Transpiration Cooling under Supersonic Rocket Engine Nozzle Flow Conditions

Abstract

In a preliminary feasibility study, transpiration cooling in the supersonic flow of a conical laval type nozzle is investigated. Hot gas flow with the flow conditions of a rocket engine nozzle is created via a detonation tube combusting hydrogen and oxygen at a stagnation pressure of 30 bar with a hot gas flow Mach number of Ma = 3.35 at the beginning of the coolant injection. In this context, a variation of the transpired coolant mass flow rate, using helium as a coolant gas, was conducted. Special emphasis is laid on the heat flux reduction on both the transpiration cooled segment as well as the solid nozzle structure in the wake flow. As result, cooling efficiencies of up to 0.85 at the end of the porous sample could be measured. For the creation of an cooling film, it could be shown that comparably high blowing ratios are needed for creating a lasting cooling film.