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Pitot pressure and heat-transfer measurements have been made in plumes of 0.5-N (conical nozzle), 2-N, and 5-N (contoured nozzles) monopropellant hydrazine thrusters. The main objectives are to check the DFVLR simple plume model and to determine reliable model input values for real thrusters. The methods used for the heattransfer measurements (applying a sphere probe), the recovery temperature determination, and the evaluation of the plume quantities relevant for plume impingement calculations are outlined. The Pitot pressure measurements showed the existence of shock disturbances in the near plume flowfield of the contoured nozzles. Stagnation temperatures between 900 and 1350 K were deduced from the measured recovery temperatures. The corresponding molecular weight range was found to be between 11 and 14.5 and the most reasonable mean effective ratio of specific heats to be K = 1.4 ±0.03 for the expansion from the stagnation chamber to the continuum plume flow. This value is proposed for simple plume model calculations. The model heat-transfer results agree well with the experiments. Nomenclaturespecific heat at constant pressure and volume, respectively d = diameter F = thrust 7 sp = specific impulse 7 sp = F/m m -mass M = molecular weight Ma = Mach number p = pressure Pr = Prandtl number <2 = heat transfer r = radius r = recovery factor, Eq. (7) R = specific gas constant Re 2 = Reynolds number, Eq. (19) St = Stanton number, Eq. (6) t = time T = temperature u = velocity x = centerline distance from nozzle exit X l = dissociation degree of NH 3 d" = momentum thickness 6 E = nozzle exit angle AC = ratio of specific heats, K = C P /C V JLI = viscosity p = density Subscripts BL = boundary-layer (continuum) theory cond = conduction E = nozzle exit condition FM = free molecule K = in the vacuum chamber lim = limiting condition for Mo-oo loss = losses by radiation and conduction off, on = thruster not firing and firing, respectively r = at recovery condition rad = radiation s = sphere u = velocity w =at the wall, sphere probe condition 0 = stagnation condition 1 = freestream 2 = condition behind a normal shock wave Superscripts 1,2,... = number of iteration ( )* = nozzle throat ( ) =time derivative IntroductionI N a previous study, existing analytical plume flow models 1 " 3 were extended to deliver all flow quantities relevant to impingement calculations. Free-molecular plume flow was included by the definition of a freezing surface. 4 ' 5 In this DFVLR model, constant composition flow is assumed with mean constant gas properties. The angular plume flow description is most sensitive to changes in the ratio of specific heats of the exhaust gases.The present work is part of an extensive study of plume flow and impingement effects on spacecraft surfaces, that serves to test, verify, and improve the model by analyzing systematically the influence of the thruster nozzle geometry, nozzle boundary layer, and ratio of specific heats using pure gases. 6 This paper deals with experiments in real hydrazine plume flows from thr...

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A plume flow model has been developed to calculate the impingement forces and heat transfer caused by the firing of attitude control thrusters on satellites. Its validity for the inviscid core flow was tested and verified by an experimental study including pitot pressure and velocity measurements in plumes with different pure test gases. Measurements in real hydrazine thruster plumes allowed the determination of essential data (especially an effective ratio of specific heats) necessary for the description of this flowfield.x,y y b Nomenclature = cross section of nozzle exit = plume constant 1 = plume constant, determined by experiment = factor of proportionality in Eq. (12) = specific heats at constant pressure and volume, respectively = surface element = diameter of pitot probe = distance from nozzle throat to nozzle exit = mass flow = molecular mass = Mach number = surface normal vector = stagnation pressure = pitot pressure = pitot pressure on centerline = (specific) gas constant = nozzle Reynolds number, = radius, polar coordinate = nozzle throat radius = nozzle exit radius = stagnation temperature = velocity = velocity at nozzle throat = velocity at nozzle exit (isentropic flow) = velocity determined by experiment = maximum gas velocity V2/c/(/c-1) -R-T Q = virtual source point of streamlines = coordinates = plume boundary coordinate = boundary-layer thickness at nozzle exit angle, polar coordinate = angle between plume axis and streamline, which separates the isentropic core and the boundary-layer expansion region = maximum expansion angle = angle between surface normal and streamline ( Fig. 9) = ratio of specific heats -viscosity Presented as Paper 85-0933 at

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