Heat Transfer in Turbulent Boundary Layers of Conical and Bell Shaped Rocket Nozzles with Complex Wall Temperature

The objective of this article is to perform detailed analysis of heat transfer in accelerated supersonic nozzle flows with cooled walls. Since most of the heat transfer occurs near the nozzle walls, correct prediction of the boundary layer under strong adverse pressure gradient is therefore required to achieve high fidelity numerical prediction. In this study, a two-equation SST-V turbulence model is used in conjunction with a second-order explicit-implicit method to solve axisymmetric compressible Navier-Stokes equations. First, the effect of the inlet pipe diameter and the associated contraction area on the heat transfer is studied in nozzles having 15 degrees and 30 degrees diverging half-angles. Then, a series of computations are conducted to examine the efficiency of the use of a constant wall temperature as a function of the stagnation temperature in heat transfer calculations. The computations are performed for nominal stagnation pressure of 208N/cm(2) and stagnation temperature of 539K. The computed heat-transfer coefficients are compared to experimental data and a good agreement is found. A pronounced increase in the throat heat transfer coefficient peak is observed accompanied with a reduction in the contraction area ratio. Also, the peak of the heat transfer coefficient for the pipe inlet diameter of 7.8cm is found to be 70% higher than the one related to the pipe of 16.51cm diameter.