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Aerofoil Experiment

Pressure distribution over a NACA 2415 aerofoil

Elankumaran Nagarajan

20th October 2013

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.

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

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

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)

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

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.

-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)

Figure 2: pressure arrow diagram for NACA 2415 aerofoil at = 2

Figure 3: pressure arrow diagram for NACA 2415 aerofoil at = 8

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)

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.

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.

Appendix 1: Boeing 747 questions

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

(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

-8 -0.625 -0.625 -0.675

-7-6 -0.4 -0.4 -0.45

-4 -0.225 -0.225 -0.225

-2 0 0 0

0 0.2 0.2 0.225

2 0.4 0.4 0.425

4 0.625 0.625 0.625

6 0.8 0.8 0.85

8 1 1.025 1.075

10 1.2 1.2 1.275

12 1.3 1.4 1.425

14 1.425 1.5 1.57

16 1.3 1.6 1.65

18 1.175 1.3 1.575

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

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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