Development and Testing of a Prototype LOX/propylene Upper Stage Engine

The paper discusses the development of a 2000 N thrust liquid oxygen/propylene rocket engine designed to power the upper stage of a Nanosat Launch Vehicle (NLV). The preliminary design is intended for space operations with an expansion ratio of 70. The targeted combustion eciency is 95% and nozzle eciency is 98%, corresponding to a specic impulse of 347 s. Consistent with the employed incremental approach, these requirements are relaxed for the rst prototype version of the engine in order to conduct a static re test (SFT) demonstration at sea-level conditions: the expansion ratio of the nozzle is reduced to 4 by truncating the nozzle, the targeted combustion eciency is 90% and nozzle eciency is 95%. Propellants are introduced and mixed in the combustion chamber utilizing an unlike doublet injector element. In addition, lm cooling is provided in order to extend the life of the ablative chamber. Ignition is accomplished with a single igniter mounted on the center face of the injector. CFD analysis has been performed to validate the design and to characterize the engine’s performance. Results show that the upper stage engine produces a thrust of 1979 N with an exit Mach number of 4.3, compared to a one-dimensional calculated Mach number of 4.12. Analysis for the truncated prototype indicate a thrust of 1467 N when the predicted value in those conditions is 1270 N. The exit Mach number is determined to be 2.7.

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