Nozzles are generally designed to produce thrust in a direction perpendicular to its exit plane. In high speed flight vehicles, rapid control of thrust direction (i.e. thrust vectoring) is an obvious requirement for high performance aerospace applications. Shock vector control is one of the efficient ways to achieve this thrust vectoring. In this present study, a bypass mass injection is used to generate shock vectoring in a planar supersonic Converging-Diverging (CD) nozzle. Injection was done by a 10mm×10mm square channel which is kept perpendicular at the diverging section of the nozzle. Investigation is carried out at Nozzle Pressure Ratio (NPR) of 2.4 which gives an overexpanded nozzle flows. The flow conditions and the size of the injection channel gives a bypass mass flow ratio of 4.9%. Reynolds-averaged Navier-Stokes (RANS) equations with k-omega SST turbulence model have been implemented through numerical computations to capture the three-dimensional steady characterstics of the flow field. Results showed a significant change in the shock structure with the fromation of recirculation zone near the bypass injection port. Consequently, a change in mass flow direction has been obtained which results a considerable thrust vectoring in a supersonic nozzle.
展开▼
