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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 253 records · Page 14
All sourcesNASA NTRSDOE OSTI
NASA NTRS1985

Static and dynamic pressure measurements on a NACA 0012 airfoil in the Ames High Reynolds Number Facility

The supercritical flows at high subsonic speeds over a NACA 0012 airfoil were studied to acquire aerodynamic data suitable for evaluating numerical-flow codes. The measurements consisted primarily of static and dynamic pressures on the airfoil and test-channel walls. Shadowgraphs were also taken of the flow field near the airfoil. The tests were performed at free-stream Mach numbers from approximately 0.7 to 0.8, at angles of attack sufficient to include the onset of buffet, and at Reynolds numbers from 1 million to 14 million. A test action was designed specifically to obtain two-dimensional airfoil data with a minimum of wall interference effects. Boundary-layer suction panels were used to minimize sidewall interference effects. Flexible upper and lower walls allow test-channel area-ruling to nullify Mach number changes induced by the mass removal, to correct for longitudinal boundary-layer growth, and to provide contouring compatible with the streamlines of the model in free air.

Mcdevitt, J. B.↗
NASA NTRS1994

Miniature, Cooled Pressure-Measuring Probe

Probe designed to reduce settling time dramatically. Pressure-sensing transducer mounted in probe and connected to tip by short tube having cross-sectional area substantially smaller than conventional connecting tubes. Probe includes stainless-steel cylindrical exterior housing holding closed pressure chamber in which piezoelectric pressure transducer mounted. Open connecting tube passes portion of high-velocity, high-temperature fluid stream into closed pressure chamber. Any change of pressure in sampled stream propagates into closed pressure chamber with settling time inversely proportional to cross-sectional area of connecting tube. Cooling chamber formed around pressure chamber connected to source of water or other cooling fluid via inlet and outlet tubes.

Ashby, George C., Jr.↗
NASA NTRS1994

Steady pressure measurements on an Aeroelastic Research Wing (ARW-2)

Transonic steady and unsteady pressure tests have been conducted in the Langley transonic dynamics tunnel on a large elastic wing known as the DAST ARW-2. The wing has a supercritical airfoil, an aspect ratio of 10.3, a leading-edge sweep back angle of 28.8 degrees, and two inboard and one outboard trailing-edge control surfaces. Only the outboard control surface was deflected to generate steady and unsteady flow over the wing during this study. Only the steady surface pressure, control-surface hinge moment, wing-tip deflection, and wing-root bending moment measurements are presented. The results from this elastic wing test are in tabulated form to assist in calibrating advanced computational fluid dynamics (CFD) algorithms.

Sandford, Maynard C.↗
NASA NTRS1974

Sonic-boom ground pressure measurements from the launch and reentry of Apollo 16

Sonic-boom pressure signatures recorded during the launch and reentry phases of the Apollo 16 mission are presented. Five measurements were obtained along the vehicle ground track: 69 km (37.3 n. mi.) 92 km (49.8 n. mi.), and 130 km (70.3 n. mi.) down range from the launch site during ascent, and at 185 km (100 n. mi.) and approximately 5.5 km (3 n. mi.) from the splash-down point during reentry. Tracings of the measured signatures are included along with values of the overpressure, impulse, time duration, and rise times. Also included are brief descriptions of the launch and recovery test areas in which the measurements were obtained, the sonic-boom instrumentation deployment, flight profiles, and operating conditions for the launch vehicle and spacecraft, surface weather information at the measuring sites, and high-altitude weather information for the general measurement areas. Comparisons of the sonic-boom overpressures from Apollo 15 and 16 along with those from current aircraft are also presented.

Henderson, H. R.↗
NASA NTRS2023

A Simple Boundary-Layer Transition Detection Method Using Mean Static Pressure Measurements

A simple method for detecting boundary-layer transition using only mean static pressure port data is presented. The method can be applied to most existing models with pressure taps, and only requires that a fine angle-of-attack sweep be performed. A small but abrupt change in the static pressure is visible when the transition front passes over the pressure tap. Results from a recent Juncture Flow test entry are used to illustrate the technique. Infrared thermography measurements of the transition front compare very well to the transition locations obtained from the static pressure ports. Some differences in behavior occur depending on the dominant transition mechanism. While the technique is somewhat qualitative, it can be an excellent tool for estimating the transition location when other tools are not readily available. Given the simplicity of the technique, and the fact that most wind-tunnel models are already designed with numerous static pressure taps, this method can be applied with very little overhead.

boundary-layer transition↗
NASA NTRS1922

Pressure Measurements During Flight

The general idea was to record the air pressure on the spot, that is, without intervening tubes, by some simple automatic registering device. This was therefore made in such small dimensions that it could be inserted into the wing between the ribs and stand in direct connection with the pressure openings. The results of different experiments are presented along with the best locations for locating the instruments for recording the pressure.

Proll, A↗
NASA NTRS1984

Wall pressure measurements for three-dimensional transonic tests

An experiment is described that provides input data for wind-tunnel wall interference assessment methods that are based on test-section wall pressure distributions. Wall pressures have been measured along orifice rows on the test-section walls during longitudinal force tests on two model sizes of the same configuration. These data were acquired at Mach numbers between 0.60 and 0.90 in a small, atmospheric transonic wind tunnel. A sample of the data and a discussion of results are presented.

Sewall, W. G.↗
NASA NTRS1978

Flight-measured pressure characteristics of aft-facing steps in high Reynolds number flow at Mach numbers of 2.20, 2.50, and 2.80 and comparison with other data

The YF-12 airplane was studied to determine the pressure characteristics associated with an aft-facing step in high Reynolds number flow for nominal Mach numbers of 2.20, 2.50, and 2.80. Base pressure coefficients were obtained for three step heights. The surface static pressures ahead of and behind the step were measured for the no-step condition and for each of the step heights. A boundary layer rake was used to determine the local boundary layer conditions. The Reynolds number based on the length of flow ahead of the step was approximately 10 to the 8th power and the ratios of momentum thickness to step height ranged from 0.2 to 1.0. Base pressure coefficients were compared with other available data at similar Mach numbers and at ratios of momentum thickness to step height near 1.0. In addition, the data were compared with base pressure coefficients calculated by a semiempirical prediction method. The base pressure ratios are shown to be a function of Reynolds number based on momentum thickness. Profiles of the surface pressures ahead of and behind the step and the local boundary layer conditions are also presented.

Powers, S. G.↗
NASA NTRS1998

Noninvasive Intracranial Volume and Pressure Measurements Using Ultrasound

Prevention of secondary brain injuries following head can be accomplished most easily when intracranial pressure (ICP) is monitored. However, current measurement techniques are invasive and thus not practical in the combat environment. The Pulsed Phase Lock Loop (PPLL) devise, which was developed and patented, uses a unique, noninvasive ultrasonic phase comparison method to measure slight changes in cranial volume which occur with changes in ICP. Year one studies involved instrument improvements and measurement of altered intracranial distance with altered ICP in fresh cadavera. Our software was improved to facilitate future studies of normal subjects and trauma patients. Our bench studies proved that PPLL output correlated highly with changes in path length across a model cranium. Cadaveric studies demonstrated excellent compact, noninvasive devise for monitoring changes in intracranial distance may aid in the early detection of elevated ICP, decreasing risk of secondary brain injury and infection, and returning head-injured patients to duty.

Hargens, A. R.↗
NASA NTRS1982

Measured pressure distributions, aerodynamic coefficients and shock shapes on blunt bodies at incidence in hypersonic air and CF4

Pressure distributions, aerodynamic coefficients, and shock shapes were measured on blunt bodies of revolution in Mach 6 CF4 and in Mach 6 and Mach 10 air. The angle of attack was varied from 0 deg to 20 deg in 4 deg increments. Configurations tested were a hyperboloid with an asymptotic angle of 45 deg, a sonic-corner paraboloid, a paraboloid with an angle of 27.6 deg at the base, a Viking aeroshell generated in a generalized orthogonal coordinate system, and a family of cones having a 45 deg half-angle with spherical, flattened, concave, and cusp nose shapes. Real-gas effects were simulated for the hperboloid and paraboloid models at Mach 6 by testing at a normal-shock density ratio of 5.3 in air and 12 CF4. Predictions from simple theories and numerical flow field programs are compared with measurement. It is anticipated that the data presented in this report will be useful for verification of analytical methods for predicting hypersonic flow fields about blunt bodies at incidence.

Miller, C. G., III↗
NASA NTRS1985

Unsteady pressure measurements on a biconvex airfoil in a transonic oscillating cascade

Flush-mounted dynamic pressure transducers were installed on the center airfoil of a transonic oscillating cascade to measure the unsteady aerodynamic response as nine airfroils were simultaneously driven to provide 1.2 deg of pitching motion about the midchord. Initial tests were performed at an incidence and angle of 0 deg and A Mach number of 0.65 in order to obtain results in a shock-free compressible flowfield. Subsequent tests were performed at an incidence angle of 7 deg and Mach number of 0.8 in order to observe the surface pressures with an oscillating shock near the leading edge of the airfoil. Results are presented for interblade phase angles of 90 and -90 deg and at blade oscillatory frequencies of 200 and 500 Hz (semi-chord reduced frequencies up to about 0.5 at a Mach number of 0.8). Results from the zero-incidence cascade are compared with a classical unsteady flat-plate analysis. Flow visualization results depicting the shock motion on the airfoils in the high-incidence cascade are discussed. The airfoil pressure data are tabulated.

Shaw, L. M.↗
NASA NTRS1986

Unsteady pressure measurements on a biconvex airfoil in a transonic oscillating cascade

Flush-mounted dynamic pressure transducers were installed on the center airfoil of a transonic oscillating cascade to measure the unsteady aerodynamic response as nine airfols were simultaneously driven to provide 1.2 deg of pitching motion about the midchord. Initial tests were performed at an incidence and angle of 0 deg and A Mach number of 0.65 in order to obtain results in a shock-free compressible flowfield. Subsequent tests were performed at an incidence angle of 7 deg and Mach number of 0.8 in order to observe the surface pressures with an oscillating shock near the leading edge of the airfoil. Results are presented for interblade phase angles of 90 and -90 deg and at blade oscillatory frequencies of 200 and 500 Hz (semi-chord reduced frequencies up to about 0.5 at a Mach number of 0.8). Results from the zero-incidence cascade are compared with a clasical usnteady flat-plate analysis. Flow visualization results depicting the shock motion on the airfoils in the high-incidence cascade are discussed. The airfoil pressure data are tabulated.

Shaw, L. M.↗
NASA NTRS1991

Heat transfer and pressure measurements for the SSME fuel turbine

A measurement program is underway using the Rocketdyne two-stage Space Shuttle Main Engine (SSME) fuel turbine. The measurements use a very large shock tunnel to produce a short-duration source of heated and pressurized gas which is subsequently passed through the turbine. Within this environment, the turbine is operated at the design values of flow function, stage pressure ratio, stage temperature ratio, and corrected speed. The first stage vane row and the first stage blade row are instrumented in both the spanwise and chordwise directions with pressure transducers and heat flux gages. The specific measurements to be taken include time averaged surface pressure and heat flux distributions on the vane and blade, flow passage static pressure, flow passage total pressure and total temperature distributions, and phase resolved surface pressure and heat flux on the blade.

Dunn, Michael G.↗
NASA NTRS1976

Force and pressure measurements on an airfoil oscillating through stall

Details of force, moment, and pressure distributions on a two-dimensional, 4-ft chord, NACA 0012 airfoil, oscillating in pitch through stall, in a 7 ft x 10 ft low-speed wind tunnel are presented. Tests were run with the airfoil in a closed test section and also in a test section having four longitudinal slots in each sidewall set to provide minimum tunnel interference on the wing in steady flow. In steady flow, differences between the results for the closed and 2% open case are small. The dynamic stall process is not triggered by the bursting of a laminar separation bubble but rather by the separation of the turbulent boundary layer downstream of the bubble.

Parker, A. G.↗
NASA NTRS1951

Characteristics of a wedge with various holder configurations for static-pressure measurements in subsonic gas streams

The characteristics of a wedge static-pressure sensing element with various holder configurations were determined and compared with the characteristics of the conventional tube. The probes were tested over a range of Mach number from 0.3 to 0.95 and at various pitch and yaw angles. The investigation showed that the spike-mounted wedge sensing element has a pressure coefficient comparable with the conventional subsonic static-pressure probe and the pressure coefficient of the wedge varied less than that of the conventional probe for corresponding change of yaw angle.

Gettelman, Clarence C↗
NASA NTRS2001

Pressure Probe Designs for Dynamic Pressure Measurements in a Supersonic Flow Field

A series of dynamic flow field pressure probes were developed for use in large-scale supersonic wind tunnels at NASA Glenn Research Center. These flow field probes include pitot, static, and five-hole conical pressure probes that are capable of capturing fast acting flow field pressure transients that occur on a millisecond time scale. The pitot and static probes can be used to determine local Mach number time histories during a transient event. The five-hole conical pressure probes are used primarily to determine local flow angularity, but can also determine local Mach number. These probes were designed, developed, and tested at the NASA Glenn Research Center. They were also used in a NASA Glenn 10- by 10-Foot Supersonic Wind Tunnel (SWT) test program where they successfully acquired flow field pressure data in the vicinity of a propulsion system during an engine compressor stall and inlet unstart transient event. Details of the design, development, and subsequent use of these probes are discussed in this report.

Porro, A. Robert↗