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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 361 records · Page 20
All sourcesNASA NTRSDOE OSTI
NASA NTRS1972

Effects of roughness on heating and boundary-layer transition. Part 2: effects of discrete roughness on heating

A preliminary investigation on the effects of roughness on heating in delta wing orbiters was conducted on the windward surface of the vehicle. The purpose of the experiment was to determine the feasibility of using parametric relationships obtained from discrete roughness effects on heating in flat plates and cones, to complex shuttle configurations. An attempt was also made to characterize what parameters were needed to influence roughness. Experimental results show that effective roughness Reynolds numbers are less than or equal to those of flat plates at the same edge Mach number, but, the roughness correlations for plates and cones cannot be applied to the shuttle configuration.

Mateer, G. G.↗
NASA NTRS1972

Effects of wall cooling and angle of attack on boundary layer transition on sharp cones at free stream Mach 7.4

Tests were conducted on 5 deg and 15 deg half-angle sharp cones at wall-to-total-temperature ratios of 0.08 to 0.4, and angles of attack from 0 deg to 20 deg. The results indicate that (1) transition Reynolds numbers decrease with decreasing temperature ratio, (2) local transition Reynolds numbers decrease from the windward to the leeward side of the model, and (3) transition data on the windward ray of cones can be correlated in terms of the crossflow velocity gradient, momentum thickness Reynolds number, local Mach number, and cone half-angle.

Mateer, G. G.↗
NASA NTRS1978

Influence of free-stream disturbances on boundary-layer transition

Considerable experimental evidence exists which shows that free stream disturbances (the ratio of root-mean-square pressure fluctuations to mean values) in conventional wind tunnels increase with increasing Mach number at low supersonic to moderate hypersonic speeds. In addition to local conditions, the free stream disturbance level influences transition behavior on simple test models. Based on this observation, existing noise transition data obtained in the same test facility were correlated for a large number of reference sharp cones and flat plates and are shown to collapse along a single curve. This result is a significant improvement over previous attempts to correlate noise transition data.

Harvey, W. D.↗
NASA NTRS1980

Aerodynamic heating for gaps in laminar and transitional boundary layers

The presence of gaps, slots, and/or steps in a surface may significantly perturb a supersonic boundary layer. The paper discusses heat-transfer distributions for a variety of gap configurations which were obtained by placing instrumented inserts in a flat-plate structural carrier. The data were obtained in a 3.50 foot hypersonic wind tunnel at a freestream Mach number of 5.10 over a range of Reynolds number from 2,570,000 to 8,110,000. The variables of the test program included the freestream Reynolds number and the gap configuration, e.g., width, depth, step height, number, and orientation.

Bertin, J. J.↗
NASA NTRS1985

On boundary layer transition in high-subsonic and transonic flow under the influence of acoustic disturbances and free-stream turbulence

The existing data base which quantifies the roles played by the Mach number, turbulence level, and acoustic noise in the flow quality of wind tunnels is reviewed. Attention is given to the effects on end (ET) and beginning (BT) transitions in flows over a 10 deg cone in transonic channels. The Re of both ET and BT flows were significantly affected by the Mach number, acoustic noise level and the noise frequency content. A unit increment in the Mach number increased the Re beyond 3,000,000, while noise below the 1 percent level in the rms pressure coefficient did not alter the transition phenomenon. More data are required on the effects of turbulence, which was a significant factor, and in the settling chamber, test section, plenum and diffuser to trace the locations and magnitudes of the acoustic sources.

Murthy, S. V.↗
NASA NTRS1987

Remote noncontacting measurements of heat transfer coefficients for detection of boundary layer transition in wind tunnel tests

An infrared measurement system is used that consists of a laser heating source, an infrared camera for data acquisition, and a video recorder for data storage. A laser beam is scanned over an airfoil, heating its surface to a few degrees above ambient. An infrared camera then measures the temperature of the airfoil over a two-dimensional field, and these temperatures are stored as a function of time on a video recorder. The resulting temperature pictures are digitized and an iterative approximation algorithm is used to extract the heat transfer coefficient. The resulting values are normalized to the natural convection condition. The technique has been applied in low-speed wind tunnel tests and compared to well-established hot-film measurements which were made simultaneously to confirm the flow conditions. Heat transfer coefficients were determined using a linear scanning pattern, to indicate the position of natural and of artificially induced transition on an airfoil, at various wind speeds. The technique is shown to be sensitive to transition at low Mach numbers. The advantages of the technique are discussed.

Heath, D. Michele↗
NASA NTRS1987

Status of a specialized boundary layer transition detection system for use in the U.S. National Transonic Facility

An improved deposition technique for cryogenic hot films used for transition detection in cryogenic tunnels is reported. Tests of the hot films in a low-speed tunnel demonstrated the ability to obtain online transition data. The capability of an enhanced hot film data acquisition system was also demonstrated. A comparison of data from the new system with stability theory shows the detection of Tollmein-Schlicting waves at transition onset.

Johnson, Charles B.↗
NASA NTRS1987

Supersonic boundary-layer transition on the LaRC F-106 and the DFRF F-15 aircraft. Part 2: Aerodynamic predictions

In the leading edge region, the measured pressure distributions exhibit extreme variations from strong suction peaks to a pressure maximum at the attachment line. These variations occur over short distances on the wing surface, and their character changes with changes in Mach number and angle of attack. The data/theory comparisons show that the character of the measured pressure distributions is well predicted for every Mach number and/or angle of attack condition considered. There is good agreement between theory and experiment for the location of the attachment line and suction peaks. The pressure magnitudes are well represented in the critical leading edge region, including the pressure maximum on the attachment line. The wing/body/inlet results agree well with the wing alone back to about 20 percent of chord where the upper surface suction peak typically occurs. The largest differences between theory and measurement always occur in the vicinity of suction peaks, with the difference being approximately 15 percent or less. In regions of largest error, the predicted pressures underestimate the suction peak strength for each case considered. The ability of the NCOREL code to reproduce wing pressure characteristics is shown.

Rose, Ollie J.↗