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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 883 records · Page 49

All sources NASA NTRS DOE OSTI

NASA NTRS1991

Computational Fluid Dynamics In Aerospace

Report reviews recent progress in computational fluid dynamics (CFD). Addresses state of art in terms of methods, examples of solutions, and computer technology. Provides examples of CFD solutions of aerospace and fluid-physics problems to illustrate recent progress. Aerospace solutions include rotor/stator velocity vectors from three-dimensional Navier-Stokes code for turbomachinery, pressure distributions from simulation of aeroelasticity of wing surface-pressure contours on four-bladed helicopter rotor in forward flight, and particle traces about several aerospace vehicles. Computer technology covers supercomputers, scientific work stations, local supercomputing networks, and remote access to supercomputers.

Kutler, Paul↗

NASA NTRS1993

An automated pressure data acquisition system for evaluation of pressure sensitive paint chemistries

An automated pressure data acquisition system for testing of pressure sensitive phosphorescent paints was designed, assembled, and tested. The purpose of the calibration system is the evaluation and selection of pressure sensitive paint chemistries that could be used to obtain global aerodynamic pressure distribution measurements. The test apparatus and setup used for pressure sensitive paint characterizations is described. The pressure calibrations, thermal sensitivity effects, and photodegradation properties are discussed.

Sealey, Bradley S.↗

NASA NTRS2004

"We Burn to Learn" About Fuel-Air Mixing Within Aircraft Powerplants

I am working with my branch s advanced diagnostics team to investigate fuel-air mixing in jet-fueled gas turbine combustors and jet-fuel reformers. Our data acquisition begins with bench-top experiments which will help with calibration of equipment for facility testing. While conducting the bench-top experiments I learned to align laser and optical equipment to collect data, to use the data acquisition software, and to process the data into graphs and images. which jet he1 is to be reformed into hydrogen. Testing will commence shortly, after which we will obtain and analyze data and meet a critical milestone for the end of September. I am also designing the layout for a Schlieren system that will be used during that time frame. A Schlieren instrument records changes in the refractive index distribution of transparent media like air flows. The refractive index distribution can then be related to density, temperature, or pressure distributions within the flow. I am working on a scheme to quantify this information and add to the knowledge of the fuel-air mixing process.

Robinson, Heidi N.↗

NASA NTRS2013

Artificial Intelligence Software for Assessing Postural Stability

A software package reads and analyzes pressure distributions from sensors mounted under a person's feet. Pressure data from sensors mounted in shoes, or in a platform, can be used to provide a description of postural stability (assessing competence to deficiency) and enables the determination of the person's present activity (running, walking, squatting, falling). This package has three parts: a preprocessing algorithm for reading input from pressure sensors; a Hidden Markov Model (HMM), which is used to determine the person's present activity and level of sensing-motor competence; and a suite of graphical algorithms, which allows visual representation of the person's activity and vestibular function over time.

Lieberman, Erez↗

NASA NTRS1972

Coolant pressure and airflow distribution in a strut-supported transpiration-cooled vane for a gas turbine engine

An analysis to predict pressure and flow distribution in a strut-supported wire-cloth vane was developed. Results were compared with experimental data obtained from room-temperature airflow tests conducted over a range of vane inlet airflow rates from 10.7 to 40.4 g/sec (0.0235 to 0.0890 lb/sec). The analytical method yielded reasonably accurate predictions of vane coolant flow rate and pressure distribution.

Kaufman, A.↗

NASA NTRS1950

High-Subsonic Performance Characteristics and Boundary-Layer Investigations of a 12 10-Inch-Inlet-Diameter Conical Diffuser

Performance and boundary-layer data were taken in a 12 degree 10-inch inlet-diameter conical diffuser of 2:1 exit- to inlet-area ratio. These data were taken for two inlet-boundary-layer conditions. The first condition was that of a thinner inlet boundary later (boundary-layer displacement thickness, delta* approximately equal to 0.034) produced by an inlet section approximately 1 inlet diameter in length between the entrance bell and the diffuser. The second condition was a thicker inlet boundary layer (delta* approximately equal to 0.120) produced by an additional inlet section length of approximately 6 diameters. Longitudinal static-pressure distributions were measured fro wall static orifices. Transverse total- and static-pressure surveys were made at the inlet and exit stations. Boundary-layer velocity distributions were measured at seven stations between the inlet and exit. These data were obtained for a Reynolds number (based on inlet diameter) range of 1 x 10(exp 6) to 3.9 x 10(exp 6). The corresponding Mach number range was from M = 0.2 to choking. At the maximum-power-available condition supersonic flow was obtained as far as 4.5 inches downstream from the diffuser inlet with a maximum Mach number of M approximately equal to 1.5. The total-pressure loss through the diffuser in percentage of inlet dynamic pressure was approximately 2.5 percent for the thinner inlet boundary later and 5.5 percent for the thicker inlet boundary later over the lower subsonic range. These valued increased with increasing flow rate- the values for the thicker inlet boundary later more than those for the thinner inlet boundary layer. The diffuser effectiveness, expressed as the ratio of the actual static-pressure rise to the ideal static-pressure rise, was about 85 percent for the thinner inlet boundary layer and about 67 percent for the thicker inlet boundary later in the lower subsonic range. These values decrease with increasing flow rate. Separated flow was observed for both inlet-boundary-layer conditions in the region of adverse pressure gradient just downstream of the transition curvature from inlet section to diffuser. The flow for the thinner-inlet-boundary-layer condition did not fully re-establish itself along the diffuser walls. The thicker inlet-boundary-layer flow, while not completely re-establishing the normal flow pattern downstream of the separated region, did re-establish more successfully than the thinner inlet boundary layer.

Little, B. H., Jr.↗

NASA NTRS1975

Improvement of propeller static thrust estimation

The problem of improving the performance estimation of propellers operating in the heavily loaded static thrust condition was studied. The Goldstein theory was assessed as it applies to propellers operating in the static thrust. A review of theoretical considerations is presented along with a summary of the attempts made to obtain a numerical solution. The chordwise pressure distribution was determined during operation at a tip speed of 500 ft/sec. Chordwise integration of the pressures leads to the spanwise load distribution and further integration would give the axial thrust.

Brusse, J.↗

NASA NTRS1980

Stress distributions around furrows and asperities in EHL line contacts

The effect on the stresses, pressure, and film thickness profile of a transversely oriented asperity ridge and furrow located on the surface of a roller in a lubricated, simple sliding, line contact is investigated. The location of these defects in the contact region, the lubricant parameters, the minimum film thickness, and the combination of tangential and normal loads are varied and their effects on the pressures, film thicknesses and stresses are plotted and discussed. It is found that the pressure distribution in the contact, and hence the stress pattern in the roller, suffers a significant change due to the presence of the asperity ridge or furrow.

Cheng, H. S.↗

DOE OSTI2021

Design of a Single Layer Metamaterial for Pressure Offloading of Transtibial Amputees

While using a prosthesis, transtibial amputees may experience pain and discomfort brought on by large pressure gradients at the interface between the residual limb and prosthetic socket. Current prosthetic interface solutions attempt to alleviate these pressure gradients by using soft homogenous liners to reduce and distribute pressures. This research investigates an additively manufactured metamaterial inlay with a tailored mechanical response to reduce peak pressure gradients around the limb. The inlay uses a hyperelastic behaving metamaterial (US10244818) comprised of triangular pattern unit cells which can be 3D printed with walls of various thicknesses controlled by draft angles. The hyperelastic material properties are modeled using a Yeoh 3rd order model. The 3rd order coefficients can be adjusted and optimized, which corresponds to a change in the unit cell wall thickness to create an inlay that can meet the unique offloading needs of an amputee. Finite element analysis simulations evaluated the pressure gradient reduction from: 1) a common homogenous silicone liner, 2) a prosthetist's inlay prescription that utilizes three variations of the metamaterial, and 3) a metamaterial solution with optimized Yeoh 3rd order coefficients. When compared to a traditional homogenous silicone liner for two unique limb loading scenarios, the prosthetist prescribed inlay and the optimized material inlay can achieve equal or greater pressure gradient reduction capabilities. These results show the potential feasibility of implementing this metamaterial as a method of personalized medicine for transtibial amputees by creating a customizable interface solution to meet the unique performance needs of an individual patient.

60 APPLIED LIFE SCIENCES↗

NASA NTRS1983

Test Report for MSFC Test No. 83-2: Pressure scaled water impact test of a 12.5 inch diameter model of the Space Shuttle solid rocket booster filament wound case and external TVC PCD

Water impact tests using a 12.5 inch diameter model representing a 8.56 percent scale of the Space Shuttle Solid Rocket Booster configuration were conducted. The two primary objectives of this SRB scale model water impact test program were: 1. Obtain cavity collapse applied pressure distributions for the 8.56 percent rigid body scale model FWC pressure magnitudes as a function of full-scale initial impact conditions at vertical velocities from 65 to 85 ft/sec, horizontal velocities from 0 to 45 ft/sec, and angles from -10 to +10 degrees. 2. Obtain rigid body applied pressures on the TVC pod and aft skirt internal stiffener rings at initial impact and cavity collapse loading events. In addition, nozzle loads were measured. Full scale vertical velocities of 65 to 85 ft/sec, horizontal velocities of 0 to 45 ft/sec, and impact angles from -10 to +10 degrees simulated.

Source record↗

NASA NTRS1985

Water impact testing of a filament wound case

A lightweight Filament Wound Case (FWC) Solid Rocket Booster (SRB) is being developed by NASA to increase the payload capability of the space shuttle. As with the steel boosters, the current plan is to recover the FWC SRB's after they impact the ocean at 65 to 85 ft/sec. As the boosters enter the ocean (nozzle first) the water moves away from the vehicle creating a cavity, which then collapses on the vehicle, and results in a significant loading event. To understand this loading event, tests were conducted on a quarter scale FWC model to measure cavity collapse pressure distributions, deflected shape and the effects of end conditions and pressure scaling.

Schmidt, A. A.↗

NASA NTRS1988

Application of finite element and remeshing technique to shock interference on a cylindrical leading edge

The problem of planar oblique shock impingement on a cylindrical leading edge in hypersonic flow is modeled using a Galerkin-Runge Kutta finite element method. The method utilizes a four stage Runge-Kutta time stepping scheme to solve the compressible Euler equations. Freestream Mach numbers of 6.5, 8.0 and 16.0 are studied. The computed surface pressure distributions consistently agree well with available experimental data. The peak pressure amplification ranges from 5.45 at M = 6.5 to approximately 17.0 at M = 16.0. Stagnation point heat transfer rate amplifications are calculated from the inviscid solution using the method of Fay and Riddell. The value and wall location of the peak pressure and heat transfer rate amplifications are extremely sensitive to the location of the impinging shock/bow shock intersection point.

Stewart, James R.↗

NASA NTRS1989

Fluctuating pressures measured beneath a high-temperature, turbulent boundary layer on a flat plate at Mach number of 5

Fluctuating pressures were measured beneath a Mach 5, turbulent boundary layer on a flat plate with an array of piezoresistive sensors. The data were obtained with a digital signal acquisition system during a test run of 4 seconds. Data sampling rate was such that frequency analysis up to 62.5 kHz could be performed. To assess in situ frequency response of the sensors, a specially designed waveguide calibration system was employed to measure transfer functions of all sensors and related instrumentation. Pressure time histories were approximated well by a Gaussian prohibiting distribution. Pressure spectra were very repeatable over the array span of 76 mm. Total rms pressures ranged from 0.0017 to 0.0046 of the freestream dynamic pressure. Streamwise, space-time correlations exhibited expected decaying behavior of a turbulence generated pressure field. Average convection speed was 0.87 of freestream velocity. The trendless behavior with sensor separation indicated possible systematic errors.

Parrott, Tony L.↗

NASA NTRS1992

Unsteady separation in sharp fin-induced shock wave/turbulent boundary layer interaction at Mach 5

Fluctuating wall-pressure measurements are made in shock-wave/turbulent-boundary-layer interactions generated by sharp/unswept fins at angles of attack of 16, 18, 20, 22, 24, 26, and 28 degrees at Mach 5. The experiment was conducted under approximately adiabatic wall temperature conditions. The mean and rms pressure distributions can be collapsed in conical coordinates. The wall-pressure signal near separation is intermittent for all angles of attack (16-28 deg) and is qualitatively similar to that measured in unswept flows. However, the shock frequencies are higher - about 5 kHz compared to 0.5-1 kHz. Over the range of sweepbacks examined, from 25-55 deg, the spectral content of the fluctuating pressures does not change. Thus, the increase in separation-shock frequency from 1 to 5 kHz occurs at lower interaction sweepback and is not a continuous process with increasing sweepback. Power spectra at the position of maximum rms in the intermittent region for interactions in different incoming boundary layers have the same center frequency. The maximum rms in the intermittent region correlates with interaction sweepback, not with overall inviscid pressure rise.

Schmisseur, J. D.↗

NASA NTRS1937

Fuselage-drag tests in the variable-density wind tunnel: streamline bodies of revolution, fineness ratio of 5

Results are presented of the drag tests of six bodies of revolution with systematically varying shapes and with a fineness ratio of 5. The forms were derived from source-sink distributions, and formulas are presented for the calculation of the pressure distribution of the forms. The tests were made in the N.A.C.A. variable-density tunnel over a range of values of Reynolds number from about 1,500,000 to 25,000,000. The results show that the bodies with the sharper noses and tails have the lowest drag coefficients, even when the drag coefficients are based on the two-thirds power of the volume. The data shows the most important single characteristic of the body form to be the tail angle, which must be fine to obtain low drag.

Abbott, Ira H↗

NASA NTRS1993

The 3-D CFD modeling of gas turbine combustor-integral bleed flow interaction

An advanced 3-D Computational Fluid Dynamics (CFD) model was developed to analyze the flow interaction between a gas turbine combustor and an integral bleed plenum. In this model, the elliptic governing equations of continuity, momentum and the k-e turbulence model were solved on a boundary-fitted, curvilinear, orthogonal grid system. The model was first validated against test data from public literature and then applied to a gas turbine combustor with integral bleed. The model predictions agreed well with data from combustor rig testing. The model predictions also indicated strong flow interaction between the combustor and the integral bleed. Integral bleed flow distribution was found to have a great effect on the pressure distribution around the gas turbine combustor.

Chen, D. Y.↗

NASA NTRS1977

Rolling moments in a trailing vortex flow field

Pressure distributions are presented which were measured on a wing in close proximity to a tip vortex of known structure generated by a larger, upstream semispan wing. Overall loads calculated by integration of these pressures are checked by independent measurements made with an identical model mounted on a force balance. Several conventional methods of wing analysis are used to predict the loads on the following wing. Strip theory is shown to give uniformly poor results for loading distribution, although predictions of overall lift and rolling moment are sometimes acceptable. Good results are obtained for overall coefficients and loading distribution by using linearized pressures in vortex-lattice theory in conjunction with a rectilinear vortex. The equivalent relation from reverse-flow theory that can be used to give economic predictions for overall loads is presented.

Mcmillan, O. J.↗

NASA NTRS1977

Rolling moments in a trailing vortex flow field

Mcmillan, O. J.↗