Abstract
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.
Highlights
Conceptualizing and understanding the shock/boundary-layer interaction (SBLI) phenomena has been a challenging task for the scientific community since the dawn of supersonic flights, due to their effects on the performance of the vehicle and its parts
The diverse consequences of SBLIs are not restricted to excessive viscous dissipation, boundary layer degradation, and frequently to flow separation
The most apparent way for the SBLIs to occur is for an externally-produced shockwave to intrude on a boundary layer, formed over a surface
Summary
Conceptualizing and understanding the shock/boundary-layer interaction (SBLI) phenomena has been a challenging task for the scientific community since the dawn of supersonic flights, due to their effects on the performance of the vehicle and its parts. At the leading edge of a fin or in front of an isolated object attached to a surface, e.g., a vertical fin This type of compression corner generally creates a compression wave or a shockwave which has its root inside the boundary layer. The incoming high-speed air is compressed to a suitable combustor-friendly subsonic speed Even in these intentional-shock-utilization scenarios, the higher shock strength and their uncontrolled interactions with the viscous boundary layer might produce large drag. This incurs a higher cost to the propulsive efficiency. One of the vital reasons why the only supersonic passenger aircraft in history, the Concorde, was retired is due to its higher cost to propulsive efficiency
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