Experimental data are presented showing heat-flux distributions measured calorimetrically with several liquid-propellant rocket thrust-chamber configurations. Thrust levels of the experimental chambers were from 300 to 5000 lb. Enzian-type and axial-stream showerhead propellant injectors were utilized with hydrazine (N2H4) and nitrogen tetroxide (N2O4) propellants. Nozzle-contraction-area ratios of 8 to 1, 4 to 1, and 1.64 to 1 were tested, each having a 5-in. inlet diameter. Characteristic chamber lengths ranged from 16.95 to 62.8 in. The comparison between the experimental heat flux and the analytical heat flux using the method of Bartz [1] was found to be closest in the nozzle-expansion region. The experimental heat-flux measurements ranged between 80 per cent above and 45 per cent below the analytical estimates at the nozzle throat, however. These differences were dependent upon thrust-chamber configuration, injector type, and chamber pressure, and apparently resulted from nonideal combustion and flow characteristics. It is concluded that a priori determination of heat-flux distribution along the thrust-chamber length was possible only to a first approximation for the conditions of these tests.
Early detection of thermoacoustic instabilities in a cryogenic rocket thrust chamber using combustion noise features and machine learning
