In order to achieve proficient combustion with the present technologies, the flow through an aircraft intake operating at supersonic and hypersonic Mach numbers must be decelerated to a low-subsonic level before entering the combustion chamber. High-speed intakes are generally designed to act as a flow compressor even in the absence of mechanical compressors. The reduction in flow velocity is essentially achieved by generating a series of oblique as well as normal shock waves in the external ramp region and also in the internal isolator region of the intake. Thus, these intakes are also referred to as mixed-compression intakes. Nevertheless, the benefits of shock-generated compression do not arise independently but with enormous losses because of the shockwave and boundary layer interactions (SBLIs). These interactions should be manipulated to minimize or alleviate the losses. In the present investigation a wall ventilation using a new cavity configuration (having a cross-section similar to a truncated rectangle with the top wall covered by a thin perforated surface is deployed underneath the cowl-shock impinging point of the Mach 2.2 mixed-compression intake. The intake is tested for four different contraction ratios of 1.16, 1.19, 1.22, and 1.25, with emphasis on the effect of porosity, which is varied at 10.6%, 15.7%, 18.8%, and 22.5%. The introduction of porosity on the surface covering the cavity has been proved to be beneficial in decreasing the wall static pressure substantially as compared to the plain intake. A maximum of approximately 24.2% in the reduction in pressure at the upstream proximal location of 0.48 L is achieved in the case of the wall-ventilated intake with 18.8% porosity, at the contraction ratio of 1.19. The Schlieren density field images confirm the efficacy of the 18.8% ventilation in stretching the shock trains and in decreasing the separation length. At the contraction ratios of 1.19, 1.22, and 1.25 (‘dual-mode’ contraction ratios), the controlled intakes with higher porosity reduce the pressure gradients across the shockwaves and thereby yields an ‘intake-start’ condition. However, for the uncontrolled intake, the ‘unstart’ condition emerges due to the formation of a normal shock at the cowl lip. Additionally, the cowl shock in the ‘unstart’ intake is shifted upstream because of higher downstream pressure.

/pdf/effects-of-wall-ventilation-on-the-shock-wave-viscous-layer-1pep2j94i9.pdf

Effects of Wall Ventilation on the Shock-wave/Viscous-Layer Interactions in a Mach 2.2 Intake

A comprehensive analysis on the structure of groove-induced shock waves in a linear turbine

A computational study has been carried out to assess the effectiveness of a porous medium as a passive control device suitable for reducing the drag in a normal-shock-wave/boundary-layer interaction at transonic speeds with a view toward application in aircraft wings. Reduction in overall drag is achieved via recirculation inside the porous medium, which primarily weakens the shock structure and hence reduces the wave drag. The study has been carried out for a Mach 1.3 normal-shock-wave/boundary-layer interaction on a flat plate in the presence of a porous medium beneath the region of interaction. The computations are performed as steady-state RANS calculations using Menter’s SST $$k-\omega /k-\epsilon $$
 model for turbulence closure. A parametric study that investigates the dependency of the effectiveness of control on dimensions of the cavity (length and depth), relative position of the cavity, and porosity of the medium has been carried out. It is observed that the change in overall drag is pronounced for parameters which result in significant changes to the size of the lambda-shock structure, such as the length of the cavity upstream of the inviscid shock location. Among the parameters investigated, porosity is seen to strongly affect the boundary-layer properties, with increase in porosity resulting in higher viscous drag.

Drag reduction in transonic shock-wave/boundary-layer interaction using porous medium: a computational study

Numerical Investigation of Transonic Flow over Porous Medium Using Immersed Boundary Method

In this chapter, we discuss the development of a sharp-interface immersed boundary method for compressible, turbulent flows and its application to transonic/supersonic flows. A direct-forcing-type immersed boundary method is elucidated wherein the flow properties in the immediate neighbourhood of the immersed surface are reconstructed using inverse distance-based interpolation procedures. The flow is assumed to be locally parallel to the immersed surface, and the tangential velocity in the vicinity of the immersed surface is assumed to obey a power-law function of the local immersed surface normal. This approach helps in mimicking the energising effects of turbulent boundary layers without excessive mesh refinement near the immersed surface for suitable choices of the power-law coefficient. Temperature reconstruction is achieved from considerations of temperature variation in compressible thermal boundary layers, and density is estimated by either solving the continuity equation or by interpolation. The turbulence variables are reconstructed using law-of-the-wall-type approach. The application of the outlined immersed boundary method to the simulation of flow control devices is also discussed. Additionally, interpolation procedures for reconstructing the pressure and shear stress at the immersed surface and its application to simple cases are also presented. This information can be useful for comparison with experimental data, performing fluid–structure interaction studies, and also identifying flow-separation and re-attachment locations on the immersed surface.

Karthik Subramaniam

Papers

Passive control of normal-shock-wave/boundary-layer interaction using porous medium: Computational study

Positioning of Normal Shock in a Novel Constant-Area Test Section: A Numerical Study