Aerofoil Experiment Pressure distribution over a NACA 2415 aerofoil
Elankumaran Nagarajan
20th October 2013
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Summary
Aerofoils are the important lift creating structure. The main objective of this experiment was to measure the pressure distribution over a NACA 2415 aerofoil for a range of angles of attack, to calculate the lift coefficient for the aerofoil and to experimentally investigate the effects created by a leading edge slat. The experiment was carried out by placing a NACA 2415 aerofoil in the wind tunnel and the air was passed over the aerofoil. For different angles of attack the lift coefficient of the aerofoil was recorded using the computer. Then the experiment was repeated by using a leading edge slat. Through calculations the Reynolds number of the flow was calculated to be 1.72*10 6. The lift coefficient of the aerofoil with slat was higher than the lift coefficient of the aerofoil without slat. The maximum lift coefficient of the aerofoil with the slat was found to be 1.459 and the maximum lift coefficient without the slat was found to be 1.226.
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INDEX
.
Page no
List of Symbols…………………………………………………………………………………………4
Introduction……………………………………………………………………………………………4
Experimental Procedure…………………………………………………………………………5
Results…………………………………………………………………………………………………...6
Discussion……………………………………………………………………………………………..10
Conclusion……………………………………………………………………………………………11
Appendix 1: Boeing 747 Questions………………………………………………………..12
Appendix 2: Data………………………………………………………………………………….15
References……………………………………………………………………………………………17
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List of Symbols
P = static pressure measured at surface
P = free stream static pressure
(U2)/2 = dynamic pressure of the free stream
L = lift force
C = the aerofoil chord
= angle of attack
S = wing area
CL = coefficient of lift
CP = coefficient of pressure
Introduction An aerofoil is the shape of a wing or a blade or a body that produces an aerodynamic force when moved through a fluid. Any object with an angle of attack in a fluid experience an aerodynamic force called lift perpendicular to the flow. Aerofoils are the most efficient lifting shapes among them, able to generate more lift and to generate lift with less drag. Aerofoil shapes are found in the fixed wings of the aircraft, vertical and horizontal stabilizers of the aircraft, helicopter rotor blades, turbines, compressors, fans, propellers and etc.
Aerofoils are the important cornerstone of aeronautical research and development . Aerofoil design is the major facet of aerodynamics. From its very beginning, the National Advisory Committee for Aeronautics (NACA) recognized the importance of aerofoils. By 1920, the Committee had published a compendium of experimental results from various sources (ref. 2) and Shortly thereafter, the development of airfoils by the NACA was initiated at the Langley Memorial Aeronautical Laboratory (ref. 3). The first series of airfoils, designated "M sections" for Max M. Munk, was tested in the Langley VariableDensity Tunnel (ref. 4). This series was significant because it represented a systematic
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approach to airfoil development as opposed to earlier, random, cut-and-try approaches. This empirical approach, which involved modifying the geometry of an existing airfoil, culminated in the development of the four- and five-digit-series airfoils in the mid 1930's (refs. 5-7).
Concurrently, Eastman N. Jacobs began work on laminar-flow airfoils. Inspired by discussions with B. Melvill Jones and G. I. Taylor in England, Jacobs inverted the airfoil analysis method of Theodore Theodorsen (ref. 8) to determine the airfoil shape that would produce the pressure distribution he desired (decreasing pressure with distance from the leading edge over the forward portion of the airfoil). This pressure distribution, it was felt, would sustain laminar flow. Thus, the basic idea behind modern airfoil design was conceived: the desired boundary-layer characteristics result from the pressure distribution, which results from the airfoil shape.
The main objectives of this experiment were to measure the pressure distribution over a NACA 2415 aerofoil for a range of angles of attack, calculate the lift coefficient for the aerofoil and compare with published NACA data, experimentally determine the effects created by a leading edge slat and to under stand the aerofoil characteristics in terms of fundamental fluid dynamics.
Experimental procedure
For the experiment only the lift forces were calculated. As the lift forces are dominated by pressure forces, the shear stress distribution was disregarded. The Bernoulli’s equation for an incompressible, inviscid fluid is given in equation (1).
P +( U
2/2)
= P+( U2/2)…………………….(1)
The local static pressure at any point on the aerofoil in non-dimensional terms of a coefficient of pessure, Cp (ref 1) is shown in equation (2).
Cp = (P - P)/( U
2/2)……………………….…(2)
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The lift force can be written in terms of a coefficient by dividing the free stream dynamic pressure (ref 1) as shown in equation (3).
CL = L/( U
2S/2)……………………………...…(3)
The NACA 2415 aerofoil was used for the whole experiment. The NACA 2415 aerofoil (chord 127 mm) was placed in the working section of the 0.3 m open- return circuit wind tunnel. The test section walls acted as end plates to maintain two dimensional flow over the wing. The wing was supported by two internal spigots passing through the bushes in the Perspex windows of the test section and a clamp allowed the aerofoil to be set any angle of attack within the range of 30, measured using a pointer and protractor. The airspeed was measured using a pitot-static tube upstream of the model. The wing was fitted with 33 pressure tappings in one chordal plane and the pressure distribution over the aerofoil was measured using a computer controlled scanivalve unit and transducer.
Before the experiment it was ensured that the pressure tubes to the model and the pitot-static tube were connected correctly. Then the tunnel was started and the speed was stabilized to approximately 20 m/s. The aerofoil model was adjusted to different angles of attack and the Cp and CL data were collected using the lab view programme in the computer. Then the leading edge slat (the leading edge slat was based upon the highly cambered NACA 22 aerofoil with chord of 38.1 mm) was attached the aerofoil and the experiment was repeated over a range of high angles of attack. The C p and CL data were collected.
Results From the data collected (shown in table 1) during the experiment, C L was plotted versus the angles of attack for both aerofoil (with and without slats). The NACA reference data (shown in table 2) was also included in the above graph and shown in figure 1.
Pressure arrow diagrams for the aerofoil at angles of attack 2, 8 and 15 are shown in figure 2, figure 3 and figure 4 respectively. Pressure arrow diagram for the aerofoil with
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slat at an angle of attack of 15 is shown in figure 5. The Reynolds number of the flow in the wind tunnel was calculated to be 1.72*106 (appendix 1, question 1a).
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1.5
1
CL without slat
0.5
CL with slat NACA CL (Re-3*10^6) NACA CL (Re-6*10^6) 0 -30
-20
-10
NACA CL (Re-9*10^6) 0
10
20
30
-0.5
-1
-1.5
Figure 1: CL versus for the NACA 2415 aerofoil with and without slat. NACA reference data included.
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Figure 2: pressure arrow diagram for NACA 2415 aerofoil at = 2
Figure 3: pressure arrow diagram for NACA 2415 aerofoil at = 8
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Figure 4: pressure arrow diagram for NACA 2415 aerofoil at = 15 (without slat)
Figure 5: pressure arrow diagram for NACA 2415 aerofoil at = 15 (with slat)
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Discussion From figure 1, it could be understood that the aerofoil lift coefficient increases linearly with the angle of attack up to a maximum. A further increase in the angle of attack lead to a precipitous drop in the lift.
Lift occurs when a fluid is deflected by a moving aerofoil. It doesn't matter if the object is stationary and the fluid is moving (as with the experiment), or if the fluid is still and the object is moving through it (as with a soaring jet on a windless day). What really matters is the relative difference in speeds between the object and the fluid. The aerofoil did split the airflow in two directions: up and over the wing and down along the underside of the wing. The shape of the wing was asymmetric that it made the air moving over it travel faster than the air underneath. As the air speeded up over the aerofoil, its pressure dropped. So the faster moving air moving over the wing exerted less pressure on it than the slower air moving underneath the wing. This resulted in an upward force called lift. This was the reason behind the aerofoil exerting some lift (figure 1) at zero angle of attack. When the angle of attack of the aerofoil was increased, the airflow encountered an obstacle (in the form of change in wing angle), its path narrowed and the flow speeded up (figure 2, figure 3 and figure 4) and hence there was a further increase in the lift. This explains the linear increase in the lift coefficient with the angle of attack.
From figure 1 one could understand that after reaching a maximum, the lift coefficient started to drop. This was due to the boundary layer separation. When the air was passed over the aerofoil in he wind tunnel, a boundary layer was formed around the aerofoil due to the viscous forces occurring in the layer of the fluid close to the aerofoil surface. As the angle of attack was increased, boundary layer separation occurred when the boundary layer travelled far enough against the adverse pressure gradient that the speed of the boundary layer relative to the object fell almost to zero. The fluid flow became detached from the surface of the aerofoil. During boundary layer separation the portion of the boundary layer closest to the leading edge reversed in flow direction. The shear stress was zero at the separation point between the forward and backward flow. The overall boundary layer thickened at the separation point and was then forced off the surface by the reverse flow at its bottom. This resulted in loss of lift and stall.
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From figure 1 it could also be seen that the maximum lift coefficient for the NACA 2415 aerofoil used in this experiment with slats was higher than the lift coefficient of the same aerofoil without slat. This was due to the fact that slat increases the stall points. The slat was deployed in front of the aerofoil. In addition to the primary airflow over the main aerofoil, there was a secondary airflow through the gap between the slat and the aerofoil leading edge. This secondary flow injected high momentum fluid into the boundary layer on the upper surface. This highly energized air energized the boundary layer and increased the lift by preventing the stall at higher angle of attack.
From figure 1, comparing the NACA reference data and the experimental data it could also be found that with the increase in the Reynolds number the angle of attack at which the stall occurs also increased. (i.e. the lift coefficient increased with the Reynolds number).
There were a lot of experimental uncertainties occurred during the experiment as one could understand it by looking at the differences in the experimental and reference data. The compression tube of the wind tunnel was so long as it would influence the speed of the air. Then the setting up of the aerofoil angle of attack was done manually using a protractor that needed to be adjusted on both side of the wind tunnel. The errors associated with this would certainly influence the outcome of the data.
Conclusion
The experiment displayed the fundamental aerofoil characteristics in a fluid. This experiment provided a very clear view of the effects created by a leading edge slat. The lift coefficient of the aerofoil was increased with the introduction of the leading edge slat. The maximum lift coefficient of the NACA 2415 aerofoil with the slat was found to be 1.459 and the maximum lift coefficient without the slat was found to be 1.226. The experiment also showed that the pressure gradient was high on the leading edge and started decreasing towards the trailing end. It also proved that the lift coefficient increases with the increase in the angle of attack and with the increase in the Reynolds number of the flow.
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Appendix 1: Boeing 747 questions
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Appendix 2: Data
Angle
CL Without Slat
-10 -9 -8 -7 -6 -5 -4 -3 -2 -1 0 1 2 3
-0.855 -0.802 -0.731 -0.712 -0.593 -0.511 -0.327 -0.114 -0.003 0.08 0.153 0.275 0.482 0.59
4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20
0.655 0.733 0.824 0.856 0.947 1.061 1.145 1.154 1.164 1.226 1.161 1.112 0.916 0.969 0.741 0.743 0.789
CL with Slat
0.954 1.091 1.086 1.276 1.348 1.43 1.459 1.229 1.254 1.029 1.207
Table 1: experimental C L data for the NACA 2415 aerofoil with and without slat
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Angle (α) -18 -17 -16 -14 -12 -10 -9 -8 -7 -6 -5 -4 -3 -2 -1 0 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 22 24
NACA CL (Re-3*10^6)
NACA CL (Re-6*10^6)
NACA CL (Re-9*10^6)
-0.825
-0.9 -1.15 -1.35 -1.25 -1.05 -0.825
-0.875
-0.625
-0.625
-0.675
-0.4
-0.4
-0.45
-0.225
-0.225
-0.225
0
0
0
0.2
0.2
0.225
0.4
0.4
0.425
0.625
0.625
0.625
0.8
0.8
0.85
1
1.025
1.075
1.2
1.2
1.275
1.3
1.4
1.425
1.425
1.5
1.57
1.3
1.6
1.65
1.175
1.3
1.575
1.075 1.025 1.05
1.125 1.075 1
1.35 1.25 1.325
Table 2: NACA data – lift coefficients for different Re numbers at various angles of attack
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References 1. Sangam, CM; Lock, GD: Laboratory Handout Year 2 MEng – Aerofoil Experiment, Department of Mechanical Engineering, University of Bath.
2. National Advisory Committee for Aeronautics: Aerodynamic Characteristics of Aerofoils. NACA Rep. 93, 1920.
3. Hansen, James R.: Engineer in Charge. NASA SP-4305, 1987.
4. Munk, Max M.; and Miller, Elton W.: Model Tests with a Systematic Series of 27 Wing Sections at Full Reynolds Number. NACA Rep. 221, 1925.
5. Jacobs, Eastman N.; Ward, Kenneth E.; and Pinkerton, Robert M.: The Characteristics of 78 Related Airfoil Sections from Tests in the Variable-Density Wind Tunnel. NACA Rep. 460, 1933. 6. Jacobs, Eastman N.; and Pinkerton, Robert M.: Tests in the Variable-Density Wind Tunnel of Related Airfoils Having the Maximum Camber Unusually far Forward. NACA Rep. 537, 1935.
7. Jacobs, Eastman N.; Pinkerton, Robert M.; and Greenberg, Harry: Tests of Related Forward-Camber Airfoils in the Variable-Density Wind Tunnel. NACA Rep. 610, 1937.
8. Theodorsen, Theodore: Theory of Wing Sections of Arbitrary Shape. NACA Rep. 411, 1932.
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