Download Free Preliminary Results From A Free Flight Investigation Of Boundary Layer Transition And Heat Transfer On A Highly Polished 8 Inch Diameter Hemisphere Cylinder At Mach Numbers Up To 3 And Reynolds Numbers Based On A Length Of 1 Foot Up To 177 X 106 Book in PDF and EPUB Free Download. You can read online Preliminary Results From A Free Flight Investigation Of Boundary Layer Transition And Heat Transfer On A Highly Polished 8 Inch Diameter Hemisphere Cylinder At Mach Numbers Up To 3 And Reynolds Numbers Based On A Length Of 1 Foot Up To 177 X 106 and write the review.

Heat-transfer rates have been measured in free flight along the stagnation line of an unswept cylinder mounted transversely on an axial cylinder so that the shock wave from the hemispherical nose of the axial cylinder intersected the bow shock of the unswept transverse cylinder. Data were obtained at Mach numbers from 2.53 to 5.50 and at Reynolds numbers based on the transverse cylinder diameter from 1.00 x 106 to 1.87 x 106. Shadowgraph pictures made in a wind tunnel showed that the flow field was influenced by boundary-layer separation on the axial cylinder and by end effects on the transverse cylinder as well as by the intersecting shocks. Under these conditions, the measured heat-transfer rates had inconsistent variations both in magnitude and distribution which precluded separating the effects of these disturbances. The general magnitude of the measured heating rates at Mach numbers up to 3 was from 0.1 to 0.5 of the theoretical laminar heating rates along the stagnation line for an infinite unswept cylinder in undisturbed flow. At Mach numbers above 4 the measured heating rates were from 1.5 to 2 times the theoretical rates.
A highly polished hemisphere-cone having a ratio of nose radius to base radius of 0.74 and a half-angle of 14.5 was flight tested at Mach numbers up to 4.70. Temperature and pressure data were obtained at Mach numbers up to 3.14 and a free-stream Reynolds number of 24 x 106 based on body diameter. The nose of the model had a surface roughness of 2 to 5 microinches as measured with an interferometer. The measured Stanton numbers were in good agreement with theory. Transition Reynolds numbers based on the laminar boundary-layer momentum thickness at transition ranged from 2,190 to 794. Comparison with results from previous tests of blunt shapes having a surface roughness of 20 to 40 microinches showed that the high degree of polish was instrumental in delaying the transition from laminar to turbulent flow.
Abstract: Equilibrium temperatures and heat-transfer coefficients for a hemispherical nose have been measured for Mach numbers from 1.62 to 3.04. Heat transfer to the surface of the hemisphere was presented as Stanton number against Reynolds number for various surface heating conditions. Heat transfer at the stagnation point has been measured and correlated with theory. Transition from a laminar to a turbulent boundary layer was obtained at Reynolds numbers of approximately 1 x 106 corresponding to a region on the body located between 45© and 60© from the stagnation point.
Heat-transfer data from four wind-tunnel experiments and two free-flight experiments with turbulent boundary layers have been examined to see whether or not they are well represented by the Reynolds analogy or a modification thereof. The heat-transfer results are put into the form of dimensionless Stanton numbers based on fluid properties at the outer edge of the boundary layer and are compared with skin-friction coefficients for the same Mach numbers and wall to free-stream temperature ratios as obtained from an interpolation of the existing skin-friction data. The effective Reynolds number is taken to be the length Reynolds number measured from the effective turbulent origin, a position which differs importantly from the leading edge of the test surface in some cases.