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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 217 records · Page 12
All sourcesNASA NTRSDOE OSTI
NASA NTRS1972

Transonic wall interference effects on bodies of revolution.

Efforts to develop a near sonic transport have placed renewed emphasis on obtaining accurate aerodynamic force and pressure data in the near sonic speed range. Comparison of wind-tunnel and flight data obtained for a blunt-nose body of revolution showed significant discrepancies in drag levels near Mach 1 - apparently due to wind-tunnel wall interference. Subsequent tests of geometrically similar bodies of revolution showed that increasing the model-to-test-section blockage ratio from 0.00017 to 0.0043 resulted in altered drag curve shapes, delayed drag divergence, and 'transonic creep' from subsonic drag levels due to increased wall interference.

Couch, L. M.↗
NASA NTRS1983

An experimental investigation of the subcritical and supercritical flow about a swept semispan wing

An experimental investigation of the turbulent, subcritical and supercritical flow over a swept, semispan wing in a solid wall wind tunnel is described. The program was conducted over a range of Mach numbers, Reynolds numbers, and angles of attack to provide a variety of test cases for assessment of wing computer codes and tunnel wall interference effects. Wing flows both without and with three dimensional flow separation are included. Data include mean surface pressures for both the wing and tunnel walls; surface oil flow patterns on the wing; and mean velocity, flow field surveys. The results are given in tabular form and presented graphically to illustrate some of the effects of the test parameters. Comparisons of the wing pressure data with the results from two inviscid wing codes are also shown to assess the importance of viscous flow and tunnel wall effects.

Lockman, W. K.↗
NASA NTRS1991

Progress toward synergistic hypermixing nozzles

Mean flow measurements were obtained for air-to-air mixing downstream of swept and unswept ramp wall mounted hypermixing nozzle configurations. Aside from the sweep of the ramps, the two nozzle configurations studied are identical. The nozzles inject three parallel supersonic jets at a 15 deg angle (relative to the wind tunnel wall) into a supersonic freestream. Mach number and volume fraction distributions in a transverse plane 11.1 nozzle heights downstream from the nozzle exit plane were measured. Data are presented for a freestream Mach number of three at a matched static pressure condition and also at underexpanded static pressure condition (pressure ratio = 5). Surface oil flow visualization was used to study the near wall flow behavior. The results indicate that the swept ramp injectors produce stronger and larger vortex pairs than the unswept ramp injectors. The increased interaction between the swept ramp model's larger vortex pairs yields better mixing characteristics for this model.

Davis, D. O.↗
NASA NTRS1991

Progress toward synergistic hypermixing nozzles

Mean flow measurements have been obtained for air-to-air mixing downstream of swept and unswept ramp wall-mounted hypermixing nozzle configurations. Aside from the sweep of the ramps, the two nozzle configurations investigated are identical. The nozzles inject three parallel supersonic jets at a 15-deg angle (relative to the wind tunnel wall) into a supersonic freestream. Mach number and volume fraction distributions in a transverse plane 11.1 nozzle heights downstream from the nozzle exit plane were measured. Data are presented for a freestream Mach number of three at a matched static pressure condition and also at an underexpanded static pressure condition (pressure ratio equal to 5). Surface oil flow visualization was used to investigate the near-wall flow behavior. The results indicate that the swept ramp injectors produce stronger and larger vortex pairs than the unswept ramp injectors. The increased interaction between the swept ramp model's larger vortex pairs yields better mixing characteristics for this model.

Davis, D. O.↗
NASA NTRS2007

Wall Boundary Layer Measurements for the NASA Langley Transonic Dynamics Tunnel

Measurements of the boundary layer parameters in the NASA Langley Transonic Dynamics tunnel were conducted during extensive calibration activities following the facility conversion from a Freon-12 heavy-gas test medium to R-134a. Boundary-layer rakes were mounted on the wind-tunnel walls, ceiling, and floor. Measurements were made over the range of tunnel operation envelope in both heavy gas and air and without a model in the test section at three tunnel stations. Configuration variables included open and closed east sidewall wall slots, for air and R134a test media, reentry flap settings, and stagnation pressures over the full range of tunnel operation. The boundary layer thickness varied considerably for the six rakes. The thickness for the east wall was considerably larger that the other rakes and was also larger than previously reported. There generally was some reduction in thickness at supersonic Mach numbers, but the effect of stagnation pressure, and test medium were not extensive.

Wieseman, Carol D.↗
NASA NTRS1992

Modeling of solid/porous wall boundary conditions for the validation of computational fluid dynamics codes

A computational study has been undertaken to investigate method of modeling solid and porous wall boundary conditions in computational fluid dynamics (CFD) codes. The procedure utilizes experimental measurements at the walls to develop a flow field solution based on the method of singularities. This flow field solution is then imposed as a boundary condition in a CFD simulation of the internal flow field. The effectiveness of this method in describing the boundary conditions at the wind tunnel walls using only sparse experimental measurements has been investigated. Position and refinement of experimental measurement locations required to describe porous wall boundary conditions has also been considered.

Beutner, Thomas J.↗
NASA NTRS1986

Design requirements for the NASA Langley supersonic low-disturbance wind tunnel

The high intensity, high frequency acoustic disturbances that cause large adverse boundary layer transition effects on test models in conventional supersonic wind tunnels consist of finite length wavelets radiating from eddies in the turbulent boundary layers of the wind tunnel walls. NASA Langley has undertaken 'quiet' supersonic tunnel research that demonstrates the ability to maintain laminarity at high unit Reynolds numbers on limited upstream regions of the nozzle wall boundary layers in small, Mach 3.0, 3.5, and 5.0 pilot tunnels. The high level acoustic disturbances are then eliminated, and the transition Reynolds numbers measured on cones approach those for atmospheric flight. Attention is presently given to the design requirements that can be extrapolated from these results for a large wind tunnel facility; high quality air filtering, noise attenuation, nozzle coordinate accuracy, and surface finish, are quantified with pilot tunnel data.

Beckwith, I. E.↗
NASA NTRS2018

Simulations of the NASA Langley 14- by 22-Foot Subsonic Tunnel for the Juncture Flow Experiment

NASA’s Transformational Tools and Technologies Program’s Juncture Flow experiment aims to provide data to improve Computational Fluid Dynamics (CFD) modeling in the juncture flow region. The experiment is planned to provide validation-quality data for CFD that focuses on the separation bubble near the wing-body juncture trailing edge region. Because wind tunnel tests associated with the Juncture Flow project have been designed for the purpose of CFD validation, considerable effort is going into modeling and simulating the wind tunnel. This is not only important because wind tunnel wall effects can play a role in integrated testing uncertainties, but also because the better the boundary conditions are known, the better CFD can accurately represent the experiment. This paper builds on the recent CFD efforts to model the NASA Langley 14- by 22-Foot Subsonic Tunnel. Current best practices in simulating wind tunnels are evaluated. The features of each method, as well as some of their pros and cons, are highlighted. Boundary conditions and modeling techniques currently used by CFD for empty-tunnel simulations are also described. Preliminary CFD studies associated with modeling the Juncture Flow model are summarized, with the intention to determine sensitivities of the flow near the wing-body juncture region of the model to a variety of modeling decisions.

Subsonic↗
NASA NTRSDate not supplied

Wall-modeled LES of the NASA Juncture Flow Experiment

We perform wall-modeled large eddy simulation (WMLES) of the NASA Juncture Flow experiment (Kegerise & Neuhart, NASA/TM–2019–20286) using an equilibrium wall model and unstructured finite volume solver, to assess its predictive ability for complex geometries. The flow condition simulated corresponds to 5 degrees angle of attack with a freestream Mach number of 0.189, and a Reynolds number based on the mean crank chord of 2.4 million. All simulations neglect the effects of wind tunnel walls, sting and mast present in the experiment. To assess the performance of WMLES on simpler flows with very coarse grids, we simulate flow in a turbulent channel at a friction Reynolds number, Reτ≈2000, and flow through a square duct at anReτ≈1000. The duct flow results using the equilibrium wall model indicate that the skin friction is inaccurate near the corner for grid topologies and resolutions typically used in WMLES, having implications for more complicated flows such as the juncture flow. For the juncture flow simulation, two different approaches are investigated. The first approach consists of a truncated-domain simulation wherein the inflow plane of the computational domain is placed at≈0.2chord length of the wing, and we prescribe the mean flow from a separate Reynolds-averaged Navier-Stokes (RANS) solution along with synthetic turbulence to initiate realistic unsteadiness in the domain. The second approach involves simulating the entire geometry with trips to trigger transition to turbulence. The truncated- and full-domain simulations contain about 62 and 90 million cells, respectively; with 8-10 points per boundary layer thickness in the wall-parallel directions, 14-20 points in the wall-normal direction and a near-wall viscous spacing (∆n+1) of≈100. Preliminary results are encouraging overall in terms of the prediction of wall pressure, wall skin friction, velocity and stresses; but indicate that further work is required to improve the predictions in the separation bubble and wing-fuselage corner regions. The grid resolution used in this study is still fairly coarse, and the results should be interpreted as work in progress.

Prahladh S Iyer↗
NASA NTRS2019

Modeling of an Atmospheric-Boundary-Layer Profile in Support of Experiments in the NASA Langley Transonic Dynamics Tunnel

This paper presents the results of the Computational Fluid Dynamics (CFD) analyses to model two atmospheric-boundary-layer (ABL) profiles inside the NASA Langley Research Center Transonic Dynamics Tunnel (TDT). The CFD models include tunnel walls and all additional hardware needed to simulate the ABL profiles. This hardware includes Irwin spires and floor-mounted roughness elements. The numerical simulations show that the application of higher-fidelity numerical methods is necessary to compute boundary-layer and turbulenceintensity profiles that match experimental data. The ABL profiles are computed both inside the numerical model of the TDT and in a classical free-air model. Both Unsteady Reynolds Averaged Navier-Stokes (URANS) with Spalart-Allmaras (SA) turbulence model and Modified Delayed Detached Eddy (MDDES) simulation methods are used. The results show that the MDDES-simulation results match the experimental data very well while URANS-SA does not. The results also show that the wind-tunnel walls have a significant effect on ABL prediction.

Chwalowski, Pawel↗
NASA NTRS1979

Effect of Mach number and Reynolds number on a normal shock-wave/turbulent boundary-layer interaction

Experimental values of wall pressure and skin friction have been obtained for Mach numbers from 1.32 to 1.48 and for Reynolds numbers from 8.5 x 10 to the 6th to 225 x 10 to the 6th. Increasing the Mach number reduces the wall shear and promotes incipient separation. Reynolds number variations have little effect except at the lowest value where the flow abruptly approaches separation. Comparisons are made with solutions to the time-dependent, Reynolds-averaged, Navier-Stokes equations incorporating a two-equation, Wilcox-Rubesin turbulence model. The computations are in agreement with the experimental results. Additional numerical results indicated that the wind-tunnel walls constrained the flow and suppressed the formation of a separation bubble at the shock wave.

Mateer, G. G.↗
NASA NTRS2006

NASA Tech Briefs, September 2006

Topics covered include: Improving Thermomechanical Properties of SiC/SiC Composites; Aerogel/Particle Composites for Thermoelectric Devices; Patches for Repairing Ceramics and Ceramic- Matrix Composites; Lower-Conductivity Ceramic Materials for Thermal-Barrier Coatings; An Alternative for Emergency Preemption of Traffic Lights; Vehicle Transponder for Preemption of Traffic Lights; Automated Announcements of Approaching Emergency Vehicles; Intersection Monitor for Traffic-Light-Preemption System; Full-Duplex Digital Communication on a Single Laser Beam; Stabilizing Microwave Frequency of a Photonic Oscillator; Microwave Oscillators Based on Nonlinear WGM Resonators; Pointing Reference Scheme for Free-Space Optical Communications Systems; High-Level Performance Modeling of SAR Systems; Spectral Analysis Tool 6.2 for Windows; Multi-Platform Avionics Simulator; Silicon-Based Optical Modulator with Ferroelectric Layer; Multiplexing Transducers Based on Tunnel-Diode Oscillators; Scheduling with Automated Resolution of Conflicts; Symbolic Constraint Maintenance Grid; Discerning Trends in Performance Across Multiple Events; Magnetic Field Solver; Computing for Aiming a Spaceborne Bistatic- Radar Transmitter; 4-Vinyl-1,3-Dioxolane-2-One as an Additive for Li-Ion Cells; Probabilistic Prediction of Lifetimes of Ceramic Parts; STRANAL-PMC Version 2.0; Micromechanics and Piezo Enhancements of HyperSizer; Single-Phase Rare-Earth Oxide/Aluminum Oxide Glasses; Tilt/Tip/Piston Manipulator with Base-Mounted Actuators; Measurement of Model Noise in a Hard-Wall Wind Tunnel; Loci-STREAM Version 0.9; The Synergistic Engineering Environment; Reconfigurable Software for Controlling Formation Flying; More About the Tetrahedral Unstructured Software System; Computing Flows Using Chimera and Unstructured Grids; Avoiding Obstructions in Aiming a High-Gain Antenna; Analyzing Aeroelastic Stability of a Tilt-Rotor Aircraft; Tracking Positions and Attitudes of Mars Rovers; Stochastic Evolutionary Algorithms for Planning Robot Paths; Compressible Flow Toolbox; Rapid Aeroelastic Analysis of Blade Flutter in Turbomachines; General Flow-Solver Code for Turbomachinery Applications; Code for Multiblock CFD and Heat-Transfer Computations; Rotating-Pump Design Code; Covering a Crucible with Metal Containing Channels; Repairing Fractured Bones by Use of Bioabsorbable Composites; Kalman Filter for Calibrating a Telescope Focal Plane; Electronic Absolute Cartesian Autocollimator; Fiber-Optic Gratings for Lidar Measurements of Water Vapor; Simulating Responses of Gravitational-Wave Instrumentation; SOFTC: A Software Correlator for VLBI; Progress in Computational Simulation of Earthquakes; Database of Properties of Meteors; Computing Spacecraft Solar-Cell Damage by Charged Particles; Thermal Model of a Current-Carrying Wire in a Vacuum; Program for Analyzing Flows in a Complex Network; Program Predicts Performance of Optical Parametric Oscillators; Processing TES Level-1B Data; Automated Camera Calibration; Tracking the Martian CO2 Polar Ice Caps in Infrared Images; Processing TES Level-2 Data; SmaggIce Version 1.8; Solving the Swath Segment Selection Problem; The Spatial Standard Observer; Less-Complex Method of Classifying MPSK; Improvement in Recursive Hierarchical Segmentation of Data; Using Heaps in Recursive Hierarchical Segmentation of Data; Tool for Statistical Analysis and Display of Landing Sites; Automated Assignment of Proposals to Reviewers; Array-Pattern-Match Compiler for Opportunistic Data Analysis; Pre-Processor for Compression of Multispectral Image Data; Compressing Image Data While Limiting the Effects of Data Losses; Flight Operations Analysis Tool; Improvement in Visual Target Tracking for a Mobile Robot; Software for Simulating Air Traffic; Automated Vectorization of Decision-Based Algorithms; Grayscale Optical Correlator Workbench; "One-Stop Shopping" for Ocean Remote-Sensing and Model Data; State Analysis Database Tool; Generating CAHV and CAHVOmages with Shadows in ROAMS; Improving UDP/IP Transmission Without Increasing Congestion; FORTRAN Versions of Reformulated HFGMC Codes; Program for Editing Spacecraft Command Sequences; Flight-Tested Prototype of BEAM Software; Mission Scenario Development Workbench; Marsviewer; Tool for Analysis and Reduction of Scientific Data; ASPEN Version 3.0; Secure Display of Space-Exploration Images; Digital Front End for Wide-Band VLBI Science Receiver; Multifunctional Tanks for Spacecraft; Lightweight, Segmented, Mostly Silicon Telescope Mirror; Assistant for Analyzing Tropical-Rain-Mapping Radar Data; and Anion-Intercalating Cathodes for High-Energy- Density Cells.

Source record↗
NASA NTRS1982

An experimental investigation and a numerical prediction of a transonic normal shock wave/turbulent boundary layer interaction

Detailed pitot, static and wall pressure measurements have been obtained for a transonic normal shock wave/turbulent boundary layer interaction at free-stream Mach numbers of 1.28, 1.37 and 1.48 and at a constant unit Reynolds number of 4.92 x 10 to the 6th per meter in an axisymmetric, internal flow. Measurements have also been obtained at a unit Reynolds number of 9.84 x 10 to the 6th per meter at a free-stream Mach number of 1.29. The interaction depends very strongly on the Mach number. The effect of Reynolds number under study is small. Flow confinement due to the wind-tunnel wall boundary layer produces a weaker interaction and a much larger supersonic tongue than is observed for planar flows. Comparisons are made with solutions to the time-dependent, mass-averaged, Navier-Stokes equations incorporating a two-equation, Wilcox-Rubesin turbulence model (1980). The computations are in agreement with the experimental results.

Om, D.↗
NASA NTRS2002

Commercial-Off-The-Shelf Microelectromechanical Systems (MEMS) Flow-Measurement Probes Fabricated And Assembled

As an alternative to conventional tubing instrumentation for measuring airflow, designers and technicians at the NASA Glenn Research Center have been fabricating packaging components and assembling a set of unique probes using commercial-off-the-shelf microelectromechanical systems (MEMS) integrated circuits (computer chips). Using MEMS as an alternative has some compelling advantages over standard measurement devices. Sensor technologies have matured through high-production usage in industries such as automotive and aircraft manufacturers. Currently, MEMS are the choice in applications such as tire pressure monitors, altimeters, pneumatic controls, cable leak detectors, and consumer appliances. Conventional instrumentation uses tubing buried in the model aerodynamic surfaces or wind tunnel walls. The measurements are made when pressure is introduced at the tube opening. The pressure then must travel the tubing for lengths ranging from 20 to hundreds of feet before reaching an electronic signal conditioner. This condition causes a considerable amount of damping and requires measurements to be made only after the test rig has reached steady-state operation. The electronic MEMS pressure sensor is able to take readings continuously under dynamic states in nearly real time. The use of stainless steel tubing for pressure measurements requires many tubes to be cleaned, cut to length, carefully installed, and delicately deburred and spliced for use. A cluster of a few hundred 1/16-in.- (0.0625-in.-) diameter tubes (not uncommon in research testing facilities) can be several inches in diameter and may weigh enough to require two men to handle. Replacing hard tubing with electronic chips can eliminate much of the bulk. Each sensor would fit on the tip of the 1/16-in. tubing with room to spare. The P592 piezoresistive silicon pressure sensor (Lucas NovaSensor, Fremont, CA) was chosen for this project because of its cost, availability, and tolerance to extreme ambient conditions. The chip is 1 mm square by 0.6 mm thick (0.039 by 0.023 in.) with 0.12-mm (0.005-in.) wire connection tabs. Three MEMS chips were used to build the first type of flow-angularity probe. This MEMS probe will be demonstrated as an alternative to a standard tube type "Cobra Probe" now used routinely in wind tunnel and aeronautical hardware applications. The response time and accuracy would allow the probe to be translated on an actuator across a flow field, yielding precision dynamic measurements not possible with conventional instrumentation. The low profile, the minimal power requirement, the rugged construction, and the moderate cost all contribute to making MEMS sensors the enticing choice instrument in future research measurement needs. The MEMS probe efforts are a continuation of work initiated by Brian Willis, without whose foresight and efforts this project would never have been realized. This task was funded through cooperation with the NASA Electronic Parts and Packaging (NEPP) program at the Jet Propulsion Laboratory.

Redding, Chip↗
NASA NTRS2004

COTS MEMS Flow-Measurement Probes

As an alternative to conventional tubing instrumentation for measuring airflow, designers and technicians at Glenn Research Center have been fabricating packaging components and assembling a set of unique probes that contain commercial off-the-shelf (COTS) microelectromechanical systems (MEMS) sensor chips. MEMS sensor chips offer some compelling advantages over standard macroscopic measurement devices. MEMS sensor technology has matured through mass production and use in the automotive and aircraft industries. At present, MEMS are the devices of choice for sensors in such applications as tire-pressure monitors, altimeters, pneumatic controls, cable leak detectors, and consumer appliances. Compactness, minimality of power demand, rugged construction, and moderate cost all contribute to making MEMS sensors attractive for instrumentation for future research. Conventional macroscopic flow-measurement instrumentation includes tubes buried beneath the aerodynamic surfaces of wind-tunnel models or in wind-tunnel walls. Pressure is introduced at the opening of each such tube. The pressure must then travel along the tube before reaching a transducer that generates an electronic signal. The lengths of such tubes typically range from 20 ft (approx.= 6 m) to hundreds of feet (of the order of 100 m). The propagation of pressure signals in the tubes damps the signals considerably and makes it necessary to delay measurements until after test rigs have reached steady-state operation. In contrast, a MEMS pressure sensor that generates electronic output can take readings continuously under dynamic conditions in nearly real time. In order to use stainless-steel tubing for pressure measurements, it is necessary to clean many tubes, cut them to length, carefully install them, delicately deburr them, and splice them. A cluster of a few hundred 1/16-in.- (approx.=1.6-mm-) diameter tubes (such clusters are common in research testing facilities) can be several inches (of the order of 10 cm) in diameter and could weigh enough that two technicians are needed to handle it. Replacing hard tubing with electronic chips can eliminate much of the bulk. Each sensor would fit on the tip of a 1/16-in. tube with room to spare. The Lucas NovaSensor P592 piezoresistive silicon pressure sensor was chosen for this project because of its cost, availability, and tolerance to extreme ambient conditions. The sensor chip is 1 mm square by 0.6 mm thick (about 0.039 by 0.039 by 0.024 in.) and includes 0.12-mm (approx.=0.005-in.) wire connection tabs. The figure shows a flow-angularity probe that was built by use of three such MEMS chips. It is planned to demonstrate this MEMS probe as an alternative to a standard tube-type "Cobra" probe now used routinely in wind tunnels and aeronautical hardware. This MEMS probe could be translated across a flow field by use of a suitable actuator, so that its accuracy and the shortness of its response time could be exploited to obtain precise dynamic measurements of a sort that cannot be made by use of conventional tubing-based instrumentation.

Redding, Chip↗