Combustion-chamber performance of a Python turbine-propeller engine with four tail-pipe configurations was determined in the NACA Lewis alti¬tude wind tunnel over a range of simulated altitudes from 10,000 to 40,000 feet and engine speeds from 6800 to 8000 rpm. Fuel flow was varied at each engine speed to give full coverage of the operable engine range.nOver the range of test conditions investigated, the combustion efficiency varied from approximately 0.95 to 0.99. Combustion efficiency decreased slightly with increased altitude and increased fuel-air ratio but was not affected noticeably by changes in engine speed and exhaust-nozzle-outlet area. The combustion-chamber total-pressure-loss ratio varied from approximately 0.037 in the medium and high range of corrected shaft horsepowers to about 0.043 at low corrected shaft horsepower. The value of combustion-chamber total-pressure-loss ratio increased slightly with increased altitude at low corrected shaft horsepower, but was not affected noticeably by changes in engine speed and exhaust-nozzle-outlet area. The combustion-chamber total-pressure loss divided by the inlet impact pressure increased uniformly with increasing combustion-chamber temperature ratio from a minimum value of 1.7 to a maximum of 2.4. At a given value of combustion-chamber temperature ratio, changes in altitude, engine speed, and exhaust-nozzle-outlet area had little if any effect on' the combustion-chamber total-pressure loss divided by the inlet impact pressure.
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