WIND TUNNEL TESTING OF A COMPLETE AIRCRAFT

BACKGROUND: THE FIRST WIND TUNNELS

The first recorded mention of a wind tunnel can be found in a lecture by F. H. Wenham to the

Aeronautical Society of Great Britain, founded in 1866. In his lecture on April 17th

, 1867, Mr.

Wenham addressed the Society in the following terms:(1)

I propose shortly to try a series of experiments by the aid of an artificial current of air of known strength, and to place the Society in possession of the results.

On May of 1870, an experimental committee including three engineers and headed by

Wenham was formed, and shortly after the first wind tunnel expressly for aeronautical

purposes was designed and built. Its purpose was To ascern the fundamental relation between velocity and pressure on surfaces of different areas and inclinations. It was described in the Society's 1870 Annual Report as follows:

For the purpose, an instrument has been designed by Mr. F.H. Wenham and approved by the Experimental Committee, which is intended to submit

to the action of a fan no less than 30 in. in diameter, capable of delivering

about 3,000 cubic feet of air per minute. A clear space of 15 ft. or more in

front of the fan will allow room for a square wooden trunk, 10 ft. long and

18 inches square, to guide the blast, ascertain its velocity and insert the

anemometer.

The tunnel was equipped with a primitive balance capable of reading horizontal and

vertical forces simultaneously, and incorporating wooden windshields for the struts.

Wenham and a fellow engineer, Mr. Spencer Browning, published the first experimental

data the following year.

We found that as the angle of attack became more acute, the center of

pressure came nearer the front edge. (...) The resultant force on the

surface was found to be, as it should be, almost exactly normal to the

surface. The results on an inclined surface, down to an angle of 15

degrees, agree almost exactly with Hutton's rule, or still better, with a rule

we should now prefer, mainly on the ground that it was very simple, the

rule of Colonel Duchemin." (2)

(1)

J. L. Pritchard, "Francis Herbert Wenham, Honorary Members, 1824-1908: An Appreciation of the First Lecturer

to the Aeronautical Society." Journal of the Royal Aeronautical Society, Volume 62, August 1958, pp. 571-596.

(2)

If P is the pressure on the inclined surface, P' is the pressure on surface at right angles to the wind and the angle

of attack, the Duchemin's formula is P P

2

1 2

sin

sin

.

The use of aeronautical wind tunnels as a valuable experimental tool would become widespread

in subsequent years. Perhaps one of the most important wind tunnel researchers of all times

were Wilbur and Orville Wright.

(3)

Photographs from "The Wright Brothers, Heirs of Prometheus", Smithsonian Institution Press.

The Wright brothers were no mere tinkers.

After flying a number of unpowered gliders

in the summer of 1901 and finding the

results unsatisfactory, they decided to

improve the gliders performance. They undertook a serious aerodynamic research

program between September 1901 and

August 1902. Their rigorous wind tunnel

work enabled them to evaluate previous

airfoil tests compiled by other

experimenters. As a result they rejected

these earlier efforts as misleading and

inaccurate, and developed their own airfoil

sections with corresponding test data.

This pioneering wind tunnel research

placed them at the forefront of

contemporary flight researchers. The

successful flight of their Flyer on the

December 17th

1903 was the result of

this rigorous and meticulous research

effort and their mechanical genius.

Above, a replica of the original

Wright wind tunnel designed and

built by the two brothers in Dayton,

Ohio. The picture on the right shows

the small wind tunnel balance they

used to measure lift and drag on

various experimental airfoils.(3)

PARKS COLLEGE LOW-SPEED WIND TUNNEL

The Low-Speed tunnel, manufactured by Aerolab of Laurel (Maryland) is designed for infinitely

variable speeds in the range of zero to 220 ft/s, by means of adjustable shutters upstream of a

Westinghouse Centerline centrifugal fan. The general design of the tunnel is of an open circuit

(Fig. 1), closed test section type using the laboratory room for air return.

The tunnel test section is rectangular, 28 in high, 40 in wide and 54 in long. Reynolds Numbers

up to 1.26 x 106 per foot may be obtained in the test section. Air enters through a 6 x 8 foot

honeycomb flow straightener and two fine mesh (removable) turbulence screens, followed by a

6.2 to 1 contraction cone. This produces uniform flow in the test section, where turbulence

levels as low as 1.25 has been measured without the screens, and 1.06 with the screens.

A six-component pyramidal balance system supports the model in the test section (see Fig. 2).

Strain-gage load cells are used to measure all forces and moments. A Lab View-based data

acquisition program incorporates first-order balance corrections and provides the experimenter

with a real-time graphical interpretation of testing results. The angle of attack is automated over

a 25 range.

Summary of Tunnel Characteristics

Open Circuit System

Closed Jet Test Section

Speed Range: 0 220 ft/s (150 mph)

Power: 125 HP electric motor driving a constant-speed centrifugal blower.

Test Section Dimensions

Width: 40 inches

Height: 28 inches

Length: 54 inches

Six component pyramidal balance.

Angle of attack adjustment: 25

Angle of yaw adjustment: fu11 360

Computerized data acquisition system

Load Cell Ranges:

Lift: +200 to -50 lb Roll Mmt: +250 in-lbs to -250 in-lb

Drag: +100 to -100 lb Yaw Mmt: +250 in-lbs to -250 in- lb

Side Force: +100 to- 100 lb Pitch Mmt: +500 to -500 in- lb

Figure 2 A six-component pyramidal wind tunnel balance. Note the complexity of the balance

mechanism.

CURRENT EXPERIMENT

OBJECTIVE:

To familiarize the student with basic wind tunnel techniques by testing a complete model in the

Low-Speed Wind Tunnel. This introduction will provide the student with the basic skills for

successful testing in the Design II course. Special emphasis is placed on the different corrections

that must be applied to , CL and CD.

INTRODUCTION:

A number of corrections must be applied to raw wind tunnel data if one wishes to obtain

meaningful results. Chapter 9 in the course textbook (Ref. 1) outlines in detail these corrections,

and a brief look at this chapter should convince the student that this is no trivial matter. The

section numbers throughout the text refer to this book. Additional information about wind tunnel

corrections may be found in a comprehensive manner in Ref. 2. A summary of the most

important corrections for a closed jet wind tunnel (Fig. 3) is given in the following sections.

Fig. 3. - Open Return Wind Tunnel

Anatomy of an open return, closed jet wind tunnel, similar to Parks Low-Speed Tunnel.

Compared to closed return tunnels, these tunnels are noisier and require more power to

run, but are lower in cost. (Reprinted from Ref. 1)

1) Corrections Arising from the Calibration of the Test Section:

a. Velocity variation across the test section (Sec. 3.12): The velocity (and hence the dynamic pressure) may not be uniform throughout the test section. This variation can

be measured by a pitot-static tube and conveniently mapped.

b. Variation of flow direction in the test section (Sec. 3.14): The existing up flow and cross flow in the test section may be measured by any of the devices outlined in

Section 3.4 (i.e. yaw head, claw). For example, an up flow across the wing span

results in an effective twist being imparted to the wing, and a cross flow gradient in

the region of the vertical tail will change the slope of the yawing moment versus

sideslip or yaw angle.

c. Longitudinal static gradient (Sec. 3.13): All tunnels with closed test sections experience a decrease in static pressure from the front to the back end of the test

section. Thus, there is a tendency for the model to be drawn downstream (Horizontal

Buoyancy). Some tunnels have slightly expanding test sections to counteract this

effect.

2) Corrections Arising from the Balance Calibration (4.13):

a. Wind tunnel balances are extremely complex contraptions and must be properly calibrated in order to provide accurate readings. No balance exists that is capable of

perfectly measuring the loads it was intended to measure. In other words, when a

pure lift load is applied, not only the lift-load sensor will be affected, but also the

other five sensors as well (to a lesser degree). This is known as balance interactions,

and must be corrected by proper calibration.

b. There are two sources of errors in balances. The first arises from misalignment of the balance parts, caused by manufacturing tolerances and wear. These errors are linear

in nature. The other arises from the elastic deformation of the various parts, and are

of the second-degree order and nonlinear. The calibration of the balance involves

arriving at a set of equations (usually in matrix form) that will appropriately correct

the raw balance readings.

c. The Parks College wind tunnel features first-order balance corrections already incorporated in the data acquisition program, therefore the student does not need to

worry about this correction in a direct manner.

3) Tare and Interference Drag Corrections (4.17):

a. Tare and interference drag: The supports that hold the model in place will affect the airflow around the model and will have some drag themselves. The effect of the

supports in the flow is called interference and the drag of the supports tare. The

evaluation of tare and interference drag is a very complex and lengthy procedure.

Shielding the supports can minimize tare, but the added size of the shielding would

probably increase the interference drag. The combined effect can be accurately

determined using the "image method". For the level of accuracy desired in the lab, it

is enough to use thin, unshielded model supports. In this fashion, the interference

drag is minimized, and the tare drag can be approximately computed by a tare run,

that is, running the tunnel with the supports only and recording the observed forces.

b. Moment transfer to the model's center of gravity: Some important definitions follow: The balance moment center is the location about which the balance measures the

three different moments (rolling, pitching, and yawing moments). The trunnion is the

point about which the model rotates in pitch. All measured moments must be

transferred from the balance moment center to the model's CG, which at times can be

complicated if the trunnion and the moment center do not coincide. In most cases,

however, balances are designed so that the trunnion and the moment center are

coincident. Note that there are no corrections needed to transfer the forces to the

model's CG. In the Parks College tunnel, the trunnion and the moment center

coincide and are located at the holes near the tip of the front mounting struts, like

shown in Fig. 2.

c. Weight Tare Correction: Usually, the balance moment center location does not coincide with the model CG location. Consequently, when the model is pitched,

yawed or rolled the displacement of the model's CG will generate a moment that the

balance will register, and that must be subtracted from the balance reading.

4) Boundary Corrections:

a. There is no difference in having the model still and the air moving instead of the other way around, as is done in wind tunnels, but there is a noticeable difference in testing

the model in the open atmosphere or in a closed, bounded test section. The presence

of the lateral boundaries produces a number of effects, the most important of which

are:

i. Solid Blockage (Sec. 6.10): The presence of the model in the closed test section poses an obstruction to the flow, thus increasing the effective dynamic

pressure in the test section. Forces and moments at a given angle of attack are

also increased. If solid blockage corrections are to be kept small, it is

recommended that the model frontal area is no greater than 7.5% of the test

section cross-sectional area. This effect is often negligible in open test section

tunnels.

ii. Wake blockage (Sec. 6.11): The test section boundaries prevent the wake left by the model from expanding freely. This effect increases with the increase of

wake size (drag). It is also negligible in an open test section, since the

airstream is then free to expand in a normal manner.

iii. Alteration to the local angle of attack along the span (streamline curvature): In a closed test section the angles of attack near the wing tips of a model with

large span are increased excessively, making the tip stall start early. The

effect of an open jet is just the opposite. In both cases, the effect becomes

negligible if the model span does not exceed 0.8 of the test section width.

Additional corrections can be found in Ch. 6 of Pope Low Speed Wind Tunnel Testing.

NOMENCLATURE:

(Coefficient)u Uncorrected coefficient

(Coefficient)C Corrected coefficient

B Test section width

b Geometric wingspan

be Effective wingspan

bv Vortex span

C(subscript) Coefficient of (subscript)

CM Moment coefficient (roll, pitch, yaw)

d Maximum diameter of fuselage

H Test section height

k Effective wingspan/tunnel width, be/B

K1 Body shape factor for blockage K2 Fuselage shape factor for blockage

l Length of body

lt Distance from CG to (1/4) MAC of tail

q Freestream dynamic pressure

qC Corrected freestream dynamic pressure

Re Reynolds number

SW Model reference area (Wing Area)

St Tail area

g Geometric (uncorrected) AOA

C Corrected AOA

Boundary correction factor

t Total solid blockage correction factor

sbB Body solid blockage correction factor

sbW Wing solid blockage correction factor

struts, windshields Strut and windshields solid blockage correction factor

1 Tunnel correction factor for blockage

2 Downwash correction factor

PROCEDURE:

The complete aircraft to be tested is the Diamond DA-20 (Fig. 4), which is the model used for

Parks flight school. The model has a removable tail and wings for multiple configurations.

Figure 4. Diamond DA-20.

Run

Number

Velocity Board

Airspeed Comments

1 -8 to +16 in 2

increments 0 mph Tare Run: Full Model, No Wind

2 same 65 mph Full Model: no roll, no yaw

3 same 65 mph Full Model: added roll, no yaw

4 same 65 mph Full Model: no roll, added yaw

5 same 65 mph Tail Off Model: no roll

6 zero AOA only 65 mph Tare Run: No Model

The Tail Off run will allow determination of the wing-only lift coefficient, CLW. This lift

coefficient will be used later in calculating different corrections to the angle of attack.

1) Familiarize yourself with the basic wind tunnel operating procedures, the speed control, the balance and the Lab View data acquisition program.

2) Lock the balance and install the model in the test section, using unshielded struts. Assemble the aluminum floor sections and seal any gaps with masking tape. Carefully, pitch the

model trough the full angle of attack range and check for any binding.

3) Determine the zero geometric angle of attack (g), by using an angle of attack indicator or level. Afterwards, TARE the AOA indicator in the computer.

4) Unlock the balance. TARE balance readings. Before starting a run make sure nothing is left in the section (i.e. tools, tape).

5) Close test section doors (both!!!).

6) Check for any persons around the exhaust of the tunnel. Say CLEAR aloud, start the tunnel, and accelerate to the proper airspeed.

7) Proceed through the pitch run, taking data at the appropriate AOA. Once the run is complete, record any observations made during the test.

8) Stop the tunnel motor and slowly close the speed control.

9) Using the Save Data icon, save the data on a USB flash drive.

a. NOTE: The data is saved in a standard text file format with the data items separated by commas, and can be read by any spreadsheet program. An electronic spreadsheet

program (Excel) is highly recommended to speed-up data reduction. When using Excel

to manipulate data (recommended), select the Text Comma Separator options from the Load File menu, prior to loading the data text file.

10) LOCK the balance before reconfiguring the model for the next run.

11) Reconfigure the model for the next run (ask instructor for assistance).

12) TARE the balance before commencing the next run.

13) When all runs are complete record or compute the following data:

Wing

Wing area (SW) Wing MAC ( ) ARW Wing Volume

Wingspan (b)

Nominal CG location (25% )

Horizontal Tail:

Tail area (ST) Tail MAC ( ) ART lT: distance from CG to 25%

Model Trunnion Location with respect to CG (see Fig. 5)

Fuselage Volume (1)

Wind Tunnel:

Height (H) and width (B) of test section

Test section Fillet dimensions (to properly calculate tunnel cross-sectional area)

Thickness and Height of Struts

(1)

A good approximation of the volume of a streamlined body of revolution is V=0.45 l d2 where l is the length of

the model and d is the largest diameter of the body

ambient pressure and temperature

ambient density

model Reynolds number

DATA REDUCTION:

It is assumed that there is zero velocity and zero flow direction variation across the test section,

therefore corrections (A.1) and (A.2) will not be applied. The horizontal buoyancy correction

(A.3) will also be neglected, although all that it is needed to apply it is the longitudinal static

pressure gradient of the test section, which is easily measured.

The corrections for balance interaction (B) have already been applied to your data by the data

acquisition program.

At this point, subtract the No Wind Tare Run values and the No Model Tare Run values from the

model runs. This will yield the uncorrected values, which will be designated by a u subscript ()u.

If the dynamic pressure for all the runs was the same, this can be done in force and moment

units. Otherwise, it will need to be done using non-dimensional coefficients.

In equation form:

Du = D - DNo Model Tare Run

Mu = M - MNo Wind Tare Run

or

CDu = CD - CDNo Model Tare Run

CMu = CM - CM,No Wind Tare Run

Solid Blockage Correction:

2

3

11 ) volumewing(

X

KsbW

2

3

13 )ebody volum(

X

KsbB

X is the test section cross-sectional area

is obtained from Fig. 6.14

K1 and K3 are obtained from Fig. 6.13

Total solid blockage is

Wake Blockage Correction:

Use the equation below to account for wake blockage.

)(4

5

4DODiDu

WDO

Wwbt CCC

C

SC

C

S

For unseparated flow, CDu = CDO + CDi, hence, making the right most term of this equation zero.

For separated flows however, CDu > CDO + CDi, and the full equation must be used (refer to

Pope). For the angle of attack range assigned, one can assume the flow remains unseparated. If

one assumes that the drag can be represented by the usual drag polar equation, CD = CDo+KCL2,

the parasite drag coefficient, CDO, can be found using a CD (y-axis) versus CL2 (x-axis) graph.

The slope of the graph will determine K, and the intercept on the CD axis will determine CDO.

Furthermore, it is known that, Re

1

AK

so Oswald's efficiency factor e can be determined.

The struts and windshields (strut fairings) also create wake and solid blockage. Their

contribution can be calculated by

AreaSection Test

Area Frontal

4

1, swindshieldstruts

The total blockage correction is then et=esbt+ewbt+estruts,windshields and the corrected value of the test

section dynamic pressure is qC=q(1+t)2. This corrected value of the dynamic pressure is to be

used in all subsequent calculations. If coefficients were computed using the uncorrected

dynamic pressure, these coefficients will need to be recalculated using the corrected dynamic

pressure.

Now the data reflects the correct aerodynamic forces on the balance (i.e. it is corrected for all

tares and weights) and has the correct dynamic pressure. Note that the lift coefficient is fully

corrected at this stage.

Moment Transfer:

The moments will be transferred now to the model's nominal CG, assuming the trunnion and the

balance moment center are coincident (ours are, indeed), as shown in Fig.5. Use the equation

CmCG,u = Cmu - x(CL cos + CDu sin ) - y(CDu cos - CL sin )

or

MCG,u = Mu - x(L cos + Du sin ) - y(Du cos - L sin )

to perform this transfer. The variables x and y are the horizontal and vertical distances from the

trunnion (Fig. 5, yellow circle) to the nominal CG location (25% ).

Figure 5. Moment transfer from trunnion to CG.

Wall Corrections:

As stated in the introduction, wall corrections will affect the angle of attack, drag, and pitching

moment coefficients.

The corrected angle of attack is c = g + up + w

Where g is the geometric (measured) AOA

up is correction (A.2), assumed zero

w = (1+2)(SW/C)(180/)CLW

CLW is the wing-only lift coefficient, obtained in run #4 or #5

is found as follows:

a) from Fig. 6.23, find bv/b

b) find effective span: be = (b+ bv)/2

c) use Fig. 6.29 to find

2 is found from Fig. 6.52, with k = be/B

The 2 factor in this correction compensates for streamline curvature induced by the tunnel walls.

When performing the calculations for the no-flap configuration, use the CLW corresponding to

the no flaps wing-only run. Likewise, use the flaps-down CLW for the flaps down corrections.

The drag coefficient corrected for blockage effects, wall effects, and tare is CDC = CDu + CDup +

Cdw, assuming the corrected value of dynamic pressure (qC) was used in all calculations. The different terms in this equation stand for:

CDu is the uncorrected (measured) drag coefficient

CDup is correction (A.2), assumed zero

Cdw = CLW2 (SW/C)

The wall correction to Cm can be significant. It is computed as follows:

CmCGC = CmCGu - CmCGt, and

CmCGT = , where

The term is the variation in pitching moment coefficient with horizontal tail incidence

angle which may be found performing another wind tunnel test (see Expt. 3). If experimental

measurements are not available, it can be estimated as follows:

= -aT TV

Assuming a two-dimensional tail lift curve slope of 0.100 per degree att = 0.0533. Also:

WW

TT

MACS

SlV .

The parameter 2 may be found from Fig. 6.52.

RESULTS:

I. Generate the following using raw data gathered (uncorrected for anything) but

transferred to the nominal CG (25% ). Doing this will help determine if corrections were reasonable or not.

a) CLu vs. g (tail off and tail on)

b) CDu vs. g (tail off and tail on) (if drag needs a sign change do it here)

c) CMpitch,u vs. g (tail off and tail on)

d) CMpitch,u vs. CLu (tail off and tail on)

II. Apply the corrections described above and obtain the following corrected plots:

e) CLC vs. C (tail off and tail on)

f) CDC vs. C (tail off and tail on)

g) CMpitch,C vs. C (tail off and tail on)

h) CMpitch,C vs. CLC (tail off and tail on)

Comment about the differences between the corrected and uncorrected data.

III. Using the corrected data, plot the following:

i) CMroll,C vs. C (no roll and roll added tests)

j) CMyaw,C vs. C (no yaw and yaw added tests)

Comment how the added roll and added yaw affects the aircrafts performance (ex. stall, yaw ability).

IV. Also find the following:

k) The static margin of the complete model (tail on) with CG at 25% and the neutral point of same in terms of percent . Does this aircraft satisfy the criteria for longitudinal static stability?

l) CLmax comment about the value.

m) Cdmin and zero-lift AOA (0L ) comment about these characteristics.

n) Lift curve slope comment about the value and how it compares to other aircraft.

o) Find the Reynolds number of the true-size aircraft cruising at 10,000 ft. (hint: find the scaling ratio of the model to the 358 true-size wingspan). Comment how the experiment could be changed based on your findings.

APPENDIX:

A word about trip strips:

As you probably know, transition from laminar to turbulent boundary layer depends greatly on

the Reynolds number (Fig. 6). The maintenance of a laminar boundary layer is favored by low

values of RN. Thus, wind tunnel models, which usually operate at a lower RN than real aircraft,

have too much laminar flow over their wings when compared to the full-size aircraft. A trip

strip, or artificial roughness, is used to fix the location of transition from laminar to turbulent

boundary layer, and in this way the flow over the models wing will more closely resemble that of the full-size aircraft. It is important to simulate the correct transition location if accurate drag

readings are desired.

A trip strip is usually 0.125 to 0.250 in wide and can be made by blowing grit over wet adhesive

placed on the wing. Sometimes, a trip strip is constructed by placing several layers of tape at the

desired chord-wise transition point; even a string glued at the proper chord-wise location will

work. Refer to Section 7.1 for more information.

Figure 6. Effect of RN on boundary layer transition. The RN is greater in case (b), thus

forcing the transition point forward. Trip strips can be used on wind tunnel models, which

usually fly at a lower RN, to simulate the transition point found on the real aircraft.

REFERENCES

1) 1.Low Speed Wind Tunnel Testing, 2nd ed., W. H. Rae and A. Pope, John Wiley & Sons, 1984

2) Determination of Boundary Corrections in the McDonnell Polysonic Wind Tunnel. D.G. Shors, McDonnell Aircraft Corporation, 1968.

FIGURE 6.52

FIGURE 6.29

PROCEDURE FOR STARTING THE DATA ACQUISITION SYSTEM ON THE

SUBSONIC WIND TUNNEL (Prepared by Frank Coffey)

1) Turn on the computer and monitor.

2) The login username is aerolab and the password is LSWT.

3) Once the desktop is up click on the balance icon. Choose aerolab2.vi from the menu.

4) To start the program click on the black arrow in the upper left-hand corner. Tare out the balance readings. MAKE SURE THE BALANCE IS UNLOCKED. To take data press

the grab button. To save the data press the yellow save button. Once the data is saved to a disk the data will be lost on the balance program.

5) There is a vertical slider bar on the left hand side of the screen. This is to smooth out the fluctuation of the data. Move the slider to 50% range. Wait a couple of seconds then

lower the slider to 25% range. Wait a couple of seconds then move it to around 10%

range. Wait a couple of seconds, now you are ready to take a data point. Once the data

point is taken, move the slider back to 50%. Move model or make adjustments to test

and then repeat the step down procedure for the next data point.

6) If you want to view the data on the graphs on the right hand side of the screen while the test is running you MUST set the x and y-axis BEFORE you begin taking data. To clear

the graphs right click on the graph and choose clear graph before you start the next data run.

7) To stop the program click on the red stop sign icon in the upper left hand corner. You will lose all data if you stop the program. To close the program click on the X in the upper right hand corner of the screen. Shut down the computer and turn off the monitor.

8) If you encounter any problems or have any questions, get Frank Coffey the lab technician. DO NOT proceed if you have any doubts!!!