Wind tunnel for a combustion test using a supersonic and hypersonic air-breathing engine model
This wind tunnel for a combustion test uses a supersonic and hypersonic air-breathing engine model, including for ramjets, scramjets and rocket based combined-cycle engines. It is also possible to conduct engine model tests under static atmospheric conditions.
| Department | Space Technology Directorate I |
|---|---|
| Installation location | Kakuda Space Center |
| Measurement items | Pressure: 256 channels (electronic scanning type) 66 channels (fixed, including facility data) Temperature: 57 channels (including facility data) Flow rate: 5 channels (sample supply propellant) Heat flux: 16 channels Optical observations: 500 mm Schlieren |
| Fuel supply | Hydrogen gas Oxygen gas Ethanol |
| Test environment | Facility nozzle exit diameter: 51 cm x 51 cm (free jet) ◆Mach 8 flight conditions: Combination of thermal storage heating and combustion heating Nozzle exit Mach numbers: 6.7 Static temperature/static pressure: 237 K/575 Pa Maximum test time: 30 seconds ◆Mach 6 flight conditions: Thermal storage heating or combustion heating Nozzle exit Mach numbers: 5.3 Static temperature/static pressure: 222 K/2.6 kPa Maximum test time: 60 seconds ◆Mach 4 flight conditions: Thermal storage heating Nozzle exit Mach numbers: 3.4 Static temperature/static pressure: 217 K/5.5 kPa Maximum test time: 60 seconds *Given the assumption for air flow after the shockwave passage from the front edge of the vehicle, the Mach number flight conditions and facility provision Mach number differ. |
| Year of completion | FY 1993 |
