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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 217 records · Page 12
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NASA NTRS1960

Heat Transfer to 36.75 and 45 degree Swept Blunt Leading Edges in Free Flight at Mach Numbers from 1.70 to 2.99 and From 2.50 to 4.05

A flight investigation has been conducted to study the heat transfer to swept-wing leading edges. A rocket-powered model was used for the investigation and provided data for Mach number ranges of 1.78 to 2.99 and 2.50 to 4.05 with corresponding free-stream Reynolds number per foot ranges of 13.32 x 10(exp 6) to 19.90 x 10(exp 6) and 2.85 x 10(exp 6) to 4.55 x 10(exp 6). The leading edges employed were cylindrically blunted wedges ', three of which were swept 450 with leading-edge diameters of 1/4, 1/2, and 3/4 inch and one swept 36-750 with a leading-edge diameter of 1/2 inch. In the high Reynolds number range, measured values of heat transfer were found to be much higher than those predicted by laminar theory and at the larger values of leading-edge diameter were approaching the values predicted by turbulent theory. For the low Reynolds number range a comparison between measured and theoretical heat transfer showed that increasing the leading-edge diameter resulted in turbulent flow on the cylindrical portion of the leading edge.

ONeal, Robert L.↗
NASA NTRS1959

Effects of Nose Radius and Extreme Cooling on Boundary-Layer Transition for Three Smooth 15°-Cone-Cylinders in Free Flight at Mach Numbers to 8.50

Three highly polished 15°- included- angle cone- cylinders with hemispherical tips of several diameters ( 2, 3, and 4 in.) have been flown in order to obtain boundary- layer transition data at very low wall to local stream temperature ratios, and heat- transfer data. All surfaces had a 2-microinch average roughness height. Laminar flow existed over the entire hemispherical nose of the 2- and 3-inch-tip- diameter models throughout the complete flight history. Extreme cooling to wall to local stream temperature ratios at the sonic point as low as 0.20 did not cause transition on the nose for diameters as large as 3 inches. However, extreme cooling did cause early transition on the 4-inch model where it appears probable that transition occurred forward of the 45° station at a wall to local stream temperature ratio of about 0.26. Variations in tip diameter influenced transition downstream of the nose under conditions of extreme cooling. The 2-inch- tip model was laminar at all cone- cylinder stations at temperature ratios as low as 0.32 whereas the 3- and 4-inch-tip models were turbulent at the same local flow conditions but at higher wall to local temperature ratios. Transition on the cone and cylinder of the 3- and 4-inch- tip bodies appeared to be sensitive to local Mach number, and occurred at higher local temperature ratios when values of local Mach number were higher. Increasing the nose diameter from 2 to 3 inches significantly changed the local flow conditions for which laminar flow existed on the cone- cylinder afterbody. However, a further increase in tip size t o a 4-inch diameter had no discernable effect on the local flow conditions at transition. The transition results of the 3- and 4-inch-nose-diameter smooth bodies are similar to those observed on a 7/8-inch-nose-diameter body with roughened surfaces. Turbulent boundary layers resulted in both cases at very low wall to local stream temperature ratios. Both laminar and turbulent heat-transfer data were in good agreement with theoretical Stanton numbers when heat-transfer reduction due to tip blunting was considered.

CYLINDRICAL AFTERBODY↗