An experimental investigation was conducted to determine the thrust coefficient of a high-area-ratio rocket nozzle at combustion chamber pressures of 12.4 to 16.5 MPa (1800 to 2400 psia). A nozzle with a modified Rao contour and an expansion area ratio of 1025:1 was tested with hydrogen and oxygen at altitude conditions. The same nozzle, truncated to an area ratio of 440:1, was also tested. Values of thrust coefficient are presented along with characteristic exhaust velocity efficiencies, nozzle wall temperatures, and overall thruster specific impulse.
Experimental Performance of a High-Area-Ratio Rocket Nozzle at High Combustion Chamber Pressure
1996
28 pages
Report
No indication
English
Rocket Engines & Motors , Divergent nozzles , Exhaust velocity , Nozzle efficiency , Nozzle geometry , Nozzle thrust coefficients , Rocket nozzles , Specific impulse , Thrust chamber pressure , Altitude tests , Atmospheric pressure , Liquid hydrogen , Liquid oxygen , Mass flow , Nozzle walls , Turbulent boundary layer , Wall temperature
Experimental Performance of a High-Area-Ratio Rocket Nozzle at High Combustion Chamber Pressure
British Library Conference Proceedings | 1995
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