Development of a Short-Duration Rocket Nozzle Flow Simulation Facility

摘要

The rising demand for reliable wind-tunnel flow data for rocket nozzle design applications led to the development of a new type of facility. Most of the former experimental data in the literature were gathered in conditions far from those encountered in real engines and combustion gas generators. Thus, they are not appropriate for the physical interpretation of complex flow phenomena and computational fluid dynamic tools validation. The new experimental platform is intended to solve this problem by its capability to nearly match most of these conditions. It allows for the investigation of several engine-relevant aerothermodynamic problems such as boundary-layer transition, nozzle flow separation, nozzle exhaust plume interactions, etc. Furthermore, innovative nozzle cooling techniques like film cooling can be studied in this new type of facility. This is demonstrated first by the experimental results given at the end of this paper. High nozzle reservoir pressures and temperatures are achieved by means of detonative combustion of a premixed gas in a confined tube. A detonation wave produces a high-energy flow that expands through a Laval nozzle to the desired experimental flow conditions. The facility simulates enginelike conditions in a combustion environment for various mass ratios of oxidizer-fuel mixtures. The short-duration flow device affords a simple and cost-effective research tool for testing under harsh flow conditions in laboratories. Hence, it will significantly contribute to the understanding of complex rocket nozzle flow phenomena and validation of numerical tools, as well as correlations.