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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*106.
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 Variable-
Density 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 thepressure 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 2
S/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 Cp and CL
data were collected.
Results
From the data collected (shown in table 1) during the experiment, CL 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).
Figure 1: CL versus for the NACA 2415 aerofoil with and without slat. NACA
reference data included.
-1.5
-1
-0.5
0
0.5
1
1.5
2
-30 -20 -10 0 10 20 30
CL without slat
CL with slat
NACA CL (Re-3*10^6)
NACA CL (Re-6*10^6)
NACA CL (Re-9*10^6)
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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 Reynoldsnumber).
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 CL with Slat
-10 -0.855
-9 -0.802
-8 -0.731
-7 -0.712
-6 -0.593
-5 -0.511
-4 -0.327
-3 -0.114
-2 -0.003
-1 0.08
0 0.1531 0.275
2 0.482
3 0.59
4 0.655
5 0.733
6 0.824
7 0.856
8 0.947
9 1.061
10 1.145 0.954
11 1.154 1.091
12 1.164 1.086
13 1.226 1.276
14 1.161 1.348
15 1.112 1.43
16 0.916 1.459
17 0.969 1.229
18 0.741 1.254
19 0.743 1.029
20 0.789 1.207
Table 1: experimental CL data for the NACA 2415 aerofoil with and without slat
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Angle
(α) NACA CL (Re-3*10^6) NACA CL (Re-6*10^6) NACA CL (Re-9*10^6)
-18 -0.9
-17 -1.15
-16 -1.35
-14 -1.25
-12 -1.05
-10 -0.825 -0.825 -0.875
-9
-8 -0.625 -0.625 -0.675
-7-6 -0.4 -0.4 -0.45
-5
-4 -0.225 -0.225 -0.225
-3
-2 0 0 0
-1
0 0.2 0.2 0.225
1
2 0.4 0.4 0.425
3
4 0.625 0.625 0.625
5
6 0.8 0.8 0.85
7
8 1 1.025 1.075
9
10 1.2 1.2 1.275
11
12 1.3 1.4 1.425
13
14 1.425 1.5 1.57
15
16 1.3 1.6 1.65
17
18 1.175 1.3 1.575
19
20 1.075 1.125 1.35
22 1.025 1.075 1.25
24 1.05 1 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.