An investigation was conducted to study the feasibility of cooling hypersonic vehicle leading edge structures exposed to severe aerodynamic surface heat fluxes using a combination of liquid metal heat pipes and surface mass transfer cooling techniques. A generalized, transient, finite difference based hypersonic leading edge cooling model was developed that incorporated these effects and was demonstrated on an assumed aerospace plane-type wing leading edge section and a SCRAMJET engine inlet leading edge section.; The hypersonic leading edge cooling model was developed using an existing, experimentally verified heat pipe model. In this study, the existing heat pipe model was modified by adding both transpiration and film cooling options as new surface boundary conditions. The models used to predict the leading edge surface heat transfer reduction effects of the transpiration and film cooling were modifications of more generalized, empirically based models obtained from the literature.; Two applications of the hypersonic leading edge cooling model were examined. An assumed aerospace plane-type wing leading edge section exposed to a severe laminar, hypersonic aerodynamic surface heat flux was studied. In the model a one inch nose diameter leading edge structure was cooled using a lithium filled heat pipe supplemented by either surface transpiration, surface film, or internal active heat exchanger cooling while executing a 2000 psf constant dynamic pressure hypersonic ascent flight trajectory. Surface coolants used in the study were gaseous air, helium and water vapor. The results of applying the cooling model to this case included transient structural temperature distributions, transient aerodynamic heat inputs and transient surface coolant distributions. The results indicated that these cooling techniques limited the maximum leading edge surface temperatures and moderated the structural temperature gradients.; A second application of the hypersonic leading edge cooling model was conducted on an assumed one-quarter inch nose diameter SCRAMJET engine inlet leading edge section exposed to both a transient laminar, hypersonic aerodynamic surface heat flux and a Type IV shock interference surface heat flux. These results indicated that the combination of liquid metal heat pipe cooling and surface transpiration or film cooling tended to mitigate the otherwise severe maximum leading edge surface temperatures expected on a SCRAMJET engine inlet structure exposed to these environments.; The investigation led to the conclusion that cooling leading edge structures exposed to severe hypersonic flight environments using a combination of liquid metal heat pipe, surface transpiration and film cooling methods appeared feasible.
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