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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.

  • PDF System diagram
  • PDF Access map (Kakuda)
    • Department

    Research Unit IV，Research and Development Directorate
    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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