Accession Number : AD0372874

Title :   OXYGEN/HYDROGEN INJECTOR THRUST PER ELEMENT SIZE MAXIMIZATION.

Descriptive Note : Final rept. Apr 63-Apr 65 on phase 2,

Corporate Author : AIR FORCE ROCKET PROPULSION LAB EDWARDS AFB CALIF

Personal Author(s) : Main ,Howard V. ; George ,Daweel ; Mahugh ,Vernon L. ; Tepe,Lester E.

Report Date : DEC 1965

Pagination or Media Count : 57

Abstract : Results of the tests and evaluations performed on highly simplified, large thrust per element (LTE) injectors utilizing cryogenic propellants are presented. Phase II of Project Scorpio was conducted to determine the maximum thrust per injector element which could be utilized and still achieve high performance. High combustion efficiency, combustion stability, and heat transfer characteristics were the major items of concern. Twenty-nine test firings were conducted. The three injector patterns investigated were the: (1) single-element concentric pentad, (2) four-element concentric pentad, and (3) four-element concentric triplet. Two thrust chamber configurations, one uncooled and the other film cooled, were utilized. The basic characteristic length of 45 inches was varied to 75 inches by installing a section between the injector and thrust chamber. After a series of tests on each injector, modifications were made to the first two and the third one was fired reversed, i.e., fuel injected through oxidizer manifold and oxidizer injected through fuel manifold, to determine the effect on performance. Also, the injector film coolant holes and chamber film coolant passages were sealed to facilitate injector performance evaluation. Chamber stagnation pressure and mixture ratio were varied from 556 to 753 psia and 3.6/ 1 to 5.7/1, respectively. Combustion efficiency above 90% characteristic velocity was achieved with the four-element concentric pentad injector over a range of mixture ratios and chamber pressures.

Descriptors :   (*COMBUSTION CHAMBERS, FUEL INJECTORS), (*FUEL INJECTORS, OPTIMIZATION), LIQUID PROPELLANT ROCKET ENGINES, THRUST, COMBUSTION, STABILITY, HEAT TRANSFER, FILM COOLING, IGNITION, TRANSIENTS, CHECKOUT PROCEDURES, CAPTIVE TESTS, HYDROGEN, OXYGEN, PRESSURE, INJECTORS, LIQUEFIED GASES, CRYOGENIC PROPELLANTS (U) ROCKET NOZZLES, EROSION

Subject Categories : Liquid Propellant Rocket Engines

Distribution Statement : APPROVED FOR PUBLIC RELEASE