Characteristics of the Supersonic Flows Over 3-D Bump

Abstract

In supersonic flows, bumps were used for the injection of fuel into the flow field due to its high mixing efficiencies without much change in the upstream conditions. Research on wave drag reduction using contour bumps in transonic aircraft wings has been an active research topic in the aerospace sector in recent years. It was estimated that about 5–15% of wave drag reduction could be achieved in a transonic aircraft with rounded contour bumps installed on to its wing surfaces. Although there are many types of research in this field, the underlying flow physics of rounded contour bumps is less well understood. Although using contour bumps could provide desire performance in drag reduction and high total pressure recovery in transonic and supersonic aircraft, it is known that adverse effects can be induced by flow separation and spanwise vortices formation appear downstream of the bump crest of the bumps. As a result, it is important to investigate the flow separation characteristics of contour bumps to have a better understanding of the physics of bump flow. The present study aims to simulate the flow conditions computationally using the finite volume solver. The flow structure and the spanwise flow patterns are analyzed for a better understanding of the flow physics when we introduce the injection to the flow field. The results show that injection effectiveness can be increased with increasing jet total pressure ratio.