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

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