MX Missile Thermal Mapping and Surface Flow Results

Abstract

Results of wind tunnel tests to determine aerodynamic heating patterns and general flowfield characteristics on the MX missile are discussed. 5% and l2 l /2% scale models of the missile configuration were tested at Mach 6 and 8. Infrared scanning and phase change paint thermal mapping techniques were used to determine smooth wall turbulent and laminar heating levels and localized heating effects due to protuberances. Protuberances included a scaled raceway, lateral support pad shear tabs, and stage II roll control jets. The flowfield about the missile during the stage I/stage II separation event was investigated using a cold gas simulation of the stage II motor plume. Measured smooth wall heating rates confirmed both laminar and turbulent theoretical predictions and significant heating augmentation in excess of five times the smooth wall values was found in the vicinity of the leading edge of the raceway and the shear tabs. An extensive interaction region was produced by the roll control system jets. Schlieren photography of the staging event showed plume induced flow separation extending to the nose of the missile. NomenclatureA = area A ref = missile reference area D m = missile or model cross-sectional diameter h = heat-transfer coefficient (see Eq. 1), Btu/ft 2 -s-°R hi/hu = ratio of interference heating to undisturbed heating i.r. = infrared camera view £ = model or missile length M = Mach number p = static pressure P t = tunnel total pressure q -convective heat-transfer rate q = dynamic pressure RCS = roll control system Re? = freestream Reynolds number based on missile or model length Re/ft = unit freestream Reynolds number S = stage separation distance T r = recovery temperature T t = tunnel total temperature T w = model wall (surface) temperature a = model angle of attack 7 = ratio of specific heats Subscripts e = stage II nozzle exit £ = local external flow over missile or based on model/missile length oo = undisturbed freestream