Investigation of Aerodynamic Heating and Ablation of Thermal Protection Systems

The purpose of this thesis is investigation of aerodynamic heating and ablation of thermal protection systems for high speed missiles. A one-dimensional solver called AeroheataBS (Aerodynamic Heating and Ablation Simulation) has been developed for this purpose. Time dependent mission profile of a missile, flight velocity, altitude and angle of attack, is used as an input for the aerodynamic heating analysis. Atmospheric properties are used as a function of altitude for the computations. Convectional heating is calculated using external flow correlations and compressibility effects are included through Eckert Reference Temperature method. The flow properties at the boundary-layer edge are computed from shock relations and the surface pressure is evaluated using Modified Newton Theory. The Arrhenius equation is used for the calculation of thermo-chemical decomposition of the insulation material. Effective heat of ablation of the material is used for the surface recession rate. The reduction of convective heating due to injection of the ablation products into the boundary layer is included in the mathematical model. The explicit finite difference method is used for the temporal discretization of the governing equations. Temperature distribution along the nodal points in the body and insulation material is evaluated at each time point. The material properties are computed as a function of temperature. Validation of the methodology developed in the present study is performed through analytical methods where available, flight telemetry data, ground testing and CFD (Computational Fluid Dynamics). The flight data obtained from X-15 and HIFiRE-5 are used for validation but as well as the telemetry data taken from a newly developed missile where the thermal design was carried out by the present methodology, AeroHeataBS. The present predictions are compared with the data that is obtained from the inner surfaces of the missile by the telemetry. The predictions are compared with the published results of the FIAT computer code, which is well known and established software in the field. The predictions under the constant ablation rate are verified by the results obtained using the MSC MARC software. The results of the AeroHeataBS are additionally compared with the results of the CFD CHT (Conjugate Heat Transfer) computations on a generic test case. The results obtained for the conical and cylindrical regions are found in good agreement. Some very important model coefficients for an engineering correlation that account for transition from laminar flow to turbulent flow are derived and validated based on the flight test results. The ablative material characterization was performed using the results obtained from the oxyacetylene torch testing. Surface and in-depth temperature histories are measured using infrared thermal camera and thermocouples. Effective heat of ablation is calculated by use of mass loss data of specimens under the applied heat flux and is used as an input in the AeroheataBS simulations. As a result, it has been found that the developed methodology can be used in parametric thermal design of high speed missiles and launch vehicles depending on trajectory, geometry and material under aerodynamic heating. High accuracy computation of trajectory-dependent aerodynamic heating, temperature history and ablation within minutes of CPU time saves significant efforts in the design process.