1. Reference 18 surveyed Wind Tunnels over the subsonic to hypersonic speed range;In
2. For flight velocities up to about 20,000 ftls (or Mach 20), the flight path of a hypersonic airbreathing vehicle can be typified by a freestream dynamic pressure in the range from 1000 to 2000 psf, (see Figure 1). Along such flight paths real gas effects behind a normal shock wave begin at about 9,000 ftls velocity where 10% of the oxygen is dissociated. Initially, this phenomenon is confined to the flow near a blunt nose cowl or leading edge. Because the vehicle is slender and at a low angle
3. v of attack, much of the flow continues to behave as an ideal gas until higher flight velocities are attained. This situation is favorable to the use of the conventional hypersonic wind tunnels which can produce flow Mach numbers up to 18 but at velocities not much greater than 7000 ftls. Essentially, these are ideal gas wind tunnels, a fact that is even more evident in a number of them where nitrogen or even helium is used as the test gar. These facilities can be useful in studying flow phenomena that can be characterized by just the Mach number, Reynolds number, specific heat ratio, `6, and the wall-to-inviscid flow temperature ratio. Two of the larger facilities in this category are the NSWC Tunnel #9 (>foot dia.) using nitrogen and the NASA ARC 3.5 It. Hypersonic Wind Tunnel. They can produce relatively high Reynolds numbers in the Mach number range from 5 to 14. Another important large hypersonic tunnel is the NASA LaRC a f t High Temperature Tunnel. Currently, this tunnel operates in the Mach number range 5.8-7.2 and the test medium is the product of combustion of methane and air. It has been used primarily for aero heating research. These hypersonic wind tunnels and others listed in References 1 and 18 can be useful for aerodynamic and aero heating studies of the NASP and similar vehicles provided that the Reynolds number is sufficiently close to the flight case. 3.2 Propulsion Test Facilities
4. There are four propulsion test facilities listed with the hypersonic wind tunnels of Reference 18. These are the NASA LaRC Scramjet Test Facility (M = 4.7 to 6.0), the GASL High Temperature Storage Heater Propulsion Wind Tunnel (M I 0.1 to 121, the GASL Vitiated-Air Heater (VAH) Propulsion Wind Tunnel (M = 2.7 to 8.0) and the GASL High Mass Flow Storage Heater Propulsion Wind Tunnel (HPB) (M = 0.1 to 7.0). They have maximum total temperature capabilities of 40000R,2000°R, 4500°R and 1700°R, respectively. These temperatures are still too low to simulate scramjet combustion processes at flight velocities above 10,000 ftls,