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A lot of studies on rotating detonation engines have been carried out due to the higher thermal efficiency. However, the number, rotating directions, and intensities of rotating detonation waves are changeful when the flow rate, equivalence ratio, inflow conditions, and engine schemes vary. The present experimental results showed that the combustion mode of a rotating detonation engine was influenced by the combustor scheme. The annular detonation channel had an outer diameter of 100 mm and an inner diameter of 80 mm. Air and hydrogen were injected into the combustor from 60 cylindrical orifices in a diameter of 2 mm and a circular channel with a width of 2 mm, respectively. When the air mass flow rate was increased by keeping hydrogen flow rate constant, the combustion mode varied. Deflagration and diffusive combustion, multiple counterrotating detonation waves, longitudinal pulsed detonation, and a single rotating detonation wave occurred. Both longitudinal pulsed detonation and a single rotating detonation wave occurred at different times in the same operation. They could change between each other, and the evolution direction depended on the air flow rate. The operations with a single rotating detonation wave occurred at equivalence ratios lower than 0.60, which was helpful for the engine cooling and infrared stealth. The generation mechanism of longitudinal pulsed detonation is developed.

Controlling and redirecting the low-kinetic-energy layers of flow on the fuselage of aircrafts, is one of the most demanding stages in the advanced aerodynamic designs. Among the applications, Inlet/body integration needs the serious considerations to prevent the boundary layer from being swallowed by the inlet entrance. Not only the classic diverters are outclass nowadays, but also the newly used bump surfaces have still faced some problems under the effects of thick boundary layer. In this paper a new concept for highly integrated inlets has been presented. The concept called Ridge, is an aerodynamic surface, involves a vortex and a pressure gap to redirect the boundary layer. The performance of the ridge in redirecting the boundary layer and its flexibility has been proven by numerical simulations for different cases. According to numerical calculations, clean entrance from the boundary layer with a significant reduction in the total cross section area of aircraft has been resulted. The new surface has shown a great potential to integrate or combine with different aerodynamic geometries like external compression surfaces and it can cover a wide range of Mach numbers. This article has focused on the usage of new concept for inlet/body integration. NomenclatureC D = drag coefficient Δe = mechanical energy lose C Dt = flight vehicle forebody drag coefficient β = diversion angle I vis = viscose force increment percentage δ = ridge blade height I D = drag force increment percentage ψ = Yaw angle D p = pressure drag force ω = vorticity D v = viscous drag force v = velocity vector D = total drag force v m = vorticity magnitude h = normal distance from fuselage P t∞ = free stream total pressure HPS = high pressure surface of ridge blade BR = boundary layer removing ability M = Mach number LPS = low pressure surface of ridge blade P = static pressure P t = total pressure R CD = percentage reduction of total drag coefficient R Pt = percentage reduction of total pressure recovery coefficient σ F = viscous drag to pressure drag ratio σ = total pressure recovery coefficient SST = shear stress transport 1 Ph.D student, College of energy and power engineering, 29 Yudao St., Nanjing 210016, china, eiman@nuaa.edu.cn 2 Professor, College of energy and power engineering, 29 Yudao St., Nanjing 210016, china, hgp@nuaa.edu.cn, and AIAA Senior Member 3 Ph.D student, College of energy and power engineering, 29 Yudao St.,

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