Effects of Internal-Area Distribution, Spike Translation, and Throat Boundary-Layer Control on Performance of A Double-Cone Axisymmetric Inlet at Mach Numbers From 3.0 to 2.0

An investigation was made to determine the effects of initial subsonic diffusion rate, variable geometry, and throat boundary-layer control on the performance of an axisymmetric double-cone inlet. Results were obtained for Mach numbers from 3.0 to 2.0, angles of attack to 12°, and a Reynolds number of 3.07 X 106 based on the capture diameter.

Major gains in performance at the higher Mach numbers were obtained with boundary-layer control in the form of a ram scoop at the throat. Compared with the no-bleed configuration at Mach 3.01, the total-pressure recovery was increased from 0.625 to 0.74, with a concomitant increase in the cowl-plus-spillage drag coefficient from 0.11 to 0.14 due to a corresponding decrease in the mass-flow ratio from 0.96 to 0.91. This resulted in an improvement of 22 percent in propulsive thrust for a hypothetical turbojet engine with such a bleed system. Below Mach 2.1 where shock detachment occurred, boundary-layer bleed was not effective.

Minor improvements in internal performance were obtained at the higher Mach numbers (above 2.44) with a more gradual area expansion in the initial portion of the subsonic diffuser. However, as the spike was retracted, internal contraction was also encountered sooner. At reduced Mach numbers, fixed-geometry configurations incurred excessively high-spillage drags and somewhat lower pressure recoveries when compared with those of a translating-spike inlet.