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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 NTRS1947

Force and Pressure-Distribution Measurements on a Rectangular Wing with Double-Hinged Nose

The previous measurements on airfoils with hinged nose disclosed a comparatively large low-pressure peak at the bend of the hinged nose; which favored the separation of flow. It was therefore attempted to reduce these low-pressure peaks by reducing the camber of the forward profile and thereby ensure a longer adherence of the flow and a maximum lift increase. The forces were measured on a rectangular wing with double-hinged nose and end plates, the pressure distributions were measured in the center section of the wing. The measurements disclosed that the highest lift attained with a single-hinged nose cannot be increased by a double-hinged nose. The sum of the deflection angles of both hinged noses related to the maximum lift is about equal to the corresponding angle of the single-hinge nose (approx. 30 deg to 40). The respective angle of attack in both cases amounts to approx. 21 deg. Even the low-pressure peak is about the same in both cases (P/q approx. -5.5). Therefore, a milder curvature of the forward portion of the profile affords no definite increase of the maximum lift.

Lemme, H. A.↗
NASA NTRS1976

Influence of surface roughness and waviness on film thickness and pressure distribution in elastohydrodynamic contacts

The Christensen theory of a stochastic model for hydrodynamic lubrication of rough surfaces was extended to elastohydrodynamic lubrication between two rollers. Solutions for the reduced pressure at the entrance as a function of the ratio of the average nominal film thickness to the rms surface roughness, were obtained numerically. Results were obtained for purely transverse as well as purely longitudinal surface roughness for cases with or without slip. The reduced pressure was shown to decrease slightly by considering longitudinal surface roughness. The same approach was used to study the effect of surface roughness on lubrication between rigid rollers and lubrication of an infinitely wide slider bearing. Using the flow balance concept, the perturbed Reynolds equation, was derived and solved for the perturbed pressure distribution. In addition, Cheng's numerical scheme was modified to incorporate a single two-dimensional elastic asperity on the stationary surface. The perturbed pressures obtained by these three different models were compared.

Chow, L. S. H.↗
NASA NTRS1973

Effect of sweep angle on the pressure distributions and effectiveness of the ogee tip in diffusing a line vortex

Low-speed wind tunnel tests were conducted to study the influence of sweep angle on the pressure distributions of an ogee-tip configuration with relation to the effectiveness of the ogee tip in diffusing a line vortex. In addition to the pressure data, performance and flow-visualization data were obtained in the wind tunnel tests to evaluate the application of the ogee tip to aircraft configurations. The effect of sweep angle on the performance characteristics of a conventional-tip model, having equivalent planform area, was also investigated for comparison with the ogee-tip configuration. Results of the investigation generally indicate that sweep angle has little effect on the characteristics of the ogee in diffusing a line vortex.

Balcerak, J. C.↗
NASA NTRS1929

Pressure Distribution on a Slotted R.A.F. 31 Airfoil in the Variable Density Wind Tunnel

Measurements were made in the variable density wind tunnel to determine the pressure distribution over one section of a R.A.F. 31 airfoil with a leading edge slot fully open. To provide data for the study of scale effect on this type of airfoil, the tests were conducted with air densities of approximately one and twenty atmospheres.

Jacobs, Eastman N↗
NASA NTRSDate not supplied

Simulation of Vacuum Chamber Pressure Distribution with Surrogate Modeling and Uncertainty Quantification

A major challenge in understanding differences in electric propulsion performance in ground tests and in space operations concerns the pressure distribution within the test vacuum chamber. The chamber backpressure is much higher than experienced in space, modifying thruster performance and plume dynamics. Numerical simulation is a key element to determining the background conditions in non-ideal vacuum chamber environments. An important parameter for the accurate simulation of chamber backpressure is the probability that an atom will stick to a cryogenic panel used to pump away the plume gases. This quantity can be used to model vacuum pumps in particle-based kinetic numerical methods. In this work, a three-dimensional direct simulation Monte Carlo code is used to model neutral xenon atoms flowing from the anode of the H9 Hall Effect Thruster within the University of Michigan’s Large Vacuum Test Facility. Simulated pressures are compared with ion gauge pressure measurements to infer the effective sticking coefficient of the chamber’s vacuum pumps. A pressure predicting surrogate model is developed for inference of pump sticking coefficients and for uncertainty quantification. This information enables accurate and useful kinetic simulations of electric propulsion thruster plasma plumes in vacuum chambers.

DSMC↗
NASA NTRS1983

Pressure distributions obtained on a 0.04-scale and 0.02-scale model of the Space Shuttle Orbiter's forward fuselage in the Langley continuous flow hypersonic tunnel

Results from pressure distribution tests on 0.04-scale and 0.02-scale models of the forward fuselage of the Space Shuttle Orbiter are presented without analysis. The tests wre completed in the Langley Continuous Flow Hypersonic Tunnel (CFHT). The 0.04-scale model was tested at angles of attack from -5 deg to 45 deg and angles of sideslip from -3 deg to 3 deg. The 0.02-scale model was tested at angles of attack from -10 deg to 45 deg and angles of sideslip from -5 deg to 5 deg. The tests were conducted in support of the development of the Shuttle Entry Air Data System (SEADS). In addition to modeling the 20 SEADS pressure orifices, the wind-tunnel models were also instrumented with orifices to match Development Flight Instrumentation (DFI) port locations currently existing on the Space Shuttle Orbiter Columbia (OV-102). This DFI simulation has provided a means for comparisons between reentry flight pressure data and wind-tunnel data.

Bradley, P. F.↗
NASA NTRS1975

Investigation of the unsteady pressure distribution on the blade of an axial flow fan

A major source of the noise generated by high bypass ratio aircraft fan engines is the unsteady pressures (forces and moments) acting on the fan blades. These unsteady pressures are the result of the interaction of the blades with the wakes of upstream blade rows, inlet flow distortions, and inlet turbulence. An experimental system was developed which will permit the measurement of the unsteady pressure distributions; a series of experiments were conducted to investigate these effects, and the results are compared with predictions from existing theoretical analyses.

Henderson, R. E.↗
NASA NTRS1983

Internal pressure distributions for a two-dimensional thrust-reversing nozzle operating at a free-stream Mach number of zero

An investigation was conducted in the static test facility of the Langley 16-Foot Transonic Tunnel to measure static pressure distributions inside a nonaxisymmetric thrust reversing nozzle. The tests were made at nozzle total pressures ranging from ambient to about eight times ambient pressure at a free stream Mach number of zero. Tabulated pressure data are presented.

Putnam, L. E.↗
NASA NTRS1929

Pressure Distribution Tests on PW-9 Wing Models from -18 Degree Through 90 Degree Angle of Attack

At the request of the Army Air Corps, an investigation of the pressure distribution over PW-9 wing models was conducted in the atmospheric wind tunnel of the National Advisory Committee for Aeronautics. The primary purpose of these tests was to obtain wind-tunnel data on the load distribution on the cellule to be correlated with similar information obtained in flight tests, both to be used for design purposes. Because of the importance of the conditions beyond the stall as affecting the control and stability, this investigation was extended through 90 degree angle of attack. The results for the range of normal flight have been given in NACA Technical Report No. 271. The present paper presents the same results in a different form and includes, in addition, those over the greater range of angle of attack, -18 degrees through 90 degrees. The results show that: (1) at angles of attack above maximum lift, the biplane upper wing pressures are decreased by the shielding action of the lower wing. (2) the burble of the biplane lower wing, with respect to the angle of attack, is delayed, due to the shielding action of the lower wing. (3) the center of pressure of the biplane upper wing (semispan) is, in general, displaced forward and outward with reference to that of the wing as a monoplane, while for the lower wing there is but slight difference for both conditions. (4) the overhanging portion of the upper wing is little affected by the presence of the lower wing.

Loeser, Oscar E , Jr↗