04/12/02
57:020 Mechanics of Fluids and Transfer Processes
Laboratory Experiment #3

Measurement of Pressure Distribution and Lift for an Airfoil
Purpose
The objectives of the experiment are to examine the surface pressure distribution and to compute the lift force acting on the airfoil.

Test Design
A body immersed in a flowing fluid is exposed to both pressure and viscous forces. The sum of the forces that acts normal to the free-stream direction is the lift, and the sum that acts parallel to the free-stream direction is the drag. The geometric and dynamic characteristics of airfoils are shown in Figure 1. This experiment is concerned with computation of the lift on a stationary airfoil mounted in the test section of a wind tunnel. We will consider only two-dimensional airfoils where tip and root effects are neglected.
Because the velocity of the flow over the top of the airfoil is greater than the free-stream velocity, the pressure over the top is negative. This follows directly from the application of Bernoulli’s equation. Similarly the velocity along the underside of the airfoil is less than the free-stream velocity and the pressure there is positive.
Hence, both the negative pressure over the top and the positive pressure along the bottom contribute to the lift.
There are a variety of ways to measure lift. In this experiment, the lift force, L, on the airfoil will be determined by integration of the measured pressure distribution over the airfoil’s surface. Typical pressure distribution on an airfoil and its projection on the airfoil normal are shown in Figure 2.
The pressure distribution on the airfoil is expressed in dimensionless form by the pressure coefficient Cp

Cp =

pi − p ∞
1
2 ρU ∞
2

(1)

where pi is the surface pressure measured at location i on the surface, p∞ is the pressure in the free stream, ρ is air density, and U∞ is the free-stream velocity given by

U∞ =

2( p stagnation − p ∞ )

ρ

Figure 1. Geometric and dynamic parameters of airfoils U∞

(2)

where pstagnation is the stagnation pressure measured at the tip of the Pitot tube.

1

Figure 2. Pressure distribution on an airfoil

The lift force is customarily expressed as a dimensionless lift coefficient per unit span length

CL =

2L
2
ρU ∞ bc

(3)

where L is lift force on the airfoil obtained by integration of the measured pressure distribution over the airfoil’s surface, b is the airfoil span, and c is the airfoil chord, as shown in Figure.
In this experiment, a Clark-Y airfoil is the shape under consideration. Airfoil characteristics are provided in Appendix A. Measurements will be made using an automated data acquisition system (ADAS) as sketched in
Figure 3.

Software
- Surface
Pressure
- Velocity
- WT Control

Digital i/o PC

A/D
Boards

Serial
Comm.
(COM1)

Metrabyte
M2521
Signal
Conditioner

Scanivalve
Position
Circuit (SPC)

Scanivalve
Controller
(SC)

RTD

Pressure
Input
Digital
Voltimeter
(DVM)

Scanivalve

Pitot Tube
(Free
Stream)

Scanivalve
Signal
Conditioner
(SSC)

Pressure
Transducer
(Validyne)

Pressure Taps

Airfoil Model
Bundle of tubes Figure 3. Schematic of the data acquisition system
Substituting in equation (3) for the measured variables and using the notations provided in Appendix A, the data reduction equation for the lift coefficient is

³ (p

i

CL =

s

− p∞ ) sin (θ )ds
2
1 / 2 ρU ∞ c

where θ is the angle of surface normal to free-stream flow.

2

(4)

Measurement Systems and Procedures
The equipment (measurement systems) used in this experiment includes:
• Protractor - angle of attack
• Measuring tape - chord length
• Resistance temperature detectors (RTD)/electronic data acquisition - temperature in the tunnel
• Pitot-static probe/scanning valve/pressure transducer/electronic data acquisition - free-stream velocity
• Pressure taps/scanning valve/ pressure transducer/electronic data acquisition - surface pressure
Pressure and temperature are measured using the ADAS.
The pressure differences pi − p∞ and pstagnation − p∞ needed in equations (2) and (3) are measured using a pressure transducer, i.e., a device which converts a pressure difference into a voltage. The pressure transducer has one of the ports connected at all times to p∞ , the reference pressure in the free stream. The other port is connected to the output of an analog multiplexer
(scanning valve), which measures pressure at various locations in test section. The scanning valve is a rotary multiple-port switch that permits multiple pressure inputs to be connected to a common output. In the present experiment the input connections to the scanning valve are as follows:
• Ports 0 to 28 are connected to the corresponding pressure taps located on the airfoil (see Figure 3)
• Port 40 is connected to the Pitot-static tube to measure the stagnation pressure
• Port 42 is connected to the Pitot-static tube to measure the static pressure (not used in experiments).
Prior to the experiments, the pressure transducer is calibrated against a Rouse manometer (precision micromanometer with a resolution of 0.001 inches). The calibration (conversion) equation valid for the entire range of measured pressures is given by h = aV + b
(5)
where h is the reading of the Rouse manometer (in inches of liquid column), V is the voltage produced by the pressure transducer, a and b are coefficients obtained through linear regression applied to the calibration points.
Figure 4 illustrates the block diagram for the measurement systems and data reduction equations for determination of the lift coefficient.

EXPERIMENTAL ERROR SOURCES

ANGLE OF
ATTACK

TEMPERATURE

FREE-STREAM
VELOCITY

SURFACE
PRESSURE

INDIVIDUAL
MEASUREMENT
SYSTEMS

Xα
B α, P α

XT
B T, P T

XV
BV,PV

Xp
B p,P p

MEASUREMENT
OF INDIVIDUAL
VARIABLES

ρair = ρ (XT ) = f (Tair) air air

(p - p ) sin(θ)ds
CL = CL(Xα , XT, XV, Xp, XG) =

CL
BC , P C
L

s

DATA REDUCTION
EQUATIONS

1/2 ρU c
2

EXPERIMENTAL
RESULTS

L

Figure 4. Block diagram of the lift coefficient determination including measurements systems, data reduction equations, and results

3

Each group of students will measure the pressure distribution on the airfoil set at an angle of attack in a pre-established value for the free-stream velocity. Each group will repeat the measurements ten times to provide information for estimation of the precision errors.
The following operations will be made during the experiment:
(1) Follow teaching assistant (TA) instructions for setting wind tunnel velocity.
(2) Follow teaching assistant (TA) instructions for powering on the wind tunnel and ADAS.
(3) Open the Fluids lab folder located on the ADAS computer monitor. Double click on the icon Tunnel Velocity
– Temperature_OTWT_MLB.vi (see Figure 5). This program is a Virtual Instrument (VI) created using
LabView for automatic display of the operating parameters of the wind tunnel. Enter the values of the atmospheric pressure and ambient temperature. These values can be obtained using the barometer-temperature device installed inside the large wind tunnel. Run the VI by clicking on the arrow button to the top left of the screen. The VI displays the velocity and temperature in the test section of the tunnel. Next select
Operate/Make Current Values Default from the Menu toolbar to transmit this information to all the running
VIs.

(4) In the fluids lab folder you will also find the Lift VI. Double click on the icon labeled Lift. This Virtual
Instrument (VI) allows automatic acquisition of the data required for calculating the lift coefficient. The screen of the VI is shown in Figure 6. Enter in the Observation Point List box the scanning valve port numbers where you would like to take measurements (see above the labeling for the input connections to the scanning valve).
Use commas or spaces between the port numbers in the Observation Point List box. Follow TAs instructions for appropriately selecting the Sampling Rate (Hz), Settling Time (s), and Length of each Sample (s) and enter these values in the corresponding boxes. Type in the conversion (regression) coefficients a and b (usually displayed on the computer cabinet or provided by TAs) the density of the Rouse manometer (provided by
TAs), and the angle of attack at which the model is set.

4

Figure 6. Computer interface for the lift measurements VI
(5) Run the VI by clicking on the arrow button to the top left of the screen. A window will appear asking about the directory where the data file will be saved. Select the file name and path of the directory to save the data.
(6) Repeat steps 4 and 5 for the remaining measurements.

Uncertainty Analysis
Using your own data, derive and present results for uncertainties estimates for pressure coefficient and lift coefficient. The methodology for estimating uncertainties follows the procedures presented in IIHR Report No. 406
(Stern et al., 1999) for Multiple Tests method (sections 6.1 and 6.3). Based on previous experiments, it was found that bias and precision limits for the ρ, U ∞ , θ, s, and c are negligible, hence for the present analysis only include the bias limit for pi − p∞ . Table 5 provides bias limit estimates for the individual variables. The data reduction equations for the pressure and lift coefficients are equations (1) and (4), respectively.
Table 1. Assessment of elemental errors for the individual variables included in the data reduction equations
Variable
Bias Limit
2.1
pi-p∞
(Pa)
0
(kg/m3)
ρ
0
U∞
(m/s)
0 θ = (β -α) (degrees)
0
s
(m)
0 c (m)

Estimation
Precision Limit
Previous meas.
Determined
Negligible
0
Negligible
0
Negligible
0
Negligible
0
Negligible
0

5

Estimation
Repeated measurements
No repeated meas.
No repeated meas.
No repeated meas.
No repeated meas.
No repeated meas.

Pressure Coefficient
The data reduction equation for the pressure coefficients (equation 1) is of the form

C p = f ( pi − p ∞ , ρ ,U ∞ )
However, here we will only consider bias limits for

pi − p∞ (see Table 1).

The total uncertainty for the pressure coefficient measured at each pressure tap is given by equation (24) in Stern et al. (1999)
2
2
2
U Cp = BCp + PCp

(6)

Bias Limits. The bias limit for each of the taps (constant for all) in equation (6) is given by equation (14) in Stern et al. (1999). Neglecting the correlated bias limits and discarding the terms with negligible bias limit j 2
BCp = ¦θi2 Bi2 = θ (2pi − p∞ ) B(2pi − p∞ )

(7)

i =1

θ i = ∂C p ∂X i

where Bi is the bias limit for the individual variable and

are sensitivity coefficients.

The

sensitivity coefficient θ ( pi − p∞ ) , evaluated using the average values for the individual variables, is given by

θ(p _ p ) = i ∞

∂C p

∂( pi − p∞ )

=

2
2
ρU ∞

Precision limits. The precision limit for each pressure tap in equation (6) is estimated using equation (23) in Stern et al. (1999) with t = 2

PCp = 2S Cp

M

(8)

where SCp is the standard deviation of the pressure coefficients at each pressure tap evaluated using equation (22) in
Stern et al. (1999) for the M = 10 repeated tests. Precision limits will have different values for each of the 28 taps.
Lift Coefficient
The data reduction equation for the lift coefficient (equation 4) is of the form

C p = f ( pi − p∞ , θ , s , ρ , U ∞ , c )
However, here we will only consider bias limits for

pi − p∞ (see Table 1).

The total uncertainty for the measurement of the lift coefficient is given by (equation 24 in Stern et al.,
1999)
2
2
2
U CL = BCL + PCL

(9)

Bias Limits. The bias limit in equation (9) is given by equation (14) in Stern et al. (1999). Neglecting the correlated bias limits and the terms assumed with negligible bias limits j 2
BCL = ¦θ i2 Bi2 = θ (2pi − p∞ ) B(2pi − p∞ )

(10)

i =1

Given the above assumptions and replacing the integral in the data reduction equation (4) with summation of elemental lift coefficients over the surface of the airfoil, the expression for the bias limit of the lift coefficient is
2
CL

B

=B

2
( pi − p∞ )

2

ª ∂C
º ª B p º
¦ « ∂( p −L p )» =« 0.5( pi−U∞2)c » i =1 ¬
∞ ¼ i ¬ ρ ∞ ¼ k 2

k

¦ [sin(θ )ds ] i =1

Notations used in equation (11) are defined in Table 1 and Appendix A (k = 29).

6

2

i

i

(11)

Precision limits. The precision limit in equation (10) is estimated using equation (23) in Stern et al. (1999) with t = 2

PCL = 2S CL

M

(12)

where SCL is the standard deviation of the lift coefficients evaluated using equation (22) in Stern et al. (1999) for the
M = 10 repeated tests.

Data Analysis and Discussions
Each group of students will obtain pressure coefficients and lift coefficients for one free-stream velocity and an angle of attack. All data required for analysis is contained in a file outputted by the ADAS. The format of the outputted file is illustrated in Figure 7

1
2
3
4
5
6
7
8
9
10
11
12
13
14
15
16
17
18
19
20
21

A
B
C
D
‘Legend’
‘File Name’
‘ Angle (degrees), Temperature (°C), Manometer Fluid Density (kg/m3)’
‘Coefficients a,b, from h = a⋅V +b conversion equation
'Coefficients a1, b1 from h = a1 * V + b1 conversion equation
‘A/D channel, Sampling rate (Hz), Settling Time (s), Sampling time (s)’
‘=====================================’
D;\vhil\data\lift_041202_0911_15.txt
10
25.1
784.8
4.14E-01
-3.56E-03
1.01E+00
7.94E-01
0
1000
2
‘=====================================’
‘Data Section – Pressure Tap Number (Scanivalve Port), Mean Voltage (V)’
0
-0.0703
1
-0.66949
2
-0.78618
Etc.
Etc.
27
0.38178
28
0.5081
40
0.53052

E

F

G

2

Figure 7. Sample of data file outputted by ADAS
The values for the items/quantities listed in rows 2 to 6 are specified in the same order in rows 8 to 12.
Quantities listed from row 15 on are those measured by ADAS. Values in column A specify the port where the differential pressure is measured. Values in column B are voltages measured by the pressure transducer. Using the conversion equation (5) these values can be converted into inches of manometer liquid (i.e., alcohol ρ = 784.8kg/m3 at 26°C) and subsequently in pressures. Note that the density of the Rouse liquid (usually alcohol) varies with the ambient (not wind tunnel) temperature.
Data reduction will include the following steps:
1. Calculate the flow Reynolds number, Re = U∞⋅c/ν (using equation 3 and the reading on port 40).
2. Calculate the pressure coefficients (using equation 2 and the readings on ports 0 to 28) on the airfoil surface.
Plot Cp versus x/c, for the angles of attack employed in the experiment. Include as error bars on this plot the calculated total uncertainties. Compare your distribution with the reference data provided in Figure 8.a.

7

AOA = 4

AOA = 0

20

40

60

80

100

0

20

Cp

x/c
-2.0
-1.5
-1.0
-0.5
0.0
0.5
1.0
1.5
2.0

60

20

40

60

80

AOA = 8

100

0

20

x/c
-3.0
-2.5
-2.0
-1.5
-1.0
-0.5
0.0
0.5
1.0
1.5
2.0
2.5
3.0

40

60

20

40

100

80

-2.0
-1.5
-1.0
-0.5
0.0
0.5
1.0
1.5
2.0
100

x/c

AOA = 12

0

80

x/c

AOA = 6

0

Cp

40

Cp

0

-2.0
-1.5
-1.0
-0.5
0.0
0.5
1.0
1.5
2.0

Cp

-2.0
-1.5
-1.0
-0.5
0.0
0.5
1.0
1.5
2.0

60

80

100

0

20

40

x/c

-2.0
-1.5
-1.0
-0.5
0.0
0.5
1.0
A0A = 14
1.5
2.0
60 80 100

Cp

Cp

a)

x/c

b)

1.25

1.00

CL

0.75

0.50

0.25
Reference data (Re = 75,000)
Reference data (Re = 100,000)
Reference data (Re = 200,000)

0.00
0

5

10

15

20

25

Angle of attack (degree)

Figure 8. Reference data (Marchman and Werme, 1984); a) Distribution of the pressure coefficients for α = 0°,
4°, 6°, 8°, 12°, 14° and Re = 75,000; b) Variation of the lift coefficient with the angle of attack

8

3.

4.

1.
2.
3.
4.

Calculate the experimental lift force on the airfoil. The lift force is obtained by integrating the measured pressure over the airfoil surface. Use of the trapezoidal rule for the numerical integration is recommended (see
Appendix A).
Calculate the lift coefficient per unit span length, CL, given by equation (4). Include the uncertainty estimate for the lift coefficient. Compare your value for the lift coefficient with the value determined from the reference data provided in Figure 8.b.
Answer the following questions:
Suggest another experimental procedure to measure the lift force on the airfoil.
If the lift L is a function of the free-stream velocity U, density ρ, chord c, angle of attack α, and viscosity µ, what are the dimensionless groups (π parameters) that characterize this problem?
How do the experimental measurements of the pressure distribution apply to a full-scale aircraft having a similar airfoil section?
How do the experimental measurements of the pressure distribution apply to a full-scale aircraft having a similar airfoil section?

References
AIAA (1995). AIAA- 071 Standard, American Institute of Aeronautics and Astronautics, Washington, DC.
Granger, R.A. (1988). Experiments in Fluid Mechanics, Holt, Rinehart and Winston, Inc. New York, N.Y.
Marchman III, J.F. and Werme, T.D. (1984). “Clark-Y Performance at Low Reynolds Numbers,” Proceedings
AIAA 22nd Aerospace Science Meeting, Reno, NE.
Robertson, J.A. and Crowe, C.T. (1993). Engineering Fluid Mechanics, 5th edition, Houghton Mifflin, Boston,
MA.
Stern, F., Muste, M., Beninati, L-M., Eichinger, B. (1999). “Summary of Experimental Uncertainty
Assessment Methodology with Example,” IIHR Report, Iowa Institute of Hydraulic Research, The
University of Iowa, Iowa City, IA.
White, F.M. (1994). Fluid Mechanics, 3rd edition, McGraw-Hill, Inc., New York, N.Y.

9

APPENDIX A
Geometry of the Clark-Y Airfoil and Positioning of the Pressure Taps on the Airfoil Surface

∆s

β

Node

Surface

Surface

Number (i)

Element

(i-1) to (i)

Length

to Chord Line

(mm)

Angle

N
Tangent to surface at P

P

8

U

β

x/C

Camber Line

(degrees)
Airfoil chord

θ

6

5
3 4

2
1

7

α (angle of attack) 8

9

10

11

12

13

14

0,29
28

24
27 26
25

0

23

22

0.2

21

20

0.4

18

19

x/c

0.6

17

16

15

0.8

1.0

Figure 1A. Positioning of the pressure taps on the airfoil surface

Airfoil Geometry:
- maximum thickness, t = 0.0254 m;
- airfoil wing span, b = 0.762 m;
- chord length, c = 0.3048 m
Specific notations:
- surface to chord line angle, β;
- angle of surface normal to free-stream flow, θ;
- angle of attack, α.

0
1
2
3
4
5
6
7
8
9
10
11
12
13
14
Trailing Edge
15
16
17
18
19
20
21
22
23
24
25
26
27
28
29 or 0

Lift force can be computed using the trapezoidal scheme for the numerical integration

L = ³ ( p ∞ − p )sin(θ )ds ≈ ¦ s i

1
[( p∞ − pi ) + ( p∞ − pi −1 )]sin(θ i −1,i )∆si −1,i
2

10

0.00
1.25
2.50
5.00
7.50
10.00
15.00
20.00
30.00
40.00
50.00
60.00
70.00
80.00
90.00
100.00
90.00
80.00
70.00
60.00
50.00
40.00
30.00
20.00
15.00
10.00
7.50
5.00
2.50
1.25
0.00

6.0
6.0
9.0
9.0
8.0
15.5
15.5
30.0
30.0
30.0
31.5
31.5
31.5
31.5
30.0
29.0
31.0
31.0
31.0
31.0
30.0
30.0
30.0
15.5
15.5
7.5
9.0
6.0
4.5
8.0

160.7
136.2
115.8
107.3
102.6
100.1
97.1
93.2
89.8
86.7
83.9
81.9
81.1
79.7
78.6
272.4
272.4
272.4
272.4
272.4
272.4
272.4
272.4
272.4
272.4
272.4
265.5
253.4
241.0
209.0