RPA – Tool for Rocket Propulsion Analysis
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Rocket Propulsion Analysis
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Solid & hybrid propulsion
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Scripting Examples

/***************************************************
 Scripting example  
 ***************************************************/
prop = Propellant();
prop.setRatio(6.0, "O/F");    // Set O/F weight ratio
prop.addOxidizer("O2(L)");  // Add oxidizer
prop.addFuel("H2(L)");        // Add fuel

chamber = Chamber(prop);
chamber.setP(10, "MPa");
chamber.solve(true);
nozzleExit = NozzleSectionConditions(chamber, 40, "A/At")
...
...

See Scripting Examples for more information about RPA scripting.

Performance Analysis Examples

See page Solid and hybrid propulsion to find examples for performance analysis of solid and hybrid propulsion.

RD-170

Design parameters [1]
Parameter Value Unit
Oxidizer Liquid Oxygen (LOX) -
Fuel Kerosine RG-1 -
Components mass ratio 2.63 O/F
Combustion chamber pressure 24.5 MPa
Nozzle inlet contraction area ratio 2.6 Ac/At
Nozzle exit area ratio 36.87 Ae/At
Frozen equilibrium flow at 1.3* Afr/At

Comparison of actual and calculated engine parameters
Parameter Actual  RPA   Percent error
Specific impulse (vac), s 337 338 0.3 %
Specific impulse (SL), s 309 309 0.0 %

Download RD-170 configuration file

RD-253

Design parameters [2]
Parameter Value Unit
Oxidizer N2O4(L) -
Fuel UDMH -
Components mass ratio 2.67 O/F
Combustion chamber pressure 15.7 MPa
Nozzle inlet contraction area ratio 2.36 Ac/At
Nozzle exit area ratio 26.2 Ae/At
Frozen equilibrium flow at 6.0* Afr/At

Comparison of actual and calculated engine parameters
Parameter Actual  RPA   Percent error
Specific impulse (vac), m/s 3160 3156 0.13 %
Specific impulse (SL), m/s 2890 2832 2.00 %

Download RD-253 configuration file

RL10A3-3A

Design parameters [3],[4]
Parameter Value Unit
Oxidizer Liquid Oxygen (LOX) -
Fuel Liquid Hydrogen (LH2) -
Components mass ratio 5.5 O/F
Combustion chamber pressure 475 psia
Nozzle inlet contraction area ratio 4.6 Ac/At
Nozzle exit area ratio 61.0 Ae/At
Frozen equilibrium flow at 3* Afr/At

Comparison of actual and calculated engine parameters
Parameter Actual  RPA   Percent error
Specific impulse (vac), s 444.4 441.3 0.7 %

Download RL10A3-3A configuration file

SSME

Design parameters [5],[6]
Parameter Value Unit
Oxidizer Liquid Oxygen (LOX) -
Fuel Liquid Hydrogen (LH2) -
Components mass ratio 6.0 O/F
Combustion chamber pressure 3280 psia
Nozzle inlet contraction area ratio 3.4 Ac/At
Nozzle exit area ratio 77.5 Ae/At
Frozen equilibrium flow at 3.0* Afr/At

Comparison of actual and calculated engine parameters
Parameter Actual  RPA   Percent error
Specific impulse (vac), s 454.4 451.95 0.53 %

Download SSME configuration file

Conclusion

An excellent agreement is obtained between the actual performance and performance predicted by PRA program. The maximum relative difference is 2.0% occurring for RD-253.

Notes

*) Estimated values; (Afr/At)LOX/RP-1 < (Afr/At)LOX/LH2 < (Afr/At)N2O4/UDMH

References

  1. RD-170 rocket engine
  2. RD-253 rocket engine
  3. RL10A3-3A rocket engine specification
  4. SSME rocket engine
Copyright © 2009-2011 Alexander Ponomarenko  |  Contact  | 
Last modified: May 29, 2011