Control-oriented models of hypersonic vehicle propulsion systems require a reduced-order model of the scramjet inlet that is accurate to within 10% but requires less than a few seconds of computational time. To achieve this goal, a reduced-order model is presented, which predicts the solution of a steady two-dimensional supersonic flow through an inlet or around any other two-dimensional geometry. The model assumes that the flow is supersonic everywhere except in boundary layers and the regions near blunted leading edges. Expansion fans are modeled as a sequence of discrete waves instead of a continuous pressure change. Of critical importance to the model is the ability to predict the results of multiple wave interactions rapidly. The rounded detached shock near a blunt leading edge is discretized and replaced with three linear shocks. Boundary layers are approximated by displacing the flow by an empirical estimate of the displacement thickness. A scramjet inlet is considered as an example application. The predicted results are compared to two-dimensional computational fluid dynamics solutions and experimental results. Nomenclature a = local sound speed, m=s c = specific heat, J=kg K H = length normal to flow, m h = specific enthalpy, J=kg L = length tangent to flow, m M = Mach number n = number of a given quantity Pr = Prandtl number p = pressure, Pa R = normalized gas constant, J=kg K R = 8314:47 J=kmol K r = radius, m T = temperature, K u = velocity magnitude, m=s W = molecular weight, kg=kmol x = forward body-frame coordinate, m Y = mass fraction z = vertical body-frame coordinate, m = shock angle = ratio of specific heats = thickness of layer, m " = ratio = ln p 0 =p = dynamic viscosity, kg=m s = flowpath angle = =2 = M 2 1 p = sin 1 1=M, Mach angle = Prandtl-Meyer function = density, kg=m 3 = wave angle = flux of subscripted quantity = reference angle Subscripts A, B, . . . = region label a, b, . . . = point label bs = curved portion of bow shock cl = property of inlet cowl e = value at edge of boundary layer ex = expansion i = species index j = index of expansion discretization k = region index le = leading edge p = constant pressure s = constant entropy sp = pertaining to species w = wall value 0 = stagnation value 1 = index for inlet portion of flow 2 = index for inlet outflow 1 = freestream Superscripts = value at Mach number of 1 = reference value for boundary layer

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This paper provides details of the combustion and inlet submodels used in the Michigan-Air Force Scramjet In Vehicle (MASIV) model. The model solves conservation equations in 1-D, using several modeling techniques to retain some of the fidelity of higher-order simulations. Inlet wave interactions, fuel mixing and finite-rate chemistry are considered. The order of the problem is reduced by physics-based, experimentally-verified algebraic scaling laws, which retains the required physics but reduces the computation time of the problem to seconds, instead of the several days required by computational fluid dynamics (CFD). Scaling coefficients and assumptions are given. The model is used to compute the performance of an experimental configuration for which real

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