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Pressure Distribution on an Aerofoil

physic

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Pressure Distribution on an Aerofoil

School of Engineering



Name of Laboratory Aerodynamics Laboratory

Laboratory Number Lab 49

Code for Experiment A1

Underpinning Module AE 2010

Title of Experiment Pressure Distribution on an Aerofoil

 

Aims

The aim of this experiment is to examine the pressure distribution on an aerofoil and its variation with incidence

 

Learning Outcomes

 

i)                    To be able to produce non-dimensional pressure distribution plots from a wind tunnel test using a spreadsheet.

ii)                  To examine the effect of incidence on pressure distribution.

iii)                To determine the lift coefficient from pressure data.

 

 

 

Introduction

An aerofoil is the two-dimensional cross section of a wing, tail or helicopter rotor blade. The lift and pitching moment on an aerofoil is determined by the pressure distribution on it. The pressure distribution changes with the angle of incidence.

Understanding the behaviour of aerofoils requires some understanding of the variation of the pressure distribution with incidence.

Apparatus

This experiment uses a pressure tapped aerofoil mounted inside a low speed open return wind tunnel. The pressure tapings are connected to a multi-tube manometer which is also used to measure the dynamic pressure.

You should check and record all the connections on the multitube manometer.

Procedure:

1.                 Level and adjust the multitube manometer in the vertical position with the indicator registering 90, then incline it to approximately 40 to the horizontal.

2.                 Check and record all of the manometer connections.

3.                 Record the NACA aerofoil designation of the aerofoil.

4.                 Set the model at zero incidence and turn the wind tunnel on. Run the fan up to about of its full speed.

5.                 Record all of the manometer heights.

6.                 Adjust the angle of incidence to -4 and repeat the previous step. Adjust the motor control if necessary to maintain a constant dynamic pressure.

7.                 Repeat the previous step for 4, 8 and 12

8.                 Determine the stall angle of the aerofoil and qualitatively observe the changes in pressure distribution near the stall angle.

Theory:

The non-dimensional pressure coefficient is defined as:

Where




Pi - pressure at tapping i

P - free stream pressure

ρ - air density

V - free stream velocity

S - wing area

c - aerodynamic mean chord

The quantity , is known as the dynamic pressure and for low speeds it is equal to the difference between the stagnation pressure Po and the free stream pressure P. This means that we calculate CP using the relation:

For an open return wind tunnel, we may assume that the stagnation pressure is equal to the room pressure and we measure the free stream pressure in the test section.

In this experiment we measure pressure differences using a multitube manometer. The difference in pressure is proportional to the difference in height of the liquid levels in the manometer. Since the pressure coefficient is a ratio of two pressure differences, it is also equal to the ratio of differences in height of liquid levels.

where h indicates the height of the liquid in the manometer.

The multitube manometer is inclined, which means that it does not read the difference in height directly. This is done to increase the difference in manometer reading. Since all of the tubes are inclined at the same angle, the difference in manometer reading is proportional to the difference in height and we can write:

where r indicates the manometer reading.

Aerofoil pressure plots are usually given with CP on the vertical axis, with negative values above the axis and on the horizontal axis.

Force Coefficients:

The force coefficient on an aerofoil may be determined from the non-dimensional pressure plot. A typical pressure plot is shown below:


The force coefficient perpendicular to the aerofoil is the area enclosed by the figure. The lift is defined as the force perpendicular to the airflow, not the aerofoil, but for small angles of attack, the difference is negligible.

Analysis of Results

1.           Plot graphs of Cp Versus X/C for your five sets of readings. ( -4 , 0, 4, 8, 12 ) this is done most easily using a spreadsheet.

2.           For each angle of attack, determine the lift coefficient. Clearly explain how you evaluated the required areas.

3.           Comment on the effect of angle of attack on the pressure distribution. In particular note the variation in the magnitude and the position of the point of minimum pressure.

4.           Briefly describe the variation in pressure distribution with angle of attack near and beyond the stalling angle.

References:

1. Fundamentals of Aerodynamics - John D. Anderson, Jr. - McGraw-Hill, Inc.

ISBN:0-07-001679-8.

2. Low speed Wind Tunnel Testing-William H.Rae, ISBN: 047 1874027.








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