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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 127 records · Page 7
All sourcesNASA NTRSDOE OSTI
NASA NTRS1990

A preliminary look at techniques used to obtain airdata from flight at high angles of attack

Flight research at high angles of attack has posed new problems for airdata measurements. New sensors and techniques for measuring the standard airdata quantities of static pressure, dynamic pressure, angle of attack, and angle of sideslip were subsequently developed. The ongoing airdata research supporting NASA's F-18 high alpha research program is updated. Included are the techniques used and the preliminary results. The F-18 aircraft was flown with three research airdata systems: a standard airdata probe on the right wingtip, a self-aligning airdata probe on the left wingtip, and a flush airdata system on the nose cone. The primary research goal was to obtain steady-state calibrations for each airdata system up to an angle of attack of 50 deg. This goal was accomplished and preliminary accuracies of the three airdata systems were assessed and are presented. An effort to improve the fidelity of the airdata measurements during dynamic maneuvering is also discussed. This involved enhancement of the aerodynamic data with data obtained from linear accelerometers, rate gyros, and attitude gyros. Preliminary results of this technique are presented.

Moes, Timothy R.↗
NASA NTRS1993

NASA Lewis 8- by 6-foot supersonic wind tunnel user manual

The 8- by 6-Foot Supersonic Wind Tunnel (SWT) at Lewis Research Center is available for use by qualified researchers. This manual contains tunnel performance maps which show the range of total temperature, total pressure, static pressure, dynamic pressure, altitude, Reynolds number, and mass flow as a function of test section Mach number. These maps are applicable for both the aerodynamic and propulsion cycle. The 8- by 6-Foot Supersonic Wind Tunnel is an atmospheric facility with a test section Mach number range from 0.36 to 2.0. General support systems (air systems, hydraulic system, hydrogen system, infrared system, laser system, laser sheet system, and schlieren system are also described as are instrumentation and data processing and acquisition systems. Pretest meeting formats are outlined. Tunnel user responsibility and personal safety requirements are also stated.

Soeder, Ronald H.↗
NASA NTRS1996

The Ames 12-Foot Pressure Tunnel: Tunnel Empty Flow Calibration Results and Discussion

An empty test section flow calibration of the refurbished NASA Ames 12-Foot Pressure Tunnel was recently completed. Distributions of total pressure, dynamic pressure, Mach number, flow angularity temperature, and turbulence are presented along with results obtained prior to facility demolition. Axial static pressure distributions along tunnel centerline are also compared. Test section model support geometric configurations will be presented along with a discussion of the issues involved with different model mounting schemes.

Peter T Zell↗
NASA NTRS2009

Mars Entry Atmospheric Data System Modelling and Algorithm Development

The Mars Entry Atmospheric Data System (MEADS) is being developed as part of the Mars Science Laboratory (MSL), Entry, Descent, and Landing Instrumentation (MEDLI) project. The MEADS project involves installing an array of seven pressure transducers linked to ports on the MSL forebody to record the surface pressure distribution during atmospheric entry. These measured surface pressures are used to generate estimates of atmospheric quantities based on modeled surface pressure distributions. In particular, the quantities to be estimated from the MEADS pressure measurements include the total pressure, dynamic pressure, Mach number, angle of attack, and angle of sideslip. Secondary objectives are to estimate atmospheric winds by coupling the pressure measurements with the on-board Inertial Measurement Unit (IMU) data. This paper provides details of the algorithm development, MEADS system performance based on calibration, and uncertainty analysis for the aerodynamic and atmospheric quantities of interest. The work presented here is part of the MEDLI performance pre-flight validation and will culminate with processing flight data after Mars entry in 2012.

Karlgaard, Christopher D.↗
NASA NTRS2023

Uncertainty Analysis of the NASA Glenn 8- by 6-Foot Supersonic Wind Tunnel 2019 Characterization Test 14-Foot, 5.8 Percent Test Section Porosity

This paper presents methods and results of a measurement uncertainty analysis that was performed for the 8- by 6-foot Supersonic Wind Tunnel located at the NASA Glenn Research Center. The uncertainty analysis is based on data collected during a characterization test following major structural modifications to the facility and upgrades to the facility’s data and control systems. The statistical methods and engineering judgments used to estimate elemental uncertainties are described in this report. The Monte Carlo method of propagating uncertainty was selected to determine the uncertainty of calculated variables of interest. A description of the Monte Carlo method as applied for this analysis is provided. The primary variable of interest (VOI) for this facility is free stream Mach number. In addition to determining the uncertainty in Mach number, the uncertainty in free stream values of static pressure, total pressure, dynamic pressure, total temperature, static temperature, Reynolds number, and air speed were also calculated. Uncertainty results are presented as random (unpredictable variation in repeated measurements), systematic (potential offset between observed and true value), and total (random and systematic combined) uncertainty for Mach number. Systematic uncertainty results are presented for the other VOIs. Individual uncertainty sources are presented both dimensionally and as percent contributions to uncertainty in all VOIs, to aid in the identification of primary uncertainty sources.

Uncertainty Analysis↗
NASA NTRS2014

Aeroelastic Analyses of the SemiSpan SuperSonic Transport (S4T) Wind Tunnel Model at Mach 0.95

Detailed aeroelastic analyses of the SemiSpan SuperSonic Transport (S4T) wind tunnel model at Mach 0.95 with a 1.75deg fixed angle of attack are presented. First, a numerical procedure using the Computational Fluids Laboratory 3-Dimensional (CFL3D) Version 6.4 flow solver is investigated. The mesh update method for structured multi-block grids was successfully applied to the Navier-Stokes simulations. Second, the steady aerodynamic analyses with a rigid structure of the S4T wind tunnel model are reviewed in transonic flow. Third, the static analyses were performed for both the Euler and Navier-Stokes equations. Both the Euler and Navier-Stokes equations predicted a significant increase of lift forces, compared to the results from the rigid structure of the S4T wind-tunnel model, over various dynamic pressures. Finally, dynamic aeroelastic analyses were performed to investigate the flutter condition of the S4T wind tunnel model at the transonic Mach number. The condition of flutter was observed at a dynamic pressure of approximately 75.0-psf for the Navier-Stokes simulations. However, it was observed that the flutter condition occurred a dynamic pressure of approximately 47.27-psf for the Euler simulations. Also, the computational efficiency of the aeroelastic analyses for the S4T wind tunnel model has been assessed.

Hur, Jiyoung↗
NASA NTRS1983

A technique for dynamically calibrating pressure transducers at cryogenic temperatures

A technique was developed for the calibration of dynamic pressure transducers at cryogenic temperatures. The calibration system utilizes an 8.9 Newton peak thrust shaker which oscillates a helium-filled bellows to generate a sinusoidal dynamic pressure to calibrate transducers immersed in a cryogenic environment. The system has a dynamic pressure measurement uncertainty of approximately 11% and is capable of producing peak-to-peak dynamic pressure amplitudes of 1.4 kPa over a frequency range of 40 to 100 hertz and a temperature range of 100 to 300 K. It provides an unprecedented capability of both static and dynamic calibration of pressure transducers from ambient to cryogenic temperature.

Gibbens, B. V.↗
NASA NTRS1972

Improvement of a large-amplitude sinusoidal pressure generator for dynamic calibration of pressure transducers

Results of research on the improvement of a sinusoidal pressure generator are presented. The generator is an inlet-area-modulated, gas-flow-through device (siren type) which was developed to dynamically calibrate pressure transducers and pressure probes. Tests were performed over a frequency range of 100 Hz to 20 kHz at average chamber pressures (bias pressure) between 30 and 50 psia (21 and 35 N/sq cm abs) and between 150 and 300 psia (104 and 207 N/sq cm abs). Significant improvements in oscillation pressure waveform were obtained but with reduction in available generator oscillation pressure amplitude range. Oscillation pressure amplitude, waveform, and waveform spectral content are given as functions of frequency for the two bias pressure conditions. The generator and instrumentation for frequency, amplitude, and spectrum measurements are described.

Robinson, R. E.↗
NASA NTRS1987

Oscillation pressure device for dynamic calibration of pressure transducers

Method and apparatus for obtaining dynamic calibrations of pressure transducers. A calibration head (15), a flexible tubing (23) and a bellows (20) enclose a volume of air at atmospheric pressure with a transducer (11) to be calibrated subject to the pressure inside the volume. All of the other apparatus in the drawing apply oscillations to bellows (20) causing the volume to change thereby applying oscillating pressures to transducer (11) whereby transducer (11) can be calibrated.

Hess, Robert W.↗
NASA NTRS1973

Surface pressure fluctuations in hypersonic turbulent boundary layers.

The surface pressure fluctuations on a flat-plate model at hypersonic Mach numbers of 5.2, 7.4, and 10.4 with an attached turbulent boundary layer were measured using flush-mounted small piezoelectric sensors. A high-frequency resolution of the pressure field was achieved using specially designed small piezoelectric sensors. The rms pressures and nondimensional energy spectra for all above Mach numbers are presented. The convective velocities, obtained from space-time correlation considerations are equal to 0.7 of the free-stream velocity. The results indicate that the rms pressures vary from 5 to 25% of the mean static pressures. The ratios of rms pressure to dynamic pressure are less than the universally accepted subsonic value of .006. The ratio decreases in value as the Mach number or the dynamic pressure is increased.

Raman, K. R.↗
NASA NTRS1974

Surface pressure fluctuations in hypersonic turbulent boundary layers

The surface pressure fluctuations on a flat plate model at hypersonic Mach numbers of 5.2, 7.4 and 10.4 with an attached turbulent boundary layer were measured using flush mounted small piezoelectric sensors. A high frequency resolution of the pressure field was achieved using specially designed small piezoelectric sensors that had a good frequency response well above 300 KHz. The RMS pressures and non-dimensional energy spectra for all above Mach numbers are presented. The convective velocities, obtained from space time correlation considerations are equal to 0.7 U sub infinity. The results indicate the RMS pressures vary from 5 to 25 percent of the mean static pressures. The ratios of RMS pressure to dynamic pressure are less than the universally accepted subsonic value of 6 x 10/3. The ratio decreases in value as the Mach number or the dynamic pressure is increased. The ratio of RMS pressure to wall shear for Mach number 7.4 satisfies one smaller than or equal to p/tau sub w smaller than or equal to three.

Raman, K. R.↗
NASA NTRS1981

The development of cryogenic wind tunnels and their application to maneuvering aircraft technology

Cryogenic wind tunnels are considered as a means of studying high Reynolds number (Re) complicated flows encountered by high maneuvering lift and high angles of attack characteristic of modern fighter aircraft. Large decreases in the viscous force while the inertial force remains constant are provided by the use of cryogenic facilities. A 2.5 m square tunnel is nearing completion at the National Transonic Facility (NTF), and will be driven by synchronous motors having a total power of 120,000 hp. The tunnel, using N2 as the cryogenic fluid, will allow large Re sweeps at constant dynamic pressure and dynamic pressure, and aeroelastic sweeps at constant Re; full altitude (air density) and acceleration force simulation will also be possible. Advances in model and strain gage balance technologies for use at the NTF are outlined, and experiments with buffet are described.

Polhamus, E. C.↗