodynamic Environments using Unsteady Pressure Sensitive Paint for the Space Launch System Block 1 Cargo Launch Vehicle

Predicting launch vehicle unsteady aerodynamic loads due to buffet remains a significant challenge. Current practices for modeling buffet environments involve the development of buffet forcing functions using discrete unsteady pressure measurements acquired during wind-tunnel tests. These practices often result in significant uncertainty in buffet environments for coupled loads analyses due to the complex spatio-temporal nature of the unsteady pressure field and the challenge of its estimation using discrete sensors. Unsteady pressure sensitive paint, on the other hand, can provide unsteady pressure data at a comparatively high spatial density and may overcome the challenge of unsteady pressure field estimation with discrete sensors and lead to improvements in the development of buffet forcing functions. In this paper, comparisons of the fluctuating pressure field are made for the Space Launch System Block 1 cargo launch vehicle measured using unsteady pressure sensitive paint and pressure transducers.

buffet↗
NASA NTRSDate not supplied

Analysis of Transonic Unsteady Aerodynamic Environments using Unsteady Pressure Sensitive Paint for the Space Launch System Block 1 Cargo Launch Vehicle

Predicting launch vehicle unsteady aerodynamic loads due to buffet remains a significant challenge. Current practices for modeling buffet environments involve the development of buffet forcing functions using discrete unsteady pressure measurements acquired during wind-tunnel tests. These practices often result in significant uncertainty in buffet environments for coupled loads analyses due to the complex spatio-temporal nature of the unsteady pressure field and the challenge of its estimation using discrete sensors. Unsteady pressure sensitive paint, on the other hand, can provide unsteady pressure data at a comparatively high spatial density and may overcome the challenge of unsteady pressure field estimation with discrete sensors and lead to improvements in the development of buffet forcing functions. In this paper, comparisons of the fluctuating pressure field are made for the Space Launch System Block 1 cargo launch vehicle measured using unsteady pressure sensitive paint and pressure transducers.

buffet↗
NASA NTRS1974

The 727 airplane side inlet low-speed performance confirmation model test for refanned JT8D engines

The results of a low-speed wind tunnel test of a 0.3 scale model 727 airplane side inlet for JT8D-100 engines are presented. The objectives of the test were to develop lines for a full-scale flightworthy inlet, to evaluate inlet total pressure recovery and steady-state total pressure distortion, and to obtain model-scale distortion data which can be used in the assessment of the compatibility of the inlet with the JT8D-100 series engines. A secondary objective was to obtain internal/external cowl static pressures for the determination of nacelle loads. Two basic inlet models were tested at static, forward speed, angle-of-attack (inflow angle), and cross-wind conditions. One model was with and one without an acoustic ring. Two modifications to the models were also tested, one with the ring closer to the inlet throat and one with a larger lip. Test measurements consisted of inlet surface static pressure, engine face total pressure, inlet airflow, tunnel total pressure, tunnel total temperature and tunnel velocity. Total pressure traverses were taken directly behind the ring and strut. No dynamic measurements were taken.

Schuehle, A. L.↗
NASA NTRS2004

Development of the Planar Inlet Design and Analysis Process (PINDAP)

The aerodynamic development of an engine inlet requires a comprehensive program of both wind tunnel testing and Computational Fluid Dynamics (CFD) simulations. To save time and resources, much "testing" is done using CFD before any design ever enters a wind tunnel. The focus of my project this summer is on CFD analysis tool development. In particular, I am working to further develop the capabilities of the Planar Inlet Design and Analysis Process (PINDAP). "PINDAP" is a collection of computational tools that allow for efficient and accurate design and analysis of the aerodynamics about and through inlets that can make use of a planar (two-dimensional or axisymmetric) geometric and flow assumption. PINDAP utilizes the WIND CFD flow solver, which is capable of simulating the turbulent, compressible flow field. My project this summer is a continuation of work that I performed for two previous summers. Two years ago, I used basic features of the PINDAP to design a Mach 5 hypersonic scramjet engine inlet and to demonstrate the feasibility of the PINDAP. The following summer, I worked to develop its geometry and grid generation capabilities to include subsonic and supersonic inlets, complete bodies and cowls, conic leading and trailing edges, as well as airfoils. These additions allowed for much more design flexibility when using the program.

Gruber, Christopher R.↗
NASA NTRS2015

Improved Regression Analysis of Temperature-Dependent Strain-Gage Balance Calibration Data

An improved approach is discussed that may be used to directly include first and second order temperature effects in the load prediction algorithm of a wind tunnel strain-gage balance. The improved approach was designed for the Iterative Method that fits strain-gage outputs as a function of calibration loads and uses a load iteration scheme during the wind tunnel test to predict loads from measured gage outputs. The improved approach assumes that the strain-gage balance is at a constant uniform temperature when it is calibrated and used. First, the method introduces a new independent variable for the regression analysis of the balance calibration data. The new variable is designed as the difference between the uniform temperature of the balance and a global reference temperature. This reference temperature should be the primary calibration temperature of the balance so that, if needed, a tare load iteration can be performed. Then, two temperature{dependent terms are included in the regression models of the gage outputs. They are the temperature difference itself and the square of the temperature difference. Simulated temperature{dependent data obtained from Triumph Aerospace's 2013 calibration of NASA's ARC-30K five component semi{span balance is used to illustrate the application of the improved approach.

wind tunnel testing↗
NASA NTRS1992

Low-speed cruise aerodynamics of the stopped rotor/disk rotorcraft concept

A low-speed wind tunnel test was completed in support of ongoing conceptual design studies of the Stopped Rotor/Disk rotorcraft concept. A one-fifth scale model was tested in the NASA Ames Low-Speed 7- by 10-Foot Wind Tunnel #1 to evaluate the low-speed cruise performance. The primary test objective was to compare performance characteristics for three possible conceptual designs of the Stopped Rotor/Disk cruise configuration: the large hub fairing (disk) alone, the disk/extended blades configuration, and the disk/conventional wing configuration. Results showed that the disk/extended blades configuration was the most efficient in low-speed cruise. Other test objecives included making parametric changes by varying the geometry of the disk and by varying the extended blade incidence angles. Studies were also conducted to examine the aerodynamic interaction between the disk and a conventional wing. An examination was made into the effects of the disk on static longitudinal stability. The wake generated by the disk impinged on a T-tail of the model and thus degraded longitudinal stability. Alternative tail geometries are required in order to improve the concept's static stability.

Swanson, Stephen M.↗