1. The forerunner of the 8-Foot High Temperature Tunnel was the 8-Foot High Temperature Structures Tunnel. This facility was designed in the 1950's and was placed in service during the mid-1960's. For the following 20-plus years, the methane/air-heated facility was used to conduct research in the areas of aerothermal loads and aeromermostructures at simulated Mach 7 flight conditions. During the late 1980's and early 1990's, the tunnel was modified with the addition of an oxygen replenishment system and newMach 4 and 5 nozzles (refs. 12-13). These modifications allowed the use of the facility as a Mach 4, 5, and 7 propulsion test facility and the facility became operational in that capacity in 1993. The facility capability for aerothermal loads and aerothermostructural research remains intact and is, in fact, enhanced by the modifications. However, in this paper, the 8-Foot HTT will only be discussed in its function as a propulsion test facility.
2. The purpose of the 8-Foot HTT is to test complete, larger scale and multiple-module, scramjet component integration models in flows with stagnation enthalpies duplicating that in flight at Mach 4, 5, and 7. The flow at the exit of the facility nozzle simulates the flow upstream of the aircraft bow shock in flight. The stagnation enthalpy necessary to simulate flight Mach number for the engine tests is achieved through methane-air combustion with oxygen replenishment to obtain a test gas with the same oxygen mole fraction as atmospheric air (0.2095).
3. Facility Configuration: A schematic of the 8-Foot High Temperature Tunnel is shown in Figure 18. Air is supplied from a 6000 psia bottle field, methane from bottles at 6000psia, and liquid oxygen (LOX) from an 8000 gallon run tank at 2290 psia. The 8-Foot HTT combustion heater (Figure 19) consists of a laminated high-strength carbon steel pressure vessel, a stainless steel outer liner, and a Nickel 201 inner liner. High pressure air from the bottle field enters thepressure vessel through a torus at the upstream end, flows downstream between the pressure vessel and the outer liner, turns 180°, and flows upstream in the annular space between the inner and outer liners, thereby cooling the inner liner which is exposed to the hot combustion products and thermally protecting the carbon steel pressure vessel. The inner liner ends at approximately the mid-point of the combustor and the outer surface of the LOX injector ring forms a new annular channel with the outer liner which accepts the air flow from the annular space between the inner and outer liners. The LOX is injected at the beginning of the 20-inchlong annular space between the LOX ring and the outer liner where it mixes with the air. At the exit of the LOX injector ring, the oxygen-enriched air flows into the area bounded by the outer liner, turns 180°, and flows downstream to the methane fuel injection region. The methane is injected from 700 fuel injection orifices located on the downstream faces of 15 concentric rings of manifold tubing. The methane is mixed with and burns in the oxygen-enriched ah- in the half-length of the combustor upstream of the facility nozzle to provide stagnation pressures to 2000psia and stagnation temperatures to 3560 R.
4. The combustion products (with an oxygen mole fraction of 0.2095) exiting the combustor are expanded through an air-transpiration-cooled nozzle throat section (Fig. 19), which has a throat diameter of 5.62 inches. The nozzle geometry downstream of this throat section depends upon the desired test Mach number. If Mach 4 or 5 tests are required, a mixer section and the appropriate throat and nozzle section downstream of the mixer are substituted fora portion of the original Mach 7 nozzle (Fig. 20). In the mixer section, ambient-temperature air is added to the combustion-heated test gas to achieve the stagnation enthalpy for either Mach 4 or Mach 5 flight-simulation propulsion tests.