Accession Number:

AD0367753

Title:

INVESTIGATION OF LARGE THRUST PER ELEMENT INJECTORS WITH CONVENTIONAL AND TWO-DIMENSIONAL THRUST CHAMBERS UTILIZING LIQUID OXYGEN AND LIQUID HYDROGEN PROPELLANTS.

Descriptive Note:

Final rept. Feb 62-Feb 65 on Phase 1,

Corporate Author:

AIR FORCE ROCKET PROPULSION LAB EDWARDS AFB CA

Report Date:

1965-09-01

Pagination or Media Count:

79.0

Abstract:

Results of the tests and evaluations performed on highly simplified, large thrust per element injectors, utilizing cryogenic propellants, are presented. Phase 1 of the Scorpio Project was conducted to investigate large thrust per element injectors, in an attempt to alleviate inherent problems of conventional multi-orifice, liquid rocket injectors. The problems of complexity, costly and time-consuming fabrication, and stringent tolerance requirements, become more acute as higher thrust injectors are required to perform future rocket missions. LTE injector feasibility, high combustion efficiency, combustion stability, heat transfer characteristics, and cost were the major items of concern. Ninety-five test firings were conducted on twelve 50,000-pound large thrust per element injectors, utilizing liquid oxygen and liquid hydrogen propellants. Two uncooled thrust chamber configurations consisting of a square chamber with a two-dimensional nozzle and cylindrical chamber with a conical nozzle were utilized with various Ls 30 to 75 inches. Chamber pressure and mixture ratio were varied from 650 to 800 psig and 3 to 7, respectively. Injector design parameters were varied to determine their effect on performance. Combustion efficiency about 95 C sub th was achieved with a single-element injector yielding 50,000 pounds of thrust, thereby demonstrating the feasibility of large thrust per element injectors utilizing LO2LH2. Rapid and smooth ignition was achieved, as was the case with the simulated liquid airLH2 studies described in Report RTD-TDR-63-1041. No high-frequency combustion instability was experienced. A significant increase in combustion efficiency was obtained as the L was increased. Author

Subject Categories:

  • Liquid Propellant Rocket Engines

Distribution Statement:

APPROVED FOR PUBLIC RELEASE