| Summary: | Abstract Numerical studies were carried out to investigate the thermal performance of the muzzle flow with strong shock discontinuities and gas–gas heat transfer mechanism in the base bleed unit (BBU) of an aerodynamic vehicle under the combined effects of muzzle flow and depressurization process. The process was modeled using the Navier–Stokes equations with chemical kinetics and interior ballistic equations. The results show that the muzzle phenomena with large temperature gradients and different wave structures appear after the supersonic vehicle enters into the free‐flight stage. The BBU also experiences a strong depressurization process with the release of high‐temperature propellant gas at 2200 K. However, the igniter combustion gas at a higher temperature of 3050 K in the BBU develops downstream and upstream repeatedly under the combined effects of the base wave and depressurization process, which makes the energy conversion process in the unit more complicated. With the projectile leaving the control of the muzzle flow, the BBU is full of the igniter combustion gas and the energy in the unit is uniformly distributed gradually. The high‐temperature area, at approximately 2750 K, propagates downstream along the axial and radial directions simultaneously until it finally dominates the entire unit. The maximum temperature of the combustion gas acting on the propellant surface is always 2740 K and the maximum heat flux remains approximately 1250 W/cm2 at later periods.
|
|---|