Numerical analysis of flow field over compound delta wing at subsonic and supersonic speeds

Abstract

A great number of supersonic aircrafts use delta wings. Since 1950, delta wings have been studied well and it has been observed that there appear two large counter-rotating leading edge vortices when delta wing structure flies at high angles of attack. These vortices are responsible for additional lift and they also provide a very high stall incidence to the wing. At high Mach numbers, compressibility advances leading edge separation and also expands the magnitude of the primary vortex. Shock waves appearing due to high-speed flow over the delta wing result in drastic changes in the flow characteristics. Although supersonic jets use delta wing configurations but most of the time they fly at subsonic speeds and hence the focus of present study is on both subsonic as well as supersonic regimes. The present study consist of numerical simulations of the flow-field over a compound delta wing at various Mach numbers ranging from 0.3 to 2.0 and angles of attack ranging from 0° to 15° using the computational model established in previous research [1]. Reynolds Averaged Navier-Stokes (RANS) based steady-state computations were carried out. Spalart Allmaras (SA) turbulence model is considered due to its less computational cost and good performance for simulating external flows. Since the study involves supersonic Mach numbers, compressibility was also incorporated in the computational model. The study with various freestream Mach numbers shows that there is a sudden change in flow fields with an increase in the Mach number for the range of 0.8 to 1.3. The flow near to the delta wing surface beneath the primary vortex becomes completely supersonic and shock waves appear when the flow reaches Mach number 0.85. Due to this shock wave formation, flow fields become more complex. At high Mach number because of the stronger magnitude of the primary vortex which prevents the evolution of the inner separation, the secondary vortex disappears. The results also confirm the hypothesis of vortex breakdown which is also responsible for the nonlinear behavior of flow characteristics over the wing when there is an increase in Mach number.