Preliminary Design of a 2D Supersonic Inlet to Maximize Total Pressure Recovery

TLDR: In this paper, a two-dimensional, mixed compression, two-ramp supersonic inlet was designed to maximize total pressure recovery and match the mass flow demand of the engine.

Abstract: *† This paper provides a method of preliminary design for a two-dimensional, mixed compression, two-ramp supersonic inlet to maximize total pressure recovery and match the mass flow demand of the engine. For an on-design condition, the total pressure recovery is maximized according to the optimization criterion, and the dimensions of the inlet in terms of ratios to the engine face diameter are calculated. The optimization criterion is defined such that in a system of (n-1) oblique shocks and one normal shock in two dimensions, the maximum shock pressure recovery is obtained when the shocks are of equal strength. This paper also provides a method to estimate the total pressure recovery for an off-design condition for the specified inlet configuration. For an off-design condition, conservative estimation of the total pressure recovery is given so that performance of the engine at the off-design condition can be estimated. To match the mass flow demand of the engine, the second ramp angle is adjusted and the open/close schedule of a bypass door is determined. The effects of boundary layer are not considered for the supersonic part of the inlet, however friction and expansion losses are considered for the subsonic diffuser. Nomenclature α = Angle of attack j β = The installation angle of the j th ramp γ = The ratio of specific heats j δ = The flow deflection angle of the j th shock (j th ramp half angle) d θ = The half expansion angle of the subsonic diffuser j θ = The shock wave angle of the j th shock * A = The cross section area of flow tube at throat where the flow is sonic j A = The cross section area of flow at j th station point 54 AR = The ratio of inlet cross section areas at station points 5 and 4 5 d = The engine diameter at station point 5 (engine face) 6 d = The engine diameter at station point 6 (fan face) H = Flight altitude c h , 0 h = The captured freestream flow tube height i h = The height of inlet at the entry, measured perpendicular to the flight direction j h = The height of j