1 Lift Measurement from Pressure Wall Distribution Aashish Lamba, Abhishek Khandelwal, Abhishek Singh, Adarsh Chandra Thakur, Aditya Duhan, Akhil Srivas Indian Institute of Space Science and Technology, Thiruvananthapuram. The experiment was done in blower type wind tunnel in aerodynamics lab in IIST to determine the lift over an airfoil by measuring the pressure distribution on the top and bottom wall of the test section. Variation of Lift and C l with Re (keeping A O A constant) and A O A (keeping Re constant) were studied. It was found that the C l decrease with Reynolds number at a given A O A whereas lift and C l increase with angle of attack at a given Re. Nomenclature RPM = revolutions per minuteR e = Reynolds’s numberv= free stream velocity of the fluid C l = Lift Coefficient Cp = Coefficient of Pressure AoA = Angle of Attack I.Introduction n airfoil is part of body or surface, such as a wing, propeller blade, or rudder, whose shape and orientation control stability, direction, lift and drag. The geometry of the airfoil is described with a variety of terms. As a wing moves through air, the air is split and passes above and below the wing. The wing’s upper surface is shaped so the air rushing over the top speeds up and stretches out. This decreases the air pressure above the wing. The air flowing below the wing moves in a straighter line, so its speed and air pressure remains the same. Since high air pressure always moves toward low air pressure, the air below the wing pushes upward toward the air above the win g. The wing is in the middle, and the whole wing is “lifted.” The faster an airplane moves, the more lift there is. And when the force of lift is greater than the force of gravity, the airplane is able to fly. An airfoil-shaped body moved through a fluid produces an aerodynamic force. It seems complex but the forces generated are due to mainly two sources which are pressure and shear stress distribution over the body surface. The contribution due to shear is small as compared to pressure forces. As a wing moves through air, the air is split and passes above and below the wing. The wing’s upper surface is shaped so the air rushing over the top speeds up and stretches out. This decreases the air pressure above the wing. The air flowing below the wing moves in much straighter line, so its speed and air pressure remains the same. Thus the pressure difference between upper and lower surface is developed. When this pressure distribution is integrated over the surface of the airfoil we get resultant forces in two direction one parallel to the flow called Drag and other perpendicular to the flow called Lift. The resultant force can be slip in axial which is along the chord and normal which is perpendicular to chord. A Figure 1. Forces on the ai rfoil
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The pressures on the top and bottom walls are denoted
by Pu(x) and Pl(x) respectively. The walls are assumed to be
close enough to the model so that the pressure on the wall is not
necessarily P∞. Also assume that the surfaces ai and bh are far
enough upstream and downstream such that P=P∞, and that thechange in orientation of the airfoil, i.e. changing the angle of
attack does not change the pressure distribution on the ai and bh.
Now using the Reynolds Transport Theorem on the control
volume shown by control surface ‘abcdefghi’ we can show that:
Where,
L’ = Lift on the airfoilPl(x) = pressure distribution on the lower wall with respect to x
Pu(x) = pressure distribution on the upper wall with respect to x
III.
Experimental Setup
A blower type wind tunnel on a modular flow apparatus is used to conduct the
experiment which is having control setup and maximum speed of 1200rpm.The end
of the wind tunnel was placed with the test section consisting of a symmetric airfoil
with possible angle adjustments.
Measurement at various points of the test section is done using a multi tubemanometer which is connected to pressure sensor nodes, the nodes are further
connected to top and bottom surfaces of the test section.
IV. Procedure
The whole experimental setup is based on the theory of force equilibrium of a control volume that is the force
due to pressure difference between the upper and lower pressures in the test section has to be equal to the lift on the
airfoil present in the test section. In the above control volume equilibrium, assumption is the test section is
sufficiently larger than the airfoil.
The airfoil is placed in the test section and at the horizontal that is at 0° angle of attack. The wind tunnel is
maintained at 500rpm.Intially angle of attack is 0° and noted the value of pressure of all the nodes at upper and
lower surface. Again repeat with 3° ,6°,9° ,12°,13° and 15°.Then repeat this experiment with rpm varying from 600