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In aircraft engines operating at high Mach numbers, it is exigent to reduce the magnitude of flow speed from supersonic to subsonic before entering the burner, to accomplish a proficient ignition. It is accomplished by a progression of oblique shocks as well as a normal shock wave in supersonic intakes. However, the advantage of shock enabled compression in intakes does not come independent but with colossal losses due to shock-boundary layer interactions, which includes intake unstart and an abrupt thickening/separation of the boundary layer. Controlling these interactions by boundary layer control using micro-vortex generators (MVGs) has gained prominence. In the present study, an attempt is made to control the interactions at the shock impact point in a Mach 2.2 mixed compression intake. Two types of MVGs; a conventional configuration and an innovative ramped-vane configuration were experimentally investigated by varying the MVG heights as 600, 400, and 200 μm. Also, the near-field effects of MVGs are quantified in terms of static pressure variation in the internal flow. It is found that the innovative MVGs of height 200 μm leads to favorable pressure drop in both the upstream and downstream region, due to enhanced flow mixing near the boundary layer. © Springer Nature Singapore Pte Ltd 2021.

Department of Thermal and Energy Engineering

School of Mechanical Engineering

Vellore Campus

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Lecture Notes in Mechanical Engineering

Near-field effectiveness of the sub-boundary layer vortex generators deployed in a supersonic intake

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. © 2020 by the authors.

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Effects of wall ventilation on the shock-wave/viscous-layer interactions in a mach 2.2 intake

In the present study, the flow characteristics of the commercially used S809 wind turbine airfoil controlled with triangular counter-rotating micro-vortex generators at stall angle of attack of 15 degrees and 10 m/s, 15 m/s and 20 m/s (speed range used in the wind turbine applications) had been computationally investigated. In addition to the controlled airfoil, an uncontrolled airfoil was also studied for the comparison. The modelling and analysis had been carried out using incompressible, Reynolds Averaged Navier Stokes equation using Spalart-Allmaras one equation turbulence model. The numerical computations were performed with SIMPLE algorithm. The velocity profiles at different locations on the suction surface were plotted for both uncontrolled and controlled airfoils. The shear stresses exerted on the upper surface of the airfoil in both the configurations were also compared. It is found that the controlled airfoil, the shear stress distribution was greatly increased near to trailing edge of the airfoil revealing the superiority of vortex generators in increasing the efficiency of wind turbine by delaying boundary layer separation. The qualitative results of flow visualization in the spanwise direction also support the quantitative findings of velocity profiles and shear stress distribution. © Published under licence by IOP Publishing Ltd.

IOP Conference Series: Materials Science and Engineering

CFD Studies on Triangular Micro-Vortex Generators in Flow Control

In jet engines, it is imperative to bring down the flow speeds from supersonic to subsonic level before entering the burner, in order to have efficient combustion. It is achieved by a series of oblique as well as normal shock waves in supersonic intakes. However, the advantage of shock enabled compression in intakes does not come standalone but with colossal losses due to shock-boundary layer interactions, which includes intake unstart, abrupt thickening/separation of boundary layer, etc. Here, an attempt is made to control the interactions at the shock impact point in a Mach 2.2 mixed compression intake. Passive control in the form of a shallow cavity covered with porous surface, is employed at the shock impingement point. The porosity of the surface is varied as 10%, 15.7% and 22.5% and their effect in controlling the SBLIs at different intake contraction ratios of 1.14, 1.19 and 1.25 is studied. The intake controlled by cavity with 15.7% porous surface reduces the interaction losses. These results are found to be in good agreement with the qualitative data. © 2017, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.

35th AIAA Applied Aerodynamics Conference, 2017

Shock boundary layer interaction control in supersonic intake using cavity with porous surface

In jet engines, it is imperative to bring down the flow speeds from supersonic to subsonic level before entering the burner, in order to achieve efficient combustion. It is achieved by a series of oblique as well as normal shock waves in supersonic intakes. However, the advantage of shock enabled compression in intakes does not come standalone but with colossal losses due to shock-boundary layer interactions, which includes intake unstart and abrupt thickening/separation of boundary layer. Controlling these interactions by boundary layer control using micro-vortex generators (MVGs) has gained prominence. Here, an attempt is made to control the interactions at the shock impact point in a Mach 2.2 mixed compression intake. Two types of MVGs; a conventional and an innovative ramped-vane configuration were experimentally investigated. In both the configurations, the height of the MVGs are varied as 600μm; 400μm and 200μm. The innovative MVGs of height 200μm, showed enhanced flow mixing near the boundary layer and minimized interaction losses. These results are found in good agreement with the qualitative data. © 2017, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.