Pressure Management on a Supercritical Airfoil in Transonic Flow

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  • Topic: Aerodynamics, Fluid dynamics, Mach number
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  • Published : March 21, 2013
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Pressure Management on a supercritical aerofoil in transonic flow

Abstract-At transonic speeds an aerofoil will have flow accelerate onwards from the leading edge to sonic speeds and produce a shockwave over the surface of its body. One factor that determines the shockwave location is the flow speed. However, the shape of an aerofoil also has an influence. The experiment conducted compared Mach flow over a supercritical aerofoil (flattened upper surface) and a naca0012 aerofoil (symmetrical). Despite discrepancies, the experiment confirmed the aerodynamic performance of a supercritical aerofoil being superior to a conventional aerofoil. A comparison of the graphical distributions demonstrates the more even pressure distribution on a supercritical aerofoil and a longer delay in shockwave formation. All of which, reflects the theory.

Table of Contents

Introduction3
Apparatus3
Induction Wind Tunnel with Transonic Test Section3
Mercury Manometer4
Procedure4
Theory and Equations5
Results6
Discussion10
Theory of Transonic Flight10
Relating the Theory to the Experiment11
Effectiveness of Supercritical aerofoils……………………………………………………………………...11 Limitations and Improvements12
Appendix13
References14

Introduction
For any object travelling through a fluid such as air, a pressure distribution over all of its surface exists which helps generate the necessary lift. Lift is an aerodynamic force which is perpendicular to the direction of the aerofoil. Transonic speeds result in the formation of shockwaves over the top surface of the aerofoil. This is due to accelerated flow over the surface region. We say this region is approximately between 0.8-0.9. Since the flow must accelerate and then will lose velocity following the shockwave the aerofoil will have a subsonic and sonic region. For the majority of commercial airlines this is not a desired region to cruise at given the instantaneous pressure distribution which passengers would otherwise experience. Particularly, the formation of shock induced boundary layer separation. Supercritical aerofoils are more efficient designed for higher Mach speeds and drag reduction. They are distinct from conventional aerofoils by their flattened upper surface and asymmetrical design. The main advantage of this type of aerofoil is the development of shockwaves further away then traditional aerofoils and thus greatly reducing the shock induced boundary layer separation. In order to truly understand the effectiveness of a supercritical aerofoil an experiment gathering supercritical aerofoil performance and raw data of a naca0012 aerofoil will be extensively analysed and compared. Following the calculation and procedureit will be assessed whether a supercritical aerofoil is more effective. Apparatus

A wind tunnel with a transonic test section was used in this experiment to study transonic flow around an aerofoil. The test section consists of liners which, after the initial contraction, are nominally parallel apart from a slight divergence to compensate for growth of the boundary layers on the wall. In order to reduce interference and blockage at transonic speeds, the top and bottom liners are ventilated by longitudinal slots backed by plenum chambers. The working section has a height and width of 178mm and 89mm respectively. The stagnation pressure, p0∞, in the tunnel is close to atmospheric pressure, and therefore it can be taken to be equal to the settling-chamber pressure as the errors are only small. To minimise the disturbance due to the model itself, the reference stagnation pressure, p∞, is taken from a pressure tapping in the floor of the working-section, well upstream of the model. The nominal ‘free-stream’ Mach number, M∞, in the tunnel can be calculated from the ratio p∞/p0∞. The Mach number in the tunnel can be controlled by varying the pressure of the injected air, pj. The maximum Mach number that the tunnel can achieve is about 0.88

Mercury...
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