Hydrogen-air supersonic combustion study by strut injectors

Abstract

The supersonic combustion experiments were conducted with seven integrated nozzle-strut injectors at the stagnation temperature around 2000 k, stagnation pressure 1-1.4 Mpa, global equivalence ratio 0.4-1.1 and a fixed entry Mach number 2.5. A one-dimensional supersonic combustion computer code SCC-1 based on HAP' code in terms of wall static pressure data was developed. The calculated results are fairly consistent with the experimental measurements. The factors affecting combustion efficiency and total pressure recovery were discussed. Introduction Many years of research have shown that the feasibility of hypersonic propulsion largely depends on the efficiency of supersonic combustion in the scramjet combustor. The crucial issues associated with the combustion efficiency are the fuel-air mixing and mixing enhancement due to the very limited available residence tune which is less than one millisecond. Variant approaches including parallel, angled and normal injection and mixing enhancement by use of hypermixers have been investigated extensively. In real scramjet combustor the flow is extremely complex which is obviously quite different from those employed in most fundamental mixing enhancement studies. For example, the shock wave in the inlet and isolator and thick turbulent boundary layer will inevitably propagate into the combustor from upstream. It will cause the interaction of shock-mixing layer and of shockboundary in combustor. The former may benefit the mixing rate enhancement, however the later may induce boundary separation to form low speed recirculating flow. Obviously it causes problem for most hypermixer-fuel injector mounted on the combustor wall in which the local flow should be supersonic without separation. The parallel injection has its advantage of fuel momentum addition , however its non-efficient mixing is well known. The defect of parallel mixing can be overcomed by normal injection in a great extent However the normal injection induces detached normal shock wave together with the boundary layer separation. The severe local Copyright © 1998 by the authors, published by the American Institute of Aeronautics and Astronautics, Inc, with permission loss of total pressure leads to a decrease in overall cycle efficiency. It seems very hard to solve the issue, however through the research of engine performance, D. M. Bushnell'' points out that in the Mach number 6 to 12 range scramjet engine produce adequate thrust margin and therefore transverse injector may be utilized providing appreciable mixing enhancement. We believe that the most straightforward way to solve this issue would be to reduce the mixing gap by placing struts within the interval air passage as well as relief the total pressure loss.. It would probably be a promising approach to meet the requirement of scramjet overall cycle efficiency. In the consideration of designing strut injector the problem, which is material and cooling how to prevent the head of strut from being burned out within the high temperature flow, was encountered. The alternative approach to circumvent this difficulty is to utilize a kind of integrated nozzle strut injector structure which is a combination of a supersonic nozzle and fuel injector. A great number experiments show that it is good for supersonic combustion study with in stream struts. However, the disadvantage is also found that the isolator could not be well simulated. Besides, from the view of experiment due to the extremely hostile condition in experimental supersonic combustion burner, the measurement of static pressure along the burner wall is relatively straightforward, accurate and reliable compared to total pressure and static temperature. Usually the one-dimension approximation to the flow can provide a reasonable description of the flow behavior. The one-dimension assumption means the flow is uniform at any cross-section. Although the one-dimension approach can never be perfect correct, it can help to reveal and understand