Exhaust nozzle recombination studies are carried out for hydrogen-fueled subsonic combustion ramjet engines. Bray's criterion is used to determine a nozzle freezing point. Nozzle performance is calculated by assuming equilibrium conditions up to the freezing point and frozen flow from the freezing point to the nozzle exit. Several engine parameters are varied in order to determine their effect on recombination and specific impulse. These parameters include nozzle throat size, throat radius of curvature, nozzle shape, Mach number, altitude, equivalence ratio, and inlet kinetic energy efficiency. The study has shown that recombination losses are not expected to be serious for Mach numbers up to 9 and freestream dynamic pressures above 500 psf. Nomenclature A= cross-sectional area A c = combustion chamber exit area A * = critical flow or throat area CD = overall drag coefficient of combustion chamber Cy = velocity coefficient DC = combustion-chamber drag 3D = dissociation rate F n et = net stream thrust g -gravitational constant H = static enthalpy /net = net fuel specific impulse J = energy conversion factor k r = forward reaction rate for reaction r k-r = reverse reaction rate for reaction r I = length M = Mach number M = mean molecular weight m = mass-flow rate, slugs/sec N = number of moles per unit mass of mixture n = number of moles tl = number of species in rth reaction equation P = static pressure 0 = net rate of molar change go = freestream dynamic pressure, p a Fo 2 /2 RC = nozzle throat radius of curvature #o = universal gas constant RT = nozzle throat cross-sectional radius (R = recombination rate= fuel injection angle j = specific-heat ratio TIKE -inlet-kinetic-energy efficiency for subsonic combustion vi rf = stoichiometric coefficient, left side of reaction equation Vi rff = stoichiometric coefficient, right side of reaction equation P = mass density $ = equivalence ratio Subscripts a = air C = combustion chamber e -nozzle exit / -fuel Presented at the AIAA-ASME Hypersonic Ramjet Conference, White Oak, Md., April 23-25, 1963; revision received July 22, 1963. * Aerospace Engineer, Mission Analysis Branch. t Aeronautical Research Scientist. i = species j = species in rth equation m = number of molecular species P = constant static pressure S = constant static entropy T = constant static temperature 0 = freestream 1 = inlet entrance 2 = combustion-chamber entrance 3 = combustion-chamber exit

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