Thermal Management at Hypersonic Leading Edges

The intense heat flux incident upon the leading edges of hypersonic vehicles traveling through the low earth atmosphere at speeds of Mach 5 and above requires creative thermal management strategies to prevent damage to leading edge components. Conventional thermal protection systems (TPSs) include the ablative coatings of NASA’s Mercury, Gemini, and Apollo vehicles and the reusable reinforced carbon-carbon (RCC) system of the Space Shuttle Orbiter. The ablative approach absorbs heat by endothermic transformation (phase and/or chemical change to the polymeric coating). The heat is dissipated from the vehicle as the single-use coating eventually vaporizes. The RCC approach manages the intense heat by operating at high temperatures and radiating heat to its surroundings. The effectiveness of both approaches is predicated on keeping the heat flux that impinges upon the susceptible aluminum airframe below a critical level.

This dissertation has explored an alternative metallic TPS concept which seeks to redistribute the heat from the leading edge, thereby eliminating local hot spots. It makes use of high thermal conductance heat pipes coupled to the leading edge so that the thermal load may be redistributed from a high heat flux location (at the stagnation point) to regions where it can be effectively radiated from the vehicle. The sealed system concept is based upon the evaporation of a fluid near the heat source that sets up a region of elevated vapor pressure inside the pipe. The latent heat is transported down the resulting pressure gradient by the vapor stream where it condenses at cooler regions, releasing the heat for removal. Replenishment of the condensed working fluid to the evaporator region is accomplished through the capillary pumping action of a porous wick which lines the interior surface of the pipe.

A design methodology for a wedge-shaped heat pipe is presented which uses a coupled flow-wall temperature model to construct design maps which relate design parameters of the leading edge system (overall length, wall thickness, and alloy) to its operating conditions (isothermal temperature, maximum temperature, maximum thermal stress). Potential bounds on heat transport due to physical phenomena linked to the sound speed within a chamber (sonic limit), capillarity, and boiling nucleation are considered by extending models developed for tube designs to the wedge geometry. A new heat flux limit is proposed which, should it be exceeded, subjects the leading edge to thermally-induced plastic deformation of the TPS.

To investigate the validity of the design approach and thermal spreading effectiveness of the proposed concept, a low temperature wedge-shaped leading edge was designed and constructed using stainless steel as the case material and water as the working fluid. Under localized tip heating, the maximum temperatures were significantly reduced compared to an otherwise identical but evacuated (no working fluid) test article. Isothermal operation was observed over its length. There was good agreement between experimental and design predictions. To test the concept at hypersonic flow enthalpies and temperatures, a high temperature Ni-based Inconel / sodium system was designed, fabricated, and tested. While there was a significant reduction in maximum temperature over an identical system containing no working fluid, isothermal operation was not observed. It is hypothesized that there is a lower bound on the wall heat flux which must be exceeded for the evaporated fluid to behave in the continuum flow regime predicted by the models. Finally, an assessment is made on three material-working fluid system combinations (Inconel/sodium, Nb-based C103/lithium, and Mo-based TZM/lithium) for a leading edge TPS that would be utilized by air-breathing hypersonic vehicles of the future.