aersp 305w aerospace technology...

AERSP 305W Aerospace Technology Laboratory

Laboratory Section 17

Laboratory Experiment Number 3 2D Airfoil Testing of S805 Airfoil

February 10, 2011 Performed in Room 8 Hammond Building

Brett Davis

Lab Partner’s Names: Titos Gosalvez Kevin Harrigan James Trexler

Zachary Watson

Lab TA: Brian Wallace

Course Instructor: Richard Auhl

Brett Davis Lab 3 of 14

Abstract

The objective of this experiment was to determine the 2D lift and drag coefficients of an S805

airfoil. Pressure taps in the airfoil were used to measure the local pressures on the airfoil’s

surface. These were used to calculate the lift coefficients. A hot-wire anemometer was used

to take a velocity survey of the airfoil’s wake in order to calculate the drag coefficients. The

coefficients were calculated from -18o to 18o at every two degrees. The angle of attach at

which stall and that at which the flow reattaches were also determined. The data from this

experiment coincided with the data presented in the NREL report on the S805 airfoil. This

experiment successfully determined the 2D lift and drag coefficients for the S805 airfoil and

was consistent with the NREL report.


Introduction

The objective of this experiment was to determine the 2D lift and drag coefficients of an S805

airfoil. The data obtained in this experiment was compared to prior experiments performed

by the National Renewable Energy Laboratory (NREL). With pressure readings from the

pressure taps in the airfoil, the local pressure coefficients were determined using Equations

(1a, 1b) for the upper and lower surface of the airfoil.

The normal force coefficient was calculated using the equations below.

(2)

(3)

Equation (3) was obtained by assuming a thin airfoil. From the normal force coefficient, the

2D lift coefficient was obtained with Equation (4) by assuming a small angle of attack.

(4)

To determine the 2D drag coefficient, Equation (5) was evaluated using Simpson’s Rule,

Equation (6), and the velocities measured by a hot-wire anemometer

.

(5)

(6)

During Lab 1, the wind tunnel’s venturi was calibrated using Equation (7). This was used to

hold the wind tunnel at a constant Reynolds number.

(7)


Experimental Procedure

To determine the 2D lift and drag coefficients for the S805 airfoil a series of experiments were

run in the 2x3 foot wind at a Reynolds number of 1,000,000 and at angles of attach ranging

from negative 18o to positive 18o. Time constraints only allowed each lab group to run tests

at two angles of attack. The data collected was shared between all lab groups. This report

concerns the experiments performed at 1o and -18o angle of attack. The details of the

procedures are outlined below.

A solid aluminum airfoil with a constant chord of 19.685” and a span of 24” was used to

perform these experiments. The airfoil had 48 static pressure ports on the upper surface and

49 ports on the lower surface. 17 of the upper and 17 of the lower ports were connected to a

manometer bank as seen in Table 1.

Table 1. S805 Pressure Tap Locations

X in. Y in. X in. Y in.

1 0.00 0.00 19

2 0.16 0.27 20 0.04 -0.08

3 0.39 0.46 21 0.24 -0.22

4 0.63 0.59 22 0.47 -0.30

5 1.26 0.86 23 0.79 -0.37

6 3.94 1.48 24 1.97 -0.55

7 6.89 1.76 25 4.92 -0.79

8 8.86 1.76 26 7.87 -0.88

9 9.45 1.73 27 9.84 -0.87

10 10.04 1.68 28 12.80 -0.77

11 10.63 1.61 29 14.17 -0.65

12 1.22 1.52 30 14.76 -0.58

13 11.81 1.42 31 15.35 -0.50

14 13.78 1.09 32 15.94 -0.41

15 16.73 0.59 33 16.34 -0.34

16 18.11 0.33 34 16.93 -0.25

17 19.29 0.07 35 17.72 -0.13

18 36 19.29 -0.01Atmosphere

Manometer

Port Number

Manometer

Port Number

Tap Location

Upper Surface Lower Surface

Tap Location

T.S. Static

Pressure transducer A was connected to the pitot static probe in the wind tunnel test section,

and transducer B measured the static pressure drop over the venturi. Transducers A and B

were set at a span of 4.3 and previously calibrated to 2.945 PSF/volt and 2.985 PSF/volt


respectively. Downstream from the airfoil a hot-wire anemometer was mounted on a sting

and motorized traverse system to measure the wake velocities as seen in Figure 1.

Figure 1. Experimental set up, location of hot-wire probe and airfoil

The hot-wire anemometer was previously calibrated to yield Equation (8).

(8)

To run the wind tunnel at a constant Reynolds number of 1 million, the venturi pressure drop

was held constant at 3.47 volts on transducer B. This pressure was determined from the

venturi calibration performed in Lab 1. The calibration constants in Lab 1 were determined

for the front, middle and rear of the test section as seen in the Appendix A. This experiment

was performed between the rear and middle locations, so these values were averaged to get

K = 0.899.

The airfoil was set at -17.95o (as close to -18o as possible) angle of attack and LabVIEW was

used to traverse the hot-wire probe vertically across the test section to measure the velocities

in the airfoil’s wake at every 0.125” over the 24” wind tunnel height. The probe made 2000

measurements at each location with a sample rate of 2 kHz. With the manometer bank tilted

at 19.8o the height of the fluid in each bank was recorded. When the hot-wire wake survey

was completed, the airfoil was rotated to 0.96o (approximately 1o). The hot-wire survey was

run again, and the fluid levels in each bank were recorded while the manometer bank was

tilted at 19.7o.

Hot-wire probe

S805 Airfoil

Traverse

direction


General observations about the stall point of the airfoil were made with the wind tunnel

running at 50%. The airfoil was rotated from -18o to stall while observing the manometer

bank. After the airfoil stalled at 20o, the airfoil was rotated back through lower angles of

attack until the flow reattached at 17.6o.

Results and Discussion

The data collected during this experiment and during the experiments performed by other lab

groups was combined to determine the 2D coefficients of lift and drag for the S805 airfoil.

The results of these experiments were compared to the data in the NREL Design and

Experimental Results for the S805 Airfoil report to assess the findings of this experiment.

During the lab, the room temperature, pressure and air density were recorded in Table 2.

Table 2. Room conditions during experiments

During Experiment Temperature (F) Pressure (in Hg) Density (slug/ft^3)

α = 1 deg 66.4 29.42 0.002305

α = -18 deg 67.1 29.42 0.002302

A time trace of the velocity at the start of surveys at 1 and negative 18 degrees was taken at

the top of the test section. As seen in Figure 2, the turbulence intensity was much larger

when the airfoil was at negative 18 degrees angle of attack than at positive 1 degree. These

results would be expected, as the larger angle of attack should produce more turbulence

downstream from the airfoil.

Figure 2. Velocity time trace downstream from S805 airfoil

70.00

80.00

90.00

100.00

110.00

120.00

130.00

0.00 0.20 0.40 0.60 0.80 1.00

Ve

loci

ty (

ft/s

)

Time (sec)

1 Deg

Re = 1x10^6 Ti (1o) = 0.003464

Ti (18o) = 0.016203


From the hot-wire anemometer’s measurements, the average velocity at each position was

calculated. These velocities were plotted as a function of vertical position to form the wake

profile at 1o in Figure 3 and at -18o in Figure 4. These profiles show that the -18o angle of

attach created a much larger velocity deficit than the 1o orientation. This large wake results in

a larger coefficient of drag, 0.2412, than the smaller wake at 1o, which had a cd of 0.0083.

These differences are consistent with what is expected.

Figure 3. Wake velocity profile at 1o angle of attack

Figure 4. Wake velocity profile at -18o angle of attack

0

5

10

15

20

88 90 92 94 96 98 100

Ve

rtic

al P

osi

tio

n (

inch

es)

Velocity (ft/s)

Re = 1x10^6 α = 0.96o

cd = 0.0083

0

5

10

15

20

0 20 40 60 80 100 120

Ve

rtic

al P

osi

tio

n (

inch

es)

Velocity (ft/s)

Re = 1x10^6 α = -17.95o

cd = 0.2412


Using the pressure readings taken from the manometer bank, the pressure distribution over

the airfoil was plotted. In Figure 5, it can be seen that at α = 1o the upper surface of the airfoil

experienced negative pressures and the lower surface had positive pressure, indicating

positive lift resulting in a cl of 0.3344. The data from this experiment closely follows the data

from the NREL report as can be seen from Figure 5. There was a favorable pressure

gradient to about 50% of the chord on the upper surface and 65% on the lower surface,

followed by adverse pressure gradients, which is close to the design specifications in the

NREL report.

Figure 5. Pressure distribution over S805 airfoil at 1o angle of attack

The pressure distribution over the airfoil at -18o can be seen in Figure 6. From this graph, it

can be seen that the airfoil generated negative lift, as the upper surface experienced positive

pressures, and the lower surface had negative pressures. There was not a large pressure

gradient on the lower surface in this orientation, while the upper surface had a large adverse

pressure gradient up to 0.5% of the chord followed by a mostly favorable pressure gradient.

At this angle, the cl was -0.6714. This data did not follow the NREL data as closely, although

-0.8

-0.6

-0.4

-0.2

0

0.2

0.4

0.6

0.8

1

0.000 0.200 0.400 0.600 0.800 1.000

Cp

x/c

Experimental Upper Surface

Experimental Lower Surface

NREL upper

NREL lower

Re = 1x10^6 cl = 0.3344

αExperiment = 0.96o

αNREL = 1o

LE TE


the general trend was consistent. The angles of attack did not exactly match the NREL data

in this case which could have contributed to the differences in the data.

Figure 6. Pressure distribution over S805 airfoil at -18o angle of attack

The compiled data from each lab group was used to generate a plot of the coefficient of lift as

a function of the angle of attack in Figure 7. The linear portion of the cl versus α curve was a

near perfect match to the NREL results. In the non-linear regions of the curve, the

coefficients of lift from this experiment were slightly greater than those from the NREL report

for positive angles of attack. At -16o and -14o the coefficients of lift are slightly less than

expected possibly due to measurement errors. Overall, the cl versus α curve is consistent

with expectations and the NREL findings.

-1.500

-1.000

-0.500

0.000

0.500

1.000

1.500

0.000 0.200 0.400 0.600 0.800 1.000

Cp

x/c

Experimental Upper Surface

Experimental Lower Surface

NREL lower

NREL upper

Re = 1x10^6 cl = -0.6714

αExperiment = -17.95o

αNREL = -18.1o

TE LE


Figure 7. Coefficient of lift vs. angle of attack of S805 airfoil

The compiled data was also used to create a plot of the coefficient of lift versus the coefficient

of drag in Figure 8. The plot shows the expected relationship between cl and cd, and the

NREL data appears to be consistent with the findings of this experiment. However, the NREL

data was over a much smaller range of drag coefficients. Figure 9 shows the data over the

range of cd presented in the NREL report. At small values of cd this experiment’s data is

relatively close to the NREL data. The differences between the two sets of data are likely

due to the difficulty in determining these very small drag coefficients. In this experiment, the

cd was evaluated by integrating the wake profile using Simpson’s Rule. The bounds of

integration for each angle of attack had to be manually selected because the velocity varied

slightly over the height of the wind tunnel. The approximations in this approach could be

responsible for the differences seen in Figure 9.

-1

-0.5

0

0.5

1

1.5

-20.00 -15.00 -10.00 -5.00 0.00 5.00 10.00 15.00 20.00

Cl

Angle of Attack (deg)

Experimental

NREL

Re = 1x10^6


Figure 8. Coefficient of lift vs. coefficient of drag of S805 airfoil

Figure 9. cl vs. cd*1000 of S805 airfoil at small cd values

-1

-0.5

0

0.5

1

1.5

0 0.05 0.1 0.15 0.2 0.25

cl

cd

Experimental

NREL

Re = 1x10^6

-0.4

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

0 5 10 15 20

cl

cd *1000

Experimental

NREL

Re = 1x10^6


With the tunnel running at 50%, the airfoil stalled at approximately 20o, however, the

protractor maxed out at 20o so this measurement could be slightly off. Figure 10 shows the

manometer bank when the airfoil was at 6o to 20o angle of attack.

Figure 10. Manometer bank showing airfoil pressures from 6o until stall

The flow reattached when the airfoil was rotated back to 17.6o. The flow separated at a

higher angle of attack than it reattached at because the flow needs more energy to reattach.

Once the flow is attached, its momentum resists separation until the adverse pressure

gradient over comes it. Once separated, the flow needs to gain enough energy to initially

reattach which is why the two angles were different.

6.0o 11.2

o 13.5

o 18.9

o ~20

o

(stall)


Conclusions

The 2D lift and drag coefficients for the S805 airfoil found from this experiment were

consistent with those presented in the NREL report. The coefficient of lift as a function of the

angle of attack yielded the same relationship as that in the NREL report with only a slight

error at large angles of attack. The lift and drag coefficients’ relationship was also consistent

with the NREL data. Some errors occurred at small coefficients of drag likely due to the

difficulty in computing such small amounts of drag. These differences were insignificant

when considered over the full range of data taken. General observations indicated that the

angle of stall is greater than the angle required to reattach the flow, which is consistent with

what was expected. Overall, the experiment successfully determined the 2D lift and drag

coefficients for the S805 airfoil.


The Appendix

Appendix A. 2x3 ft. wind tunnel venturi calibration

∆P = 0.9976q - 0.0937

∆P = 0.8976q + 0.1094

∆P = 0.8766q + 0.0078

∆P = 0.9897q

0

5

10

15

20

25

30

0 5 10 15 20 25 30

Co

ntr

acti

on

Pre

ssu

re D

rop

(P

SF)

Test Section Dynamic Pressure (PSF)

Front

Middle

Rear

Theoretical

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