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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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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↗
NASA NTRS1979

Pressure distributions on three different cruciform aft-tail control surfaces of a wingless missile at mach 1.60, 2.36, and 3.70 Volume 1: Trapezoidal tail

The results of pressure distribution tests conducted in the Langley Unitary Plan wind tunnel are presented. The data were obtained for three sets of cruciform aft-tail control surfaces on a wingless missile model at Mach numbers of 1.60, 2.36, and 3.70 for angles of attack from -4 degrees to 20 degrees, model roll angles from 0 degrees to 90 degrees, and tail deflections of 0 degrees and 15 degrees. The test Reynolds number used was 6.6 million per meter.

Lamb, M.↗
NASA NTRS1979

Pressure distributions on three different cruciform aft-tail control surfaces of a wingless missile at Mach 1.60, 2.36, and 3.70. Volume 1: Trapezoidal tail

The results of pressure distribution tests conducted in a wind tunnel are presented without analysis. The data were obtained for trapezoidal aft tail control surfaces on a wingless missile model at Mach numbers of 1.60, 2.36, and 3.70 for angles of attack from -4 to 20 deg model roll angles from 0 to 90 deg and tail deflections of 0 and -15 deg. The test Reynolds number used was 6.6 million per meter.

Lamb, M.↗
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 20-inch Mach 6 air tunnel

Results from pressure distribution tests on 0.04-scale and 0.02-scale models of the forward fuselage of the Space Shuttle Orbier are presented without analysis. The tests were completed in the Langley 20-Inch Mach 6 Tunnel. The 0.04-scale model was tested at angles of attack from 0 to 35 and angles of sideslip from 0 to -4. The 0.02-scale model was tested at angles of attack from -10 to 45 and angles of sideslip from 0 to -4. 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 to 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 had provided a means for comparisons between reentry flight pressure data and wind-tunnel data.

Bradley, P. F.↗
NASA NTRS1937

Pressure-Distribution Measurements at Large Angles of Pitch on Fins of Different Span-Chord Ratio on a 1/40-Scale Model of the U. S. Airship "Akron."

Report presents the results of pressure-distribution measurements on a 1/40-scale model of the U. S. Airship "Akron" conducted in the NACA 20-foot wind tunnel. The measurements were made on the starboard fin of each of four sets of horizontal tail surfaces, all of approximately the same area but differing in span-chord ratio, for five angles of pitch varying from 11.6 degrees to 34 degrees, for four elevator angles, and at air speeds ranging from 56 to 77 miles per hour. Pressures were also measured at 13 stations along the rear half of the port side of the hull at one elevator setting for the same five angles of pitch and at an air speed of approximately 91 miles per hour. The normal force on the fin and the moment of forces about the fin root were determined. The results indicate that, ignoring the effect on drag, it would be advantageous from structural considerations to use a fin of lower span-chord ratio than that used on the "Akron."

Mchugh, James G↗
NASA NTRS1988

Aerodynamic performance and pressure distributions for a NASA SC(2)-0714 airfoil tested in the Langley 0.3-meter transonic cryogenic tunnel

This report presents in graphic and tabular forms the aerodynamic coefficient and surface pressure distribution data for a NASA SC(2)-0714 airfoil tested in the Langley 0.3-Meter Transonic Cryogenic Tunnel. The test was another in a series of tests involved in the joint NASA/U.S. Industry Advanced Technology Airfoil Tests program. This 14% thick supercritical airfoil was tested at Mach numbers from 0.6 to 0.76 and angles of attack from -2.0 to 6.0 degrees. The test Reynolds numbers were 4 million, 6 million, 10 million, 15 million, 30 million, 40 million, and 45 million.

Jenkins, Renaldo V.↗
NASA NTRS1991

Design of 3-dimensional complex airplane configurations with specified pressure distribution via optimization

A subcritical panel method applied to flow analysis and aerodynamic design of complex aircraft configurations is presented. The analysis method is based on linearized, compressible, subsonic flow equations and indirect Dirichlet boundary conditions. Quadratic dipol and linear source distribution on flat panels are applied. In the case of aerodynamic design, the geometry which minimizes differences between design and actual pressure distribution is found iteratively, using numerical optimization technique. Geometry modifications are modeled by surface transpiration concept. Constraints in respect to resulting geometry can be specified. A number of complex 3-dimensional design examples are presented. The software is adopted to personal computers, and as result an unexpected low cost of computations is obtained.

Kubrynski, Krzysztof↗