An experimental investigation of a two-dimensional Scramjet inlet at flow Mach numbers of 8 to 25 and stagnation temperatures of 800 to 4,100 K

The development of air-breathing hypersonic vehicles, such as the National Aerospace Plane, NASP, and NASP Derived Vehicles, NDV, requires the detail understanding of the physics of the hypersonic internal and external flowfields. The internal flow through the Supersonic Combustion Ramjet, Scramjet, is particularly important for these vehicles since hypersonic velocities to Mach = 25 must be achieved by the use of this type of propulsion system. The limited available experimental data for the internal air flow in a Scramjet inlet at high Mach numbers and stagnation temperatures motivated this study. In this investigation, a 0.25 m span variable geometry 2-D Scramjet inlet was designed, constructed and tested in the Rensselaer Polytechnic Institute 0.61 m diameter Hypersonic Shock Tunnel. For these tests, the flow Mach number in the test section was varied from 8 to 25 and the reservoir temperatures investigated were in the 800-4,100 K range. The free stream until Reynolds number varied from 1.3 {dollar}times{dollar} 10{dollar}sp4{dollar} to 2.3 {dollar}times{dollar} 10{dollar}sp6{dollar} m{dollar}sp{lcub}-1{rcub}{dollar} and the leading edge free stream Knudsen number was in the 0.1-31 range. The lower end of the reservoir temperature, 800-1,100 K, was generated by operating the shock tunnel in the Reflected Mode. On the other hand, the Equilibrium Interface Mode of operation, with helium in the driver section, was required to produce the upper end of the stagnation temperature range of 4,100 K, with real gas effects.; The tests included surface and pitot pressure measurements along the Scramjet inlet centerline, and Schlieren and Air-Luminosity photographs were taken simultaneously with the pressure measurements to provide the internal flow visualization of the shock waves and the boundary layer. Both the pressure measurements and the flow visualization indicated a very complex internal flow structure in the duct passage. This complexity is caused by the hypersonic shock wave and the boundary layer interactions as well as the large free stream Knudsen numbers and the high reservoir temperatures.