Surface Pressure Measurements on an Aerofoil in Transonic Flow

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DEN 302 Applied Aerodynamics
SURFACE PRESSURE MEASUREMENTS ON AN AEROFOIL IN TRANSONIC FLOW

Abstract
The objective of this exercise is to measure the pressure distribution across the surface on an aerofoil in a wind tunnel. The aerofoil is tested under several different Mach numbers from subsonic to supercritical. The purpose of measuring the pressure distributions is to assess the validity of the Prandtl-Glauert law and to discuss the changing chracteristics of the flow as the Mach number increases from subsonic to transonic. As a result of the experiment and computation of data, the aerofoil was found to have a critical Mach number of M=0.732. Below this freestream Mach number the Prandtl-Glauert law predicted results very successfully. However, above this value, the law completely breaks down. This was found to be the result of local regions of supersonic flow and local shockwaves. Contents

Abstract2
Apparatus2
1.Induction Wind Tunnel with Transonic Test Section2
2.Aerofoil model3
3.Mercury manometer3
Procedure3
Theory3
Results4
Discussion8
Transonic Flow8
Analysis9
Conclusion11
Bibliography11

Apparatus
1.Induction Wind Tunnel with Transonic Test Section
The tunnel used in this experiment has a transonic test section with liners, which, after the contraction, remain nominally parallel bar a slight divergence to accommodate for boundary layer growth on the walls of the test section. The liners on the top and bottom are ventilated with longitudinal slots backed by plenum chambers to reduce interference and blockage as the Mach number increase to transonic speeds. The working section dimensions are 89mm(width)*178mm(height). The stagnation pressure , p0∞, is close to the atmospheric pressure of the lab and with only a small error ,is taken to be equal to the settling chamber pressure. The reference staticpressure, p∞, is measured via a pressure tapping in the floor of the working section, well upstream of the model so as to reduce the disturbance due to the model. The ‘freestream’ Mach number, M∞, can be calculated by the ratio of static to stagnation pressure. The tunnel airspeed is controlled by varying the pressure of the injected air, with the highest Mach number that can be achieved by the tunnel being 0.88. 2.Aerofoil model

The model used is untapered and unswept, having the NACA 0012 symmetric section. The model chord length, c, is 90mm and the model has a maximum chord/thickness ratio of 12%. Non-dimensionalised co-ordinates of the aerofoil model are given in table 1 below. Pressure tappings, 1-8 , are placed along the upper surface of the model at the positions detailed in table 1. An additional tapping, 3a, is placed on the lower surface of the aerofoil at the same chordwise position as tapping 3. The reason for including the tapping on the lower surface is so that the model can be set at zero incidence by equalizing the pressures at 3 and 3a 3.Mercury manometer

A multitube mercury manometer is used to record the measurements from the tappings on the surface of the model. The manometer has a ‘locking’ mechanism which allows the mercury levels to be ‘frozen’ so that readings can be taken after the flow has stopped. This is useful as the wind tunnel is noisy. The slope of the manometer is 45 degrees. Procedure

The atmospheric pressure is first recorded, pat, in inches of mercury. For a range of injected pressures, Pj, from 20 to 120Psi, the manometer readings are recorded for stagnation pressure (I0∞), reference static pressure (I∞), and surface pressure form tappings on the model (In, for n=1-8 and 3a). Theory

These equations are used in order to interpret and discuss the raw results achieved from the experiment. To convert a reading, I, from the mercury manometer into an absolute pressure, p, the following is used: p=pat±l-latsinθ(1)

For isentropic flow of a perfect gas with γ=1.4, the freestream Mach number,M∞, is related to the ratio between the...
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