Design, Tests, and Verification of a Diffuser System for a Mach 4 Scramjet Test Facility

Abstract

R directed toward developing the technology for an airframe-integrated modular scramjet engine concept currently being conducted at NASA Langley Research Center, involves experimental investigations of hydrogen-burning scramjet models at simulated Mach 4 and 7 flight conditions. The present concept of the airframeintegrated scramjet engine has modules of rectangular cross section with swept leading edges that produce an asymmetric downward flow that sharply increases locally near the cowl when the scramjet inlet unstarts. Tests on a heat-sink, hydrogen-burning model representing one module of this concept have been conducted at Langley in a facility which duplicates Mach 7 flight conditions. For Mach 4 tests, an engine test cell is being modified to contain a freejet, blowdown tunnel that will exhaust to the atmosphere. To provide shock-free tunnel flow to the modular scramjet engine and to meet the atmospheric exhaust condition, a diffuser system is required. The available mass flow rates and the size of existing tunnel hardware dictated that the scramjet model block up to 33% of the tunnel nozzle exit area. A diffuser preliminary design was defined based on published experimental results. However, most of the literature results are not directly applicable to conditions where the model blockage is as high as 33%, and where the model design produces asymmetry by creating a downstream flow that increases when the model inlet unstarts. An experimental investigation was, therefore, undertaken using unheated air and a subscale model of the tunnel-scramjetdiffuser system. A description of the subscale system and some data results obtained during tests with this system are presented. These results were subsequently verified, as a data comparison will show, by full-scale engine combustion tests in a Mach 4 facility utilizing a diffuser system based upon the best arrangement of the subscale tunnel scramjetdiffuser system.