On the Reflection of Weak Shock Waves from Nozzles with No-Flow and Critical Flow through the Nozzle

Abstract

High-frequency combustion instability in rocket engines is known to result from pressure waves traveling within the combustion chamber. For the longitudinal mode of instability, the pressure wave reflects directly from the rocket nozzle and this mode behaves acoustically as though the engine were an organ pipe closed at both ends. Very little is known about the reflective characteristics of the nozzle in terms of its geometry. The reflection of shock waves from nozzles having contraction ratios of 2, 4, and 16 and convergence half-angles of 15°, 30°, and 60° was studied in a shock tube. Two sets of experiments were conducted. In one, a series of no-flow tests were made and the results compared to a theoretical analysis (Otto Laporte, Los Alamos Scientific Laboratory of the University of California; Report LA 1740, August, 1954). In the other, a critical flow of air through the nozzle was supplied. This condition simulates more closely the condition of the rocket nozzle throat during combustion. For no-flow tests, the reflected pressure amplitude compared favorably with that predicted by the analysis. Nozzle convergence angle had no observable effect. Under critical flow conditions, reflected pressure amplitude was essentially that expected for a closed tube and was independent of nozzle contraction ratio and convergence angle.