Two Phase Thermal Protection of the Hypersonic Leading Edge

Abstract

Hypersonics is the flight regime characterized by conditions in which high temperature and extreme heat flux dominate the flow physics. This regime spans high speed aircraft around Mach 5 through atmospheric entry of spacecraft at Mach 25 to 30. Stagnation temperatures frequently climb into the many thousands of degrees, well beyond the melt temperature of any known materials. Conventional thermal protection systems for the most extreme conditions employ ablative shielding, making use of latent heat for phase transition and gaseous advection to maintain vehicle airframe temperatures to within a manageable range for the short duration of entry. More recently, space planes and high-lift entry vehicles have enabled significantly lower heating by bleeding off airspeed at higher altitudes before descending into the dense lower atmosphere. These lower heat fluxes are managed with temperature resistant, extremely-low-thermal-conductivity heat shields during the transient entry process. The concepts explored in the present work include variations on past efforts and incorporation of existing component technologies into new applications, particularly in the context of liquidvapor evaporation-condensation cycles for thermal energy transport. This paper is intended as concept map summary, providing flight conditions and feasibility for two-phase thermal protection systems, including conventional ablation, transpiration, internal spray cooling, thermosyphons and heat pipes, loop heat pipes, and solid-liquid phase change media. With heat transport demonstrated into the kilowatts per square centimeter, two-phase systems provide a promising class of reusable and continuous technologies for managing the extreme heat fluxes encountered by hypersonic vehicles. Nomenclature= entry vehicle reference area, [m 2 ] = leading edge wall thickness, [m] = drag coefficient, [-] = friction coefficient, [-] = Stanton number, [-] = lift coefficient, [-] = linear elasticity modulus, [Pa] = drag force, [N] = friction force, [N] = lift force, [N] = gravity, [m/s 2 ] ℎ = altitude, [m] ℎ ( ) = (adiabatic) wall enthalpy, [J/kg] = thermal conductivity, [W/(m-K)] = Boltzmann constant, [m 2 kg/(s 2 K)] 1 Mechanical Engineer, Thermal Systems and Analysis Section 2 Branch Head, Space Mechanics Systems Development Branch 3 President. / = lift to drag ratio, [-] = free stream Mach number, [-] = vehicle mass, [kg] = heating rate, [W] = flight dynamic pressure, [Pa] = wall heat flux, [W/m 2 ] = vehicle reference area, [m 2 ] = temperature, [K] = velocity, [m/s] = thermal diffusivity, [m 2 /s] = ballistic coefficient, [kg/m 2 ] /( / ) = equilibrium glide entry parameter, [-] = ratio of specific heats, [-] = heat dissipation efficiency, [-] = enthalpy of vaporization, [J/kg] Downloaded by PURDUE UNIVERSITY on June 24, 2016 | http://arc.aiaa.org | This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States. AIAA Aviation = dynamic viscosity, [J/kg] = Poisson's ratio, [-] = density, [kg/m 3 ] = surface ...