NASA Contractor Report 181 871 

AN EVALUATION OF THREE EXPERIMENTAL PROCESSES 

FOR TWO-DIMENSIONAL TRANSONIC TESTS 

Gennaro ZUPPARDI 

NCC1-47 

JIAFS Core Program August 1989 

NASA 

National Aeronautics and 

Space Administration 

Langley Research Center 

Hampton, Virginia 23665-5225 

[RASA-CR-t81871) A11 EVALUATIOIS OF TBPBB 1989-27674 

EXPEBZBEPTIhL PROCESSES RDE TUO-DIHEISIOIAL 

TRARSOHXC TBSTS [Eaples tlniv,) 71 p 

CSCL 14B Uaclas 

G3/09 0224575 
SUMMARY 

Aerodynamic measurements in conventional wind tunnels suffer from the 

interference effects of the sting supporting the model and the test section walls. These 

effects are particularly severe at transonic speeds where they can significantly affect the 

aerodynamic measurements. 

Sting interference can be completely overcome by using Magnetic Suspension 

techniques. Wall interference can be reduced by; testing airfoils in conventional, ventilated 

tunnels at relatively small model to tunnel size ratios, treatment of the tunnel wall 

boundary layers, or by using Adaptive Wall Test Sections (AWTS). 

This paper assesses the operating capabilities and results from two of the foremost 

two-dimensional, transonic, AWTS tunnels in existence. These are the NASA Langley 

0.3-m Transonic Cryogenic Tunnel and the ONERA-CERT T-2 tunnel in Toulouse, 

France. In addition, this paper evaluates the results derived from the well known 

conventional Canadian facility, the NAE 5’ x 5’ wind tunnel. The CAST 10/DOA2 Airfoil 

results are used in all of the evaluations. 

i 
CONTENTS 

1 .................................................... . INTRODUCTION 1 

2 .................................... . PRINCIPLE OF WALL STREAMLINING 2 

3 . COMPARISON OF THE NASA 0.3-m TCT AND THE ONERA T-2 ................. 3 

3.1 Phvsical Characteristics ............................................... 3 

3.1.1 Tunnel Circuit ................................................. 3 

3.1.2 Test Section .................................................. 4 

3.1.3 Pressure Measurement ........................................... 5 

3.1.4 Wind Tunnel Computer .......................................... 6 

3.15 Sidewall Boundary Layer Control System (BLC) .......................... 

3.2 Adantation Strategies ................................................ 7 

3.2.1 NASA0.3-mTCT ............................................. 7 

3.2.2 ONERAT-2 ................................................. 9 

3.2.3 Comparison of the NASA and ONERA Adaptation Strategies ................ 

6 

11 

4 . NAESx5'WINDTU"EL .............................................. 13 

4.1 Phvsical Characteristics ............................................... 13 

4.1.1 Tunnel Circuit ................................................. 13 

4.1.2 Test Section .................................................. 14 

4.2 Interference Correction Method ......................................... 14 

4.2.1 Mach Number Correction ......................................... 4.2.2 Angle of Attack Correction ........................................ 16 

16 

5 . MODELDESCRIPTION ................................................ 17 

6 . ANALYSIS OFTHEDATA ............................................. 18 

6.1 Evaluation of Secondarv Effects ......................................... 18 

6.1.1 Transition Tripping ............................................. 19 

6.1.2 BLC Suction Velocity ............................................ 6.1.3 Sidewall Interference Effect ........................................ 19 

19 

6.2 Comparison and Validation ............................................ 20 

6.2.1 General Remarks .............................................. 20 

6.2.2 GRUMFOILCode ............................................. 20 

6.2.3 Results of the Validation Process ................................... 21 

6.2.4 Pressure Distribution ............................................ 21 

6.2.5 Shock Location ................................................ 21 

6.2.7 Lift Curve Slope ............................................... 6.2.6 Drag Curve .................................................. 22 

22 

7 . CONCLUDINGREMARKS 24 

8 . REFERENCES ...................................................... 25 

............................................. 

9 . ACKNOWLEDGEMENTS ............................................... 27 

ii 
LIST OF ILLUSTRATIONS 

Fig. 1 Flow chart of wall adaptation principle. 

Fig. 2 

Fig. 3 

Artistic illustration of 2-D wall adaptation principle (from Ref. 5). 

Schematics of NASA 0.3-m TCT and ONERA T-2. 

Fig. 4 Operating envelopes of NASA 0.3-m TCT and ONERA T-2 (from Ref. 11) at several stagnation 

pressures and temperature. 

Fig. 5 Schematics of test sections of NASA 0.3-m TCT and ONERA T-2. 

Fig. 6 Representation of physical wall as a flat vortex sheet by NASA wall adaptation strategy. 

Fig. 7 

Fig. 8 

Flow chart of ONERA T-2 wall adaptation strategy. 

Schematic of NAE 5’s tunnel. 

Fig. 9 Estimated test time and Reynolds number per meter of NAE 5’s tunnel as a function of stagnation 

pressure at M, = 0.75. 

Fig. 10 Coordinate system of Mokry‘s method (from Ref. 26). 

Fig. 11 Sketch of CAST 10-2/DOA-2 airfoil. 

Fig. 12 Artistic illustration of lift curve of a super-critical airfoil. 

Fig. l3 Lift curves of Canadian CAST 10 airfoil model from NAE 5’s tunnel, M, w 0.765, R, w 20 x lo6. 

Fig. 14 Drag curves of Canadian CAST 10 airfoil model with free and fmed transition from NAE 5’s tunnel, 

M, w 0.765, Rc u 20 x lo6. 

Fig. 15 Lift curves of Canadian CAST 10 airfoil model from NASA 0.3-m TCT, M, w 0.765, R, ss 20 x lo6, 

fmed transition. 

Fig. 16 Drag curves of Canadian CAST 10 airfoil model from NASA 0.3-m TCT, M, = 0.750, Rc = 20 x lo6, 

fmed transition. ’ 

Fig. 17 Lift curves of Canadian and French CAST 10 airfoil model from NASA 0.3-m TC”, M, ss 0.765, Rc w 

20 x lo6, fmed transition. 

Fig. 18 Pressure distribution on Canadian CAST 10 airfoil model from NASA 0.3-m TCT, NAE 5’xS tunnel, 

and GRUMFOIL code, M, w 0.765, R, ~20 x lo6, fmed transition. 

Fig. 19 Pressure distribution on French CAST 10 airfoil model from NASA 0.3-m TCT, ONERA T-2 tunnel, 

and GRUMFOIL code, M, N 0.765, Rc w 20 x lo6. 

Fig. 20 Pressure distribution on French CAST 10 airfoil model from NASA 0.3-m TCT and ONERA T-2 

tunnel, M, w 0.765, RE w 20 x lo6, C, w 0.62, free transition. 

Fig. 21 Artistic illustration of criterion to locate shock wave on airfoil upper surface. 

Fig. 22 Shock location on Canadian CAST 10 airfoil model as a function of lift coefficient from NASA 0.3-m 

TCI’, NAE 5’s tunnel, and GRUMFOIL code, M, RI 0.765, Rc w 20 x lo6, fmed transition. 

... 

111 
Fig. 23 Shock location on French CAST 10 airfoil model as a function of lift coefficient from NASA 0.3-m 

TCT, ONERA T-2 tunnel, and GRUMFOIL code, M, w 0.765, R, w 20 x lo6, fmed transition. 

Fig. 24 Drag curves of Canadian CAST 10 airfoil model as a function of Mach number from NASA 0.3-m 

TCT, NAE 5%’ tunnel, and GRUMFOIL code, R, w 20 x lo6, C, w 0.5, fmed transition. 

ONERA T-2 tunnel, and GRUMFOIL code, R, w 4 x lo6, C, w 0.5, fmed transition. 

TCT, NAE 5’s tunnel, and GRUMFOIL code, R, w 20 x lo6, fmed transition. 

Fig. 25 Drag curves of French CAST 10 airfoil model as a function of Mach number from NASA 0.3-m TCT, 

Fig. 26 Lift curve slope of Canadian CAST 10 airfoil model as a function of Mach number from NASA 0.3-m 

Fig. 27 Lift curve slope of Canadian CAST 10 airfoil model as a function of Reynolds number from NASA 

0.3-m TCT, NAE 5’s tunnel, and GRUMFOIL code, M, w 0.765, fmed transition. 

Fig. 28 Lift curve slope of French CAST 10 airfoil model as a function of Reynolds number from NASA 0.3-m 

TCT, ONERA T-2 tunnel, and GRUMFOIL code, M, w 0.765, fmed transition. 

CAST 10 airfoil, M, w 0.765, R, w 20 x lo6, fmed transition. 

Fig. 29 Experimental lift curve from NASA 03-m TCT and theoretical lift curve from GRUMFOIL code for 

Fig. 30 Maximum lift coefficient of Canadian CAST 10 airfoil model as a function of Reynolds number from 

NASA 0.3-m TCT and NAE 5’s tunnel, M, w 0.765, fmed transition. 

Fig. 31 Maximum lift coefficient of Canadian CAST 10 airfoil model as a function of Mach number from 

NASA 0.3-m TCT and NAE Yx5’ tunnel, Rc w 20 x lo6, fmed transition. 

Fig. 32 Drag profile of Canadian CAST 10 airfoil model as a function of test section span from NASA 0.3-m 

TCT and NAE Yx5’ tunnel, M, w 0.765, RE w 20 x lo6, C, N 0.27, fmed transition. 

TCT and NAE Sx5’ tunnel, M, w 0.765, R, w 20 x lo6, fmed transition. 

Fig. 33 Drag profile of Canadian CAST 10 airfoil model as a function of test section span from NASA 0.3-m 

LIST OF TABLES 

I. Physical and operating characteristics of the NASA 0.3-m TCT and ONERA T-2 test sections. 

11. Physical and operating characteristics of the 2-D, transonic test section of the NAE Sx5’ tunnel. 

111. Geometrical parameters of the model-wind tunnel systems. 

iv 
NOMENCLATURE 

A 

%b 

C 

SF 

T 

Ta 

U 

x', y' 

Y 

(L 

B 

7 

A(L 

AM 

Au, Av 

AZ 

ACi 

non-dimensional cross sectional area of an airfoil 

intercept and slope of a least square linear fit 

model chord 

drag coefficient 

lift coefficient 

lift curve slope 

normal force coefficient 

pressure coefficient 

coupling factor (CF = 035) 

cubic spline function of the pressure coefficients 

ratio of the specific heats 

length, height and width of the test section 

Mach number 

iterations number 

stagnation pressure 

independent flow variables 

weighing function 

Reynolds number based on model chord 

Reynolds number per meter 

correlation coefficient of a least square linear fit 

scaling factor (SF=0.8) 

stagnation temperature 

ambient temperature 

velocity along the tunnel axis 

streamwise and normal components of the perturbation velocity 

local velocity parallel to the wall 

BLC suction velocity 

wall/jack coordinates 

integration limits of the wall adaptation strategies 

pressure measurement limits of the ONERA T-2 tunnel 

airfoil coordinates 

tunnel reference system of the Mokry's technique 

spanwise coordinate of the test section 

angle of attack 

Prandtl-Glauert coefficient 

vorticity strength 

correction to angle of attack 

deviation from linear lift curve 

correction to Mach number 

corrections to the components of the perturbation velocity 

wall/jack aerodynamic movement 

wall/jack actual movement 

distance between C and the wall 

boundary layer displacement thickness 

convergence criterion 

coordinate along C 

doublet strength 

non-jack coordinate along the representation plane of the wall 

reduced coordinates of the Mokry's method 

control plane near the wall 

potential of the perturbation velocity 

flow angle 

weighing factor 

V 
Subscripts 

c, F 

f 

n 

r 

C 

1 

S 

sw 

W 

00 

ceiling, floor 

compressible 

incompressible 

number of streamlining jacks 

real 

shock 

sidewall 

wall interference 

free stream 

imaginary 

Superscripts 

I Ith iteration step 

P weighed 

w' wall projection 

special symbols 

AWTS adaptive wall test section 

BL boundary layer 

BLC 

DGT/IGT direct/hverse green transform 

GN2/LN2 gaseous/liquid nitrogen 

sidewall boundary layer control system 

vi 
1. INTRODUCTION 

The aerodynamic measurements in conventional wind tunnels suffer from the interference effects of: 

the sting supporting the model and the test section walls. These effects are particularly severe at transonic 

speeds where they can significantly affect the aerodynamic measurements. 

Sting interference can be overcome by using Magnetic Suspension techniques (Ref. 1). This technique 

allows the researcher to test models without any physical support or contact with the tunnel. 

Tunnel wall interference has been typically alleviated by decreasing the solid blockage by physical 

methods or by applying analytical corrections. The decrease of the solid blockage involves: 

1) Enlarging the test section. This, obviously, increases the initial investment, and the operating 

and maintenance costs of the facility. 

2) Reducing the size of the model. This often decreases the test Reynolds number to 

unacceptable values. 

The effective use of the corrective methods is restricted to Correctable flow fields. A flow field is 

correctable when we can consider that in the region of the model both the streamwise (u,) and the normal (vw) 

component of the wall interference velocity are constant. For correctable flows we can achieve a strict 

equivalence of the free air and the constricted flow fields. Simple formulas are available to compute, in 

subsonic flows, the correction terms Au, and Av, to uw and v, respectively (Ref. 2). The computation of Auw, 

Av, in transonic flows suffers from several drawbacks: 

1) It requires powerful computers and time consuming procedures. In addiction, the corrected 

data are usually not available at the end of the test because the complexity requires correcting 

the data after the testing is finished (post run processing). 

2) The inability to establish accurately the boundary conditions on ventilated (slotted and/or 

perforated) test sections increases the computational difficulty. 

3) The corrections are unreliable when the test is performed under operating conditions not 

matching the theoretical hypothesis of the correcting method. 

The Adaptive Wall Test Section (AWTS) can overcome the above mentioned disadvantages. This type 

of facility minimizes the solid blockage without physically increasing the test section size or decreasing the 

model size. The AWTS eliminates or reduces to correctable values the induced wall interference velocity. 

The concept of the AWTS is not recent (Ref. 3). It was developed around 1937. Researchers built the 

first AWTS (12.7 cm x 5.08 cm) with a flexible ceiling and floor at the National Physical Laboratory (England) 

in 1937-38. Even though the tests were encouraging, adaptive wall researches did not continue beyond the 

1950s. This was due probably to the lack of computing power and the advent of ventilated test sections for 

transonic tunnels. The lack of computing power involved manual operation of the adaptive procedure and 

therefore lengthy and laborious working. 

The interest in the AWTS was renewed in the early 1970s. Both computers and technology were ready 

to allow full-automatic AWTS operation. New and more sophisticated adaptation strategies were developed. 

1 
Several such facilities were built both in Europe (England, France, Germany) and in the United States. In 1982 

Ashill (Ref. 4) quantified the benefits of the AWTS over the conventional rigid wall wind tunnel in reducing the 

wall interference. He computed the correction terms Auw and Avw from two sets of data for a NACA 0012 

airfoil. This airfoil was tested in the ONERA T-2 tunnel at several test conditions. One set of data was 

obtained with straight walls and the other one with adapted walls. Typically, for the test case M, = 0.73 and a 

= 0, the computations showed that the variation along the model chord of both the streamwise and the normal 

wall interference velocities were substantially smaller for the adapted walls. The correction to Mach number 

(that is, to uw) was 0.023 and 0.002 for the measurements with straight and adapted walls, respectively. The 

correction to angle of attack (that is, to vw) was -0.007 and 0.005 for the measurements with straight and 

adapted walls, respectively. These advantages over conventional tunnels allow the AWTS to become a unique 

test facility for making accurate 2-D transonic tests. 

The purpose of the present study is to assess the operating capabilities and the reliability of the 

measurements of two of the best 2-D, transonic AWTSs: the NASA 0.3-m TCT and the ONERA T-2, and one 

of the most advanced conventional tunnels of the western world, the NAE 5’s wind tunnel in Ottawa. For the 

purpose of this study we will refer to these tunnels as the NASA, ONERA, and NAE tunnels. The report will 

develop along three main directives: 

1. Description of the technical features and the related operating capabilities of the three wind 

tunnels. 

2. Description and comparison of the wall adaptation strategies for the NASA and ONERA 

AWTSs and a description of the post-test analytical correction method for the NAE tunnel. 

3. Comparison and validation of the measurements, made in the three tunnels on the CAST 

10-2/DOA-2 airfoil, in transonic, high Reynolds number flows. We will refer to this airfoil as 

CAST 10. 

The last part of this report is still today the subject of a cooperative, international program between 

NASA, NAE, ONERA, and DFVLR to assess methods to alleviate the wall interference effects in transonic 

flows. 

2. PRINCIPLE OF WALL STREAMLINING 

The basic working principle of an AWTS is to eliminate the wall interference at its source by adjusting 

the test section boundaries (Ref. 2,5,6,7). This concept involves the shaping (streamlining) of the test section 

boundary,S, (at or near the test section wall) to a free air streamtube that encloses the model for each test 

condition. The shaping of the far field tunnel boundary cancels the wall aerodynamic interferences. 

The boundary S splits the flow field into two parts: 

1. The real flow field inside the boundary S. 

2. An imaginary, irrotational flow field, extending from the boundary S to infinity. 

The boundary S has to be located away from the model in order to avoid considering the viscous effects 

and, at the same time, to attenuate the model perturbation. The imaginary flow field is computed by the small 

2 
perturbation theory relative to the wall. 

The general streamlining principle relies on matching on S two independent aerodynamic variables p 

and q in the real (p, q,) and imaginary flow field (pp qf). The variables p, and q, (that is, the components u, 

and v, of the perturbation velocity) are measured on S. One of them (for example, qr) is used as a boundary 

condition to compute pf on S. The streamlining of the boundary S is achieved if pf and p, are equal within a 

convergence criterion e. The deformation of the boundary S influences, at the same time, both the real and the 

imaginary flow field. The matching of the imaginary and real flow field is complex and necessarily iterative. 

The streamlining procedure follows the flow chart on figure 1. The remarkable feature of the adaptation 

procedure is that the wall shaping can be determined without computing the flow field about the model. 

The streamlining or adaptation procedure is simpler for 2-D flow fields where only two 2-D control 

surfaces are needed to control the boundary shape. The flow field is broken into three parts: the real part, 

which contains the model, and two imaginary parts. One imaginary part is above and the other is below the test 

section. Figure 2 explains graphically this concept. 

The streamlining principle has been shown to be feasible with two different approaches: i) ventilated 

(perforated and/or slotted) wall test section and ii) impermeable, flexible walled test section. In the first 

approach, the control boundary S is inside the test section. The streamlining is fulfilled by modifying the 

velocity vector on the boundary S. Sucking or blowing, as appropriate, through the porous ceiling and floor 

control the velocity vectors on S. In the second approach, the control boundary S coincides physically with the 

ceiling and the floor (neglecting the wall boundary layer). The walls are streamlined by simply controlling the 

wall shape by a system of jacks. 

3. COMPARISON OF THE NASA 0.3-m TCT AND THE ONERA T-2 TUNNEL 

3.1 Phvsical Characteristics 

3.1.1 Tunnel Circuit 

The NASA 0.3-m TCT (Ref. 8,9) and ONERA T-2 (Ref. 7,10,11,12,13) wind tunnel were designed 

to provide adequate Reynolds number and Mach number to simulate the design flight conditions of the current 

transonic transportation aircraft. Both tunnels are closed circuit, pressurized, and cryogenic. The Mach 

number operating range is about O.lcM, cO.95 for the NASA tunnel and 0.3cM, c 1.1 for the ONERA tunnel. 

The operating range in terms of Reynolds number per meter (at T= 120 K) is about 107cRmc3x108 for the 

NASA tunnel and 3x107cRmc3~108 for the ONERA tunnel. 

The circuits of the NASA and ONERA tunnels are about 28m and 25m long, respectively. We show 

the schematics of the tunnels in figure 3 and the operating envelope diagrams of the NASA (Ref. 14) and 

ONERA (Ref. 11) tunnel in figure 4. The data shown in the diagrams are related to empty test section. 

The most important difference is in the working of the circuits. The ONERA tunnel is intermittent 

and the flow is produced by induction. Air is pressurized by a compressor in a 45m3 reservoir to a maximum 

pressure of 80 bars. The velocity is regulated by adjusting the area of the throat downstream of the test section 

(see figure 5). The NASA 0.3-m TCT is a continuous flow fan-driven tunnel which uses gaseous nitrogen 

(GNJ as the test gas. The fan speed, and therefore the gas velocity and the Mach number, is automatically 

controlled by a computer. The liquid nitrogen (LNJ used to cool the tunnel is sprayed into the tunnel circuit 

3 
downstream of the test section. 

The short duration (90 to 120 seconds in transonic regime (Ref. 13)) of the runs, in T-2 causes some 

difficulties in making accurate tests. It causes problems with i) the control of the flow parameters, ii) the wall 

adaptation process, E) the data acquisition, and iv) the surface temperature of the model and the test section 

walls. In its original configwation, a typical run of the ONERA tunnel developed along the following steps: 

1) Starting the flow at low pressure (P = 1.1 bar) and low Mach number (Moo = 0.3), but at the 

right temperature. 

2) Introducing rapidly the model into the stream. An auxiliary facility pre-cools the model to 

obtain a thermal equilibrium between the model and the flow. 

3) Achieving the selected Mach number. 

4) Adapting the walls. 

5) Probing the wake and the pressure distribution on the model. 

A complete test requires typically two runs. The first run (steps 1 through 4) is to adapt the walls and 

the second run (step 5) is to probe the wake (Ref. 7). 

The injection flow feature allows the ONERA tunnel to achieve a lower turbulence level in the test 

section than the NASA tunnel. The turbulence, measured in the empty test section at T = 120 K and M, w 0.8, 

is about 0.0015 and about 0.011 for the ONERA (Ref. 10,15) and NASA (Ref. 16) tunnel, respectively. The 

turbulence level is defined as the root mean square of the pressure coefficient fluctuation. 

3.1.2 Test Section 

We show the schematics of the NASA and ONERA test sections in figure 5 and summarize the related 

physical and operating characteristics in Table I. 

The dimensions in centimeters (Length x Height x Width) of the NASA and ONERA test sections are 

185 x 33 x 33 and 132 x 37 x 39, respectively. The height and the width of both test sections are related to the 

entrance of the test sections. The flexible walls of both tunnels are made from special steel: 304 stainless steel 

for the NASA tunnel and INVAR steel for the ONERA tunnel. The walls must remain flexible down to 

cryogenic temperatures (u 100 K). The wall thickness of the NASA test section varies along the wall length. 

The walls are less thick in the area close to the model. This is where the maximum wall curvature occurs. The 

wall thickness is: i) 0.32 cm from the entrance to the jack #6 and from the jack #17 to the exit, ii) 0.16 crn 

between the jacks #6 and #17. The wall thickness of the ONERA test section is constant (0.13 cm) along the 

entire length. The walls of the NASA test section (Ref. 8) are fixed i) vertically and longitudinally at the 

upstream of the test section and ii) only vertically downstream. The walls can translate longitudinally as the 

wall shape changes. The downstream extremities of the walls of the ONERA test section are free in the vertical 

movement. Two moving flaps adapt the connection between the test section and the diffuser as the walls are 

deformed. 

The first 18 jacks of the NASA test section are active. The jacks #19, #20, and #21 are not used in the 

streamlining process. These jacks are used only to smooth the connection between the streamlining position 

4 
and the diffuser. The connection is fulfilled on the basis of a prefured criterion by starting from the position of 

the jack R18 (Ref. 17); thus the aerodynamic length of the test section is reduced to 142 cm. The first three 

jacks (Xl, X2, X3) and the last jack (#16) of the ONERA tunnel participate in the streamlining process only in 

a passive way. The movement of these end jacks is calculated from an extrapolation of the wall adaptation 

strategy beyond the active jacks #4 through Xl5. The aerodynamic length of the test section is about 83 cm 

(Ref. 18). No pressure information is used upstream of the point X, and downstream of the point Xu (see 

figure 5). The points X, and Xu are located dong the tunnel axis at a distance of 1.1 H (40.7 cm) upstream 

and downstream of the turntable center, respectively (Ref. 12). H is the test section height. The greater length 

of the flexible walls upstream of the model in the NASA tunnel should i) guarantee a more precise definition 

of the undisturbed flow conditions and ii) reduce the source of error in obtaining interference free flow. This is 

linked to the finite length of the adjustable walls. The larger number of jacks also should guarantee a better 

wall adaptation. 

The actuators deforming the walls of the ONERA and NASA tunnels are respectively hydraulic and 

screw jacks. All of them are controlled by electric stepper motors. For each tunnel the movement is 

transmitted by the jacks to the walls through a system of pull/push rods. Each actuator acts evenly on two 

points of the same rod. The local bending radius of the walls in both tunnels is checked before setting the new 

wall shape. The check is necessary to avoid exceeding the mechanical limit of the wall material. The two kinds 

of actuators are characterized by different features. The setting of the hydraulic jacks is faster but less fine than 

the screw jacks. Typically, in their current configuration: 

1) The minimum jack displacement of the ONERA test section is 0.02 cm and the displacement 

speed is 0.6 cm/s. 

2) The minimum jack displacement of the NASA test section is 0.01 cm and its displacement 

speed is 0.024 cm/s. 

The jack displacement transducers of the NASA and ONERA tunnels are linear variable differential 

transformers and potentiometric comparators, respectively. The accuracy of the systems is k0.013 cm and 

50.005 cm, respectively. The maximum movement of the jacks of the NASA test section is 7.62 cm upward and 

2.54 cm downward. The maximum movement of the jacks of the ONERA test section is 2.5 cm both upward 

and downward. The larger wall/jack movement gives the NASA tunnel an adaptation capability at more severe 

operating conditions: that is, higher angle of attack, higher Mach number, and larger test model. 

3.1.3 Pressure Measurement 

Both tunnels are equipped to measure static pressure on the model and on the test section walls, and 

total pressure in the wake downstream of the model. Each flexible wall of the ONERA tunnel has 104 pressure 

holes (Ref. 13); 58 are drilled near the centerline and 23 on each side. The last ones are used to make 

measurements in 3-D tests. The pressure holes are more closely spaced in the model area. Each flexible wall 

of the NASA tunnel has 18 pressure taps along the centerline of the flexible walls at the same longitudinal 

position as the active jacks. The wake rake of the NASA tunnel is formed by 6 probes traversing the semispan 

of the test section. The system can translate vertically about 25.4 cm. It can be located at three different 

positions downstream of the turntable center: 31.8 cm, and 44.5 cm, 57.2 cm (Ref. 8). The wake rake of the 

ONERA tunnel consists of one probe located at the centerline of the test section. The original papers (Ref. 7, 

10, 11,12,13) do not specify the position downstream of the model or the vertical movement of the wake rake. 

NASA and ONERA facilities use, in the pressure acquisition process, ten and seven 48-port scanning pressure 

valves, respectively. The pressure taps of the model and the wake rake in the NASA tunnel are linked directly 

5 
to individual pressure transducers. 

The pressure transducers in the NASA tunnel are capacitive type (Barocell). The transducers are 

mounted on thermostatically controlled heater bases to maintain a constant working temperature. The 

transducers used to measure the pressure on the model and on the flexible walls work in the range 26.9 bars. 

The transducers used to measure the pressure in the model wake work in the range k 1.4 bar. The average 

accuracy is +- 15 mbar (Ref. 8). The pressure value is typically the average of 20 measurements. The pressure 

transducers in the ONERA facility are differential type (Kulite). The transducers are only thermally isolated 

but not thermostatically controlled; temperature corrections are necessary for these transducers (Ref. 12). The 

average accuracy of these transducers is 22 mbars. The pressure value is typically the average of 80 

measurements (Ref. 10). 

3.1.4 Wind Tunnel Computer 

Both tunnels are equipped with two communicating minicomputers to acquire and process the data and 

to manage the tunnel operations. 

ONERA T-2 (Ref. 10): 

HFZlMX-F to acquire and manage data both on the model and walls and to run the numerical program of 

the wall adaptation strategy. 

HP21MX-E to monitor and control the wind tunnel test conditions. 

NASA 0.3-m TCT (Ref. 9,17): 

MODCOMP CLASSIC 7863 (CPU-A): 

MODCOMP CLASSIC 7863 (CPU-B): 

to acquire and manage wall data and run the numerical program of 

the wall adaptation strategy. 

to acquire and manage model data and input and check the test 

conditions: that is, Mach number, Reynolds number, and angle of 

attack. 

One MOTOROLA 6800 microprocessor controls the tunnel working parameters (temperature, pressure, and 

velocity). 

3.1.5 Sidewall Boundary Layer Control System (BLC) 

In 2-D tests, the model spans the test section from wall to wall; the sidewall BL makes a perfectly 

two-dimensional flow impossible. The sidewall BL also plays an important role in the aerodynamic 

interference. The effects due to the sidewall BL are particularly strong in transonic flows. In this situation, the 

boundary layer can separate from the sidewall, particularly where shock waves are present. This separation, in 

turn, introduces 3-D effects over the entire span of the model. Both test sections have a system to reduce the 

sidewall BL thickness. Both systems are still in development today. 

The systems.rely on the suction of the boundary layer through a porous plate. The suction region is 

located upstream of the model on each sidewall. Both the porous plates of the NASA (Ref. 8) and ONER4 

(Ref. 7) tunnels are located over the entire height of the test section. The length of the porous plate of the 

NASA tunnel is about 18 cm. The distance between the center line of the plate and the midpoint of the model 

6 
chord is 36 cm. The plate porosity is about 10 percent. The length of the plate of the ONERA tunnel is 18 cm. 

The distance between the center line of the plate and the midpoint of the model chord is 375 an. 

The NASA system is more powerful and versatile than the ONERA system. The NASA system, in 

fact, can work either in a passive or in an active mode (Ref. 19). The ONERA system can work only in a 

passive mode. The passive mode relies on the difference in pressure between the test section and the 

atmosphere. This system suffers from the following limitations: 

The static pressure in the test section must be higher than the ambient pressure. 

The maximum mass removal is limited by the LN2 injected to maintain steady operating 

conditions. The active mode overcomes these shortcomings. The gas is removed by a 

re-circulation system using a cryogenic centrifugal compressor. The re-circulation system 

re-injects the gas into the tunnel at the high-speed diffuser. The sucked mass is controlled, in 

both modes, by digital valves. The selected mass flow removal rate is input to the 

microprocessor that controls the digital valves. The microprocessor also controls the 

compressor motor speed and the gas temperature at the compressor inlet. 

3.2 Adaptation StrateeieS 

Even though the NASA and the ONERA adaptation streamlining strategies rely on the same 

theoretical principle, the computations are quite different. Both strategies predict the ceiling and floor 

streamlining shapes. These shapes are set in the walls by a system of jacks. We will describe here the most 

important features of each strategy and then we will compare some operating aspects. 

3.2.1 NASA 0.3-m TCT 

Goodyer, Judd, and Wolf (Ref. 5,6,17,20) developed the basic formula of the NASA strategy. The 

formula is related to one isolated flat wall in incompressible flow. The adaptation process is theoretically one 

step. The NASA strategy predicts the wall movement for streamlining by representing the physical wall with a 

vortex sheet. The local vortex intensity, at the streamwise jack position X, is given by the local aerodynamic 

loading: 

where VAX) and V,(X) are the local velocities, parallel to the wall, of the imaginary and real flow field, 

respectively. The strategy considers the vortex sheet as a flat plane (see figure 6) and approximates the 

computation of 7(X) by: 

The strategy computes the velocity normal component to the flat plane, induced by the vortex sheet, by the 

potential flow equation for vortex induced velocity 

This perturbation velocity normal to the wall is balanced by a component of the wall velocity of the real flow 

7 
field. The balancing component is obtained by modifying locally the slope of the wall. This change in slope is 

computed by 

Equation 4 is integrated along the wall to obtain the uerodynumic movement AZ at position E. The 

strategy approximates the theoretical integration extremes -00, +oo with the points X, and X, respectively (Ref. 

20). These are located on both the flexible walls approximately at the beginning and the end of the test section, 

respectively (see figure 5). Finally the strategy interpolates the aerodynamic movement at the jack position X 

from AZ(i$). 

When two walls are considered in the same streamlining process, the simple use of equation 4 for each 

wall is inadequate, because of the strong aerodynamic interaction between the two flexible walls. The strategy 

considers this effect by using a coupling factor CF. This takes into account the effect of one wall on the 

adaptation of the other: 

AZ',(X) = AZ,(X) + CFAG(X) (54 

AZB(X) = AG(X) + CFAZ,(X) (5.b) 

where the subscripts C and F refer to the ceiling and the floor, respectively. Since AZ'(X) tends to overshoot, 

the strategy uses a scaling factor SF to improve the convergence rate. The actual waU/jack movement is 

computed by: 

AZ,(X) = SF (AZ,(X) + CF AG(X)) (64 

AZWF(X) = SF(AG(X) + CFAZc(X)) (6-b) 

Goodyer and Wolf evaluated experimentally the numerical values of both the coupling and the scaling factors to 

give satisfactory convergence (Ref. 5,6). These values are the same for both the ceiling and the floor (CF = 

0.35, SF = 0.8). 

The aerodynamic interaction between the flexible walls also involves a change of flow at the model. 

This change necessitates an iterative procedure. The strategy computes, for each iteration step I 

1) the new wall/jack positions at X for the next iteration by: 

2) the component u of the velocity in the imaginary flow field over the new wall position as an average between 

those creating the vorticity: 

The velocity of the imaginary flow field is proportional to the jack movement. The velocity, computed by 

equation 8, can be corrected by formulas similar to equations 6.a and 6.b to consider the interference of the 

8 
other wall: 

ufC(X)'+' = SF (ufC(X)*+' + CF u,(X)'+') 

uIF(X)'+' = SF (uIF(X)'+' + CF ufc(X)'+') 

The strategy also applies to compressible flows. The compressible effects are considered by scaling the 

measured pressure coefficient to the incompressible value by 

/3 is the Prandtl-Glauert coefficient: (B = dl - MW2). The velocity component u, is then computed from Cpi 

before use in the strategy computation. 

3.2.2 ONERA T-2 

Chevallier, Mignosi et al. (Ref. 7,12,l3) developed the ONERA adaptation strategy. This strategy 

relies on a linearized theory and applies to one isolated, straight wall in incompressible flow. The strategy 

matches the real and the imaginary flow fields along two fictitious, horizontal flat control planes E, and 2,. 

These are chosen near the ceiling and the floor of the test section, respectively. The strategy computes: 

1) The normal (v,) and the tangential (u,) component on each plane CCp These components 

are the projections of the measured velocity on the ceiling and the floor onto the normal and 

parallel directions to C, and C, respectively. 

2) The normal (vf) and the tangential (uf) component on each plane CCp of the imaginary flow 

field. These components are computed by the Inverse Green Transform (IGT) and the Direct 

Green Transform (DGT) of the component u, and vr, respectively 

(1l.a) 

(1l.b) 

The strategy approximates the integration limits 200 with f 100 H (2 3.7xld cm) with respect to the 

turntable center. The strategy breaks down the interval [-00, +00] in three parts: [-a, XJ, [X,, Xu], [Xu, +=I. It 

integrates equations 1l.a and 1l.b in the interval: 

1) [XL, Xu] by using the projections onto C, and CF of the measured velocity on the ceiling and 

the floor, respectively. 

2) [a, X,] and [Xu, +00] by using an estimation of the components u,, v,. 

The strategy estimates the components u,, v, outside the interval [X,, Xu] by representing the 

model by segments. Each segment has a uniform distribution of singularities: vortex, source, 

doublet, and vortex doublet. The strengths of each singularity are adjusted to match the 

velocity measured on the ceiling and the floor at the points X, Xu (Ref. 12). 

9 
The strategy weighs the u and v components in the real and imaginary flow field to get an average value 

up, vp. The components of both the real and imaginary velocity are broken down into four contributions by a 

linear transformation. Unfortunately, the original papers (Ref. 7,U) do not provide enough information to 

understand the meaning of the four contributions. The strategy computes a weighing factor w for each 

contribution and weighs the corresponding velocity. The four contributions of the u component, for instance, 

read 

The inverse transformation is applied to calculate up and vp. The strategy projects up and vp onto the walls to 

create uy, v:. It computes the aerodynamic movement of the wall by the formula: 

for each jack location X. 

The strategy computes the components vy and up' by the potential equation (Ref. 7): 

1) The vy component: 

By integrating the equation: 

p2a+&! -0 

ax az 

with respect to Z, between C and the wall, the strategy obtains: 

where AZc is the actual distance between C and the wall. 

2) The u;" component: 

The irrotational flow field condition: 

can be similarly written as: 

a!=& 

az ax 

10 
The strategy obtains, by integrating equation 17 with respect to Z between C and the walk 

= ufp + V: AZz 

ax 

uf 

The strategy computes the derivative aup/aX and &:/ax as the slope of a least square curve, fitting the 

component up and vp along E,, Both the derivatives are assumed constant along AZc. 

The strategy (Ref. U) computes U, at each iteration step far upstream of the test section by 

integrating on both C, and CF the equation: 

where: 

U(<) is the velocity along the tunnel axis, 

a(<) = l/cosh(lrr//3H) is a weighing function. Q(<) weighs heavily the points near the airfoil and decreases very 

rapidly as < moves away from the airfoil. 

The strategy computes the actual wall movement by also taking into account the BL displacement 

thickness both on the flexible walls (6*C,F) and sidewalls (6*w). This is computed by an undefined integral 

technique (Ref. 7,s). This technique considers the flexible walls like flat plates and uses the measured velocity 

distribution. Finally, the strategy computes the actual jack/wall movement, with respect to the flat wall 

position, by 

The flow chart in figure 7 summarizes the ONERA strategy. 

3.2.3 Comparison of the NASA and ONERA Adaptation Strategies 

The ONERA strategy relies on a more sophisticated and rigorous theory. The strategy considers the 

actual shape of the walls and the numerical values of the integration limits better approximate the theoretical 

values +-. Wolf (Ref. 5) assessed the streamlining errors on the NASA test section, related both to the finite 

length of the integration interval and to the hypothesis of flat vortex sheet, and found them to be very small. A 

conservative estimate of the error, related to the hypothesis of flat vortex sheet, on the velocity component uf is 

given by: 

Auf/U, = -0.5 (ma wall slope)2 (21) 

This leads, even at high lift coefficient (C, = 5) and by using a shallow test section, to an error in estimating the 

interference effect on C, of less than 2 percent. 

The ONERA strategy is much more complex from the computational point of view: 

1) The ONERA strategy achieves numerically the aerodynamic coupling between the ceiling and 

11 
the floor by integrating equations 1l.a and 1l.b at the same time over the planes C, and C, 

The NASA strategy aerodynamically couples the walls by experimental coefficients. 

The ONERA strategy relies, for the computation of the wall/jack movement, on the weighing 

coefficients wj, j = 1,..,4. These coefficients can be very critical for the convergence rate like the 

scaling factors in the NASA strategy. 

The ONERA strategy computes the free stream velocity U,. The NASA strategy operates 

with the measured free stream velocity. The computation of the free stream velocity is, 

indeed, necessary in the ONERA tunnel. The relatively short aerodynamic length of the test 

section does not allow accurate free stream measurements. 

The ONERA strategy computes the boundary layer displacement thickness along the flexible 

walls and the sidewalls. The NASA strategy does not need to compute this parameter. It 

assumes the change in 60CP and 6*,, due to the model, is small (as shown next). 

Both the strategies can start from an cld hoc initial wall shape to accelerate the convergence process. 

The strategies can: i) compute the initial wall shape from a singularity representation of the tested model 

and/or 6) retrieve it from a previous test. The NASA strategy can retrieve the initial wall shape either: 

1) from a data base. The wall shape is selected by assigning the actual test conditions: that is, the 

angle of attack, the Mach number, and the Reynolds number, or 

2) from the aerodynamically straight wall shape. This wall shape considers the effects of the BL 

displacement thickness on the four walls. It achieves a nearly constant velocity distribution 

along the tunnel axis: 

U(X) = u, 

Wolf (Ref. 17,21) determined four aerodynamically straight, contours, each corresponding to the 

following Mach number intervals: up to 0.5, from 0.5 to 0.75, from 0.75 to 0.85, and above 0.85. He did not 

consider the effect of Reynolds number. He found that the effect of Reynolds number on the aerodynamically 

straight wall shape are negligible with respect to the Mach number effects. He fixed intuitively the 

aerodynamically straight contours because the contours obtained experimentally displayed excessive waviness. 

The convergence criterion of the NASA strategy is safer and more efficient. It guarantees a more direct 

evaluation of the wall adaptation quality. The iterative process of the ONERA strategy stops when the number 

of iterations, I, reaches the pre-selected value, N. The progress of the wall adaptation is assessed by checking, 

at successive iterations the distribution of: i) the wall shapes, ii) the Mach number at the walls, and iii) the 

pressure on the model. Archambaud (Ref. 13) says that usually the wall is considered to be adapted if the 

change of the wall position at each jack location is less than 0.01 cm. If this criterion is fulfilled, the other two 

are also satisfied. The adaptation process is repeated to satisfy the convergence criteria. 

The iterations of the NASA strategy stop when the modulus of four errors in wall streamlining are 

simultaneously below prefured values (shown in parenthesis): 

1) the average of the imbalance in the pressure coefficients, at each jack position, between the 

real and imaginary flow field both on the ceiling and the floor (0.01), 

12 
2) the average induced pressure along the model chord (0.007), 

3) the induced camber of the model (0.07 deg), 

4) the induced angle of attack at the model leading edge (0.015 deg). 

Due to the use of hydraulic jacks the overall ONERA adaptation process is much faster than the 

NASA process. One iteration of the ONERA and NASA process typically takes 7 seconds (Ref. 12) and 41 

seconds (Ref. 21), respectively, according to the following table: 

ONERA NASA 

Data Acquisition 5 sec 10 sec 

Wall Setting 0.2 sec 30 sec 

Adaptation Software 2 sec 1 5ec 

The NASA process typically converges in 2-3 iteration steps. The pre-selected number of iterations of the 

ONERA process is usually 45. The effects of these strategies on the model data are discussed in section 6.2.3. 

4. NAE 9x5' WIND TUNNEL 

4.1 Physical Characteristics 

4.1.1 Tunnel Circuit 

The NAE 9x5' tunnel (Ref. 22,23,24,25) is a modern, up-to-date large scale, high speed, blowdown 

tunnel. Both 2-D and 3-D tests can be made over a wide range of Mach numbers: 0.1 c M, c 4.25. Figure 8 

shows the schematic of the tunnel. 

The test gas is filtered, dried air. The air is compressed in a 1428 m3 reservoir at the maximum 

pressure of 21 bars. The tunnel can independently set the Reynolds number and the Mach number. It can, in 

fact, independently control: 1) the temperature, 2) the density, and 3) the velocity. 

1) A matrix of large thermal capacity tubes, set at the exit of the high pressure reservoir keep the 

gas temperature constant during the run. 

2) A valve regulates the airflow from the reservoir into the setting chamber. The valve is 

controlled by a fully automatic electro-hydraulic system. It generates a constant stagnation 

pressure and therefore a constant gas density during the run. The control system can keep the 

stagnation pressure within 0.5 percent of the set value over the entire range of blowing 

pressure (1.2-15 bars). Baffles and screens reduce the noise and the turbulence level in the 

settling chamber. 

3) A 2-D, about 15 m long nozzle accelerates the gas. The ceiling and the floor of the nozzle are 

flexible and are deformed by a system of 22 jacks per wall. This system of jacks is necessary in 

supersonic flow to shape the walls to the best contour for the selected Mach number. 

13 
The adjustment of the velocity in subsonic and transonic 2-D flow is achieved by i) setting the walls of 

the nozzle parallel and ii) changing the ratio of the area of the diffuser second throat and the test section. This 

regulation is made using a servo-controlled hydraulic circuit. This is controlled by a microprocessor, dedicated 

to the control of the Mach number. The accuracy in setting and keeping the Mach number is k0.002 (Ref. 23). 

The intermittent working of the tunnel limits the test time. This is a function of the stagnation 

pressure. We plotted in figure 9 the estimated Reynolds number per meter and the test time (Ref. 22) as a 

function of the stagnation pressure at M, = 0.75. The test time, shown in the figure also takes into account the 

time (about 2 seconds) necessary to establish a steady state flow inside the tunnel. 

4.1.2 Test Section 

The test section in the subsonic (0.1 c M, c 0.70) and supersonic (1.4 c M, c 4.25) regime is formed 

by the downstream end of the flexible nozzle. The ceiling and the floor are almost parallel in that part of the 

nozzle. Tests at transonic speeds (0.7 c M, c 1.4) are achieved in an ad hoc test section. This can easily be 

included in tandem between the flexible nozzle and the variable diffuser. The turbulence level measured in this 

test section at M, = 0.8 is 0.007 (Ref. 25). The turbulence level is defined as the root mean square of the 

pressure coefficient fluctuation. 

The ceiling and floor of the transonic test section are perforated. The geometrical porosity is 21 

percent. A fine mesh covers the perforated walls to eliminate the edge tone, generated by the perforation; the 

actual porosity is about 8.4%. The test section is provided with a passive BLC for 2-D tests. The suction area 

(61 cm x 46 cm) is located in the vicinity of the model and is covered by a porous plate with high flow resistance. 

This allows the system to perform a nearly uniform suction over the whole area. The system operates at a 

chosen suction velocity V, = 0.85% U, (Ref. 24,25). 

The tunnel uses, in the pressure acquisition process, four 48-port scanning valves: 2 for the model and 

1 for the ceiling and the floor. The pressure transducer is a quartz crystal oscillator with a nominal frequency of 

about 40 kHz. The resolution is 0.13 mbar. The transducer is mounted on a temperature controlled, thermally 

isolated base to get the maximum stability. The pressure value is the average of 100 measurements. The wake 

rake has four pressure probes traversing the semispan of the test section. The rake is located at 41 cm 

downstream of the model trailing edge. The rake can translate vertically 51 cm above the tunnel centerline and 

25 cm below (Ref. 24,s). 

A DEC PDP 11/35 computer processes all the test activities (data acquisition, tunnel operation and 

control). A VAX 11/735 computer reduces and displays the data and also runs the wall interference correction 

program. 

We summarize the more important features of the NAE transonic test section in Table 11. 

4.2 Interference Correction 

Mokry and Ohman (Ref. 26,27) developed the method to correct the wall interference in the NAE 

tunnel. This method applies to arbitrary 2-D test sections, including the ventilated test sections. It considers 

only the ceiling and the floor interference effects. The method computes the correction term to the streamwise 

induced velocity (that is, the correction term AM to the Mach number) and to the normal induced velocity (that 

is, the correction term Act to the angle of attack). It relies on the measurement oE 

14 
1) the pressure distribution along the ceiling and the floor, 

2) the model lift coefficient, and 

3) the flow angle at a reference point far from the model. 

The working hypothesis is that the flow field is correctable and sub-critical at the ceiling and floor. The 

relatively high value of the ratio H/c of the model-NAE tunnel system supports this hypothesis. The iiearized 

equation of the potential: 

describes the flow field in the inviscid part of the test section (unshaded area R inside the rectangle in figure 

10). The origin of the axis system (x‘, y’) is located at the quarter-chord of the airfoil. The potential of the 

disturbance velocity is decomposed in two parts: 

where ’ps and p,,, are the potential of the perturbation due to the airfoil in free air and to the wall interference, 

respectively. The method computes ‘ps as the principal part of the Laurent series expansion: 

The first and the second term consider the disturbance of the lift and the displacement effect of the model, 

respectively. No wake effect (modeled by a source term) appears in equation 25. Mokry says this term is 

negligible up to the stall (Ref. 26). The method computes: 

1) The strength of the vortex 7 from the Joukowski theorem: 7 = OScC,, where c is the model 

chord and C, is the lift coefficient. 

2) The strength of the doublet I( from: I( = c2A, where A is the airfoil non-dimensional 

cross-section area. 

As a next step we describe the computation of the wall interference potential ‘pw. 

The wall interference potential fulfills equation 23: 

By using the linear transformation: 

15 
equation 26 is reduced to the Laplace equation: 

By differentiating equation 28 with respect to i?: 

and with the definition: 

the method obtains: 

The method integrates equation 31 by the Fast Fourier Transform. This computation is described in 

references 26 and 27. The boundary conditions along the sides (%’, y2’) - (%’, y2’) and (x,’, y,’) - (5’’ y,’) of the 

rectangle (see figure 10) are computed by: 

r 1 

where F(x‘, y’) is a cubic spline function. It interpolates and smoothes the pressure coefficients measured on 

the ceiling and on the floor. No pressure measurement is available across the stream; the boundary condition 

along the sides (%’, y2)) - (%’, y,’) and (%’, y2)) - (%’, y,’) are computed by interpolating linearly the corner 

values. The streamwise induced velocity and the normal induced velocity are represented by the derivatives 

uw(O,O) = apw(0,O)/ax‘ and vw(O,O) = aCp,(O,O)/ay’ at the airfoil quarter of the chord, respectively. 

4.2.1 Mach Number Correction 

The method (Ref. 26) computes the Mach number correction by differentiating the adiabatic relation 

between the velocity and the Mach number: 

4.2.2 Angle of Attack Correction 

The difficulty to apply the boundary conditions to the normal velocity component complicates the 

computation of vw(O,O) by the direct integration of equation 26. The method computes this component as the 

harmonic conjugate of uw at the point (0,O). Equation 28 can be written as: 

16 
The method computes vw(O,O) by integrating equation 34 with respect to 9' between a reference point (xi, yo') 

and (0,O): 

f0 

The method computes vw(xo',yi), with the definition vw = a(p,/ay', by differentiating equation 24 with respect 

to y': 

If (xi, yi) is inside the linearized flow region: 

where $(xi, yo') is the flow angle at the reference point. This angle is usually measured by a laser velocimeter. 

Mokry suggests $(xi, yo') = 0 if the distance between the airfoil and the upstream side of the rectangle is large 

enough and 

(38) cp(x;* YJ - Cp(xl** YJ - 0 

From equation 36 and 37: 

The correction term to the angle of attack finally reads: 

5. MODEL DESCRIPTION 

We used the aerodynamic measurements on the CAST 10 airfoil to compare, on the operative 

background, the ONERA, NAE, and NASA processes. We show the sketch of the airfoil and the related 

geometrical and aerodynamic design characteristics in figure 11. CAST 10 is an unclassified, German airfoil, 

designed by DORMER for academic purposes and to gain knowledge and experience on the super-critical 

airfoil technology. 

17 
The aerodynamic behavior of the CAST 10 airfoil is typical of the super-critical airfoils. The lift curve 

exhibits a linear part at low angles of attack and a non-linear part (super-lift region) before the tlow breakdown 

region. We show in figure 12 an artistic illustration of the lift curve of a super-critical airfoil. During previous 

tests this airfoil was found to be extremely sensitive to changes in the Mach number, Reynolds number, and to 

the wall interference. Because of its high sensitivity to the wall interference, it is today considered an efficient 

tool to test the flow quality in wind tunnels (Ref. 28). 

5.1 AirfoilModel 

The tests, on which our analysis relies, were made using two different models of the CAST 10 airfoil. 

The models were built at ONERA and NAE, respectively. We will refer to the ONERA and NAE models as 

the French and Canadian models, respectively. 

The French model has a chord of 18 cm. It is made from steel MARVAL 18 and is fitted with 103 

pressure taps. The Canadian model has a chord of 22.9 cm. It is made from steel VASCOMAX 200 and is 

fitted with 80 pressure taps. The pressure taps on the French model are swept along the centerline. Of all the 

orifices on the Canadian model, 67 are staggered along the centerline, the others are located spanwise on each 

side of the centerline. The models were built according to the cryogenic specifications and the manufacturing 

accuracy is very high. The average deviation (Ref. 29) from the design conditions of the upper surface 

coordinates of the Canadian model is Az/c = 0.00058. The average deviation of the coordinates of the upper 

surface between the NAE and ONERA model is Az/c = O.OOO1. The technicians of NASA LaRC checked 

both the models to assure that the aero-thermo-elastic effects did not spoil the model shape. 

The tunnel set up for the two models differs because of the choice of the reference plane of the angle 

of attack. The reference plane of the French model links the leading and the trailing edge. The reference plane 

of the Canadian model links the mid-points between the upper and lower surface at x/c = 0.194 and x/c = 

0.639. The geometrical angle between the two reference planes is 0.882 deg (Ref. 29). 

The models were tested with both free and fued transition on both the upper and lower airfoil 

surfaces. Transition was fuced on the Canadian model by using a carborundum strip located at 5 percent of the 

chord. The average grit size was 0.028 mm. Transition was fued on the French model for the ONERA tests by 

using a carborundum strip located at 6 percent of the chord. The average grit size was 0.045 mm. Transition 

was fued for the NASA tests by using a strip of glass beads located at 5 percent of the chord. The average bead 

size was 0.053 mm. 

We show the geometrical parameters W/c and H/c of the model-wind tunnel systems in Table 111. 

6. ANALYSIS OF THE DATA 

6.1 Evaluation of Secondarv Effects 

The comparison and the validation of the data from the NAE (Ref. 24,25), the NASA (Ref. 30,31,32) 

and the ONERA (Ref. 33,34) tunnels is mainly aimed at an evaluation of both the adaptation processes and 

Mokry's method in alleviating the ceiling and floor interference effects. We preliminarily examined the available 

experimental data. This was necessary to evaluate secondary effects and therefore to remove variables from the 

validation analysis. The nominal operating test conditions of both the preliminary analysis and the following 

18 
validation analysis are: M, = 0.765, R, = 20~10~. We show the actual test conditions on the related figures. 

The normal force coeffcient and the drag coefficient from the NASA, NAE, and ONERA tunnels are 

computed by integrating the pressure distribution along the model surfaces and across the model wake, 

respectively. Further, additional force measurements from a balance are also available for the NAE tunnel. 

The low values of the angles of attack (-3 c a c 3), according to the reference system of the French model, 

allowed us to approximate the lift coefficient with the normal force coefficient (C, M Cn). 

We considered the secondary effects linked to: i) the transition tripping, ;i) the BLC suction velocity, 

and iii) the sidewall interference effects. We also preliminarily checked two sets of the corrected, aerodynamic 

coefficients C, and Cd, from the NAE tunnel (Ref. 24). One set was obtained by the pressure integration. The 

other set was measured by an aerodynamic balance. The comparison showed only slight discrepancies. The 

root mean square deviation, we computed for the test case M, = 0.766, R, = 20.2x106, in the range of the angle 

of attack (-2 c a c 2) was 0.004 for C, and 0.001 for Cd. The plots show the first set of data. 

6.1.1 Transition Tripping 

As expected artificially fming transition showed only negligible effects on integral aerodynamic 

coeficients (C,, Cd), at test Reynolds numbers (R,) on the order of 20~10~. We verified this by comparing the 

lift (see figure 13) and the drag curve (see figure 14) of the Canadian model from the NAE tunnel. The drag 

coefficients were computed from the pressure measured by the pitot of the wake rake, located at the center line 

of the test section. In order to quantify this check we computed the least square linear fit of the two sets of the 

lift coefficients in the range -2 c Q c 2. The curve intercept a and slope b of the two linear fits are the same up 

to the second decimal place (a = 0.33, b = 0.18). 

6.1.2 BLC Suction Velocity 

We found the BLC suction velocity does not strongly affect the lift curve up to the flow breakdown 

region. We verified this by comparing the lift curves of the Canadian model from the NASA tunnel (see figure 

15). Murthy (Ref. 31) measured the C, coefficients at the minimum (V, = 0) and at the maximum (V, = 0.85% 

U,) suction velocity. The parameters of the least square linear fit, in the range -2.5 c a c 1.5, are the same up 

to the second decimal place (a = 035, b = 0.19). 

We plotted the lift coefficients obtained by Mineck (Ref. 30) at V, = 0 in the same figure to get a 

preliminary idea of the repeatability of the data from the NASA tunnel. The check turned out to be satisfactory 

up to the flow break region (a c 1.5). We found only a slight discrepancy in the slope of the curve. The slope of 

the least square fit of Mineck’s data is: b = 0.18. The drag curve (see figure 16) also shows a reasonable 

repeatability at low and moderate angles of attack (a c 1.5). 

6.1.3 Sidewall Interference Effect 

Stanewsky (Ref. 28) found that the sidewall interference reduces the magnitude of the non-linearity 

feature of the lift curve in ventilated test sections. The reduction in AC, is inversely proportional to the ratio 

W/c. We assumed Stanewsys remark is true also for solid wall test sections. 

We show in figure 17 the lift curves of the French (Ref. 32) and the Canadian (Ref. 30) model from the 

NASA tunnel. These lift curves correspond to the maximum (1.8) and the minimum (1.4) ratio W/c of the 

available experimental data. As expected 

19 
1) The low values of the ratio W/c strongly affect the super-lift feature of the two models. 

2) The reduction in AC, is almost the same for the two models. The ratio W/c is almost the same 

for the two model-wind tunnel systems. 

We quantified these effects by comparing, once again, the parameters of the least square linear fit. 

The reduction in AC,, that is, the linearization of the lift curve, is quantified by the correlation coefficient r. 

This parameter is a measure of the fitting quality of the curve. It ranges in the interval 0 < r < 1. The 

correlation coefficients for the the lift curve of the French and Canadian model in the range -1.09 < CY e 1.85 

(according to the reference system of the French model) are 0.98 and 0.83, respectively. The intercept and the 

slope also are almost the same for the models a = 0.33, b = 0.17 for the French model and a = 0.35, b = 0.18 

for the Canadian model. 

62 Validation a nd ComDarisoq 

6.2.1 General Remarks 

The validation analysis is the technical tool to assess the flow quality in wind tunnels. The data are 

usually validated either from the comparison with interference free data, or from the evaluation of the residual 

interference. 

The validation of the NASA, NAE, and ONERA data was particularly hard because we could not: 

use experimental data from other tunnels as reference data. The tunnels of our concern are 

supposed to be among the best in the western world. The data of these tunnels are often used 

to validate the data from other tunnels. 

compute by a common technique and then compare the residual interference. We considered 

the possibility of using the numerical program TWINTN4 WIAC (Ref. 35) to compute the 

residual interference in the three tunnels. The code is quite general and can be used both for 

flexible and ventilated walls test sections. The use of this program on a routine basis is 

somewhat difficult. Its use requires user intuition and intervention as well as substantial 

computing time. 

We validated the experimental data from the comparison with the theoretical ones. These are 

computed by the most recent version (MCMJ-11) of the GRUMFOIL code (Ref. 36,37). Bearing in mind the 

limitations involved by any numerical approach, the analysis gives only some indications of the experimental 

data quality. 

6.2.2 GRUMFOIL Code 

GRUMFOIL is a computer program for calculating the steady, Viscous, transonic flow over 

two-dimensional airfoils. The solution relies on an iterative scheme of the strong interaction between the 

boundary layer and the inviscid flow field. The boundary layer is computed by integral entrainment methods. 

The inviscid flow field is computed by integrating, by a multigrid method, the full non-linear potential equations. 

The use of GRUMFOIL for the present application suffers from the limitation that the code does not 

adequately deal with large separated flow regions. In this condition the iteration scheme loses some 

20 
effectiveness. The solution, when possible, can be obtained only by using very low relaxation factors with a 

remarkable increase in computer time. This shortcoming restricts the use of the code for small/moderate 

angles of attack and/or Mach numbers. 

We run the code in the inverse mode such that the angle of attack is computed from the input lift 

coefficient. We always: i) input the code with the NASA test conditions: Mach number, Reynolds number, and 

lift coefficient, and ii) fixed the transition location between x/c = 0.002 and x/c = 0.05. 

6.2.3 Results of the Validation Process 

We made the validation analysis by comparing with the theoretical results: 

1) local quantities such as the pressure distribution and the shock location on the model surface, 

2) (as suggested by McCroskey (Ref. 38)), integral quantities such as the drag curve as a function 

of the Mach number, and the slope of the linear part of the lift curve as a function of the 

Reynolds number and the Mach number. 

We also compared: 1) the NASA and NAE maximum lift coefficient as a function of the Reynolds 

number and Mach number, 2) the two-dimensionality of the flow in the test section of the NASA and NAE 

tunnel. Unfortunately, the above limitation of the GRUMFOIL code prohibits its effectiveness in determining 

Clmax. The sparsity of the available data did not allow us to do the analysis on Clmax for the ONERA data. The 

intermittent working of the ONERA tunnel makes it more difficult to do systematic tests in terms of Mach 

number, Reynolds number, and angle of attack. 

6.2.4 Pressure Distribution 

We show the pressure distribution on the French and Canadian model in figures 18 and 19, 

respectively. The agreement of the pressure coefficient of the NASA and NAE tests with the GRUMFOIL 

results is excellent. We can estimate the agreement of the NASA, ONERA, and GRUMFOIL pressure 

distribution reasonably good. Note that the NASA and the ONERA data were obtained with fured and free 

transition, respectively. The mismatching of this test condition gives rise to the small discrepancy in the C 

coefficients on the model upper surface, even though the test Reynolds number is high enough (Rc w %lo6). 

The discrepancy starts from x/c w 0.10 and increases constantly up to x/c w 0.70. The tripping anticipates the 

increase of the displacement thickness. This gives rise, in turn, to an increase of the local velocity and therefore 

to a decrease of the local pressure. 

P 

We compared the pressure distributions on the French model also at other test conditions (see figure 

20). As expected the better matching of the test conditions (both sets of data were obtained at free transition) 

involves a better agreement of the pressure distribution. The agreement is even more surprising because of the 

severity of the test conditions. The lift coefficient is very close to the maximum value (Clmax w 0.655, see figures 

13,15 and 17). It is well known that the flow field is very sensitive to the wall interference near stall conditions. 

The severity of the test conditions discouraged us to attempt to run the GRUMFOIL code. 

6.2.5 Shock Location 

We used the classical criterion to determine the abscissa XJC of the shock wave on the model upper 

surface. This criterion locates the shock at the point where the slope of the pressure profile changes sign. We 

21 
computed this location as the abscissa of the cross-point of the curves, interpolating the pressure coefficients 

before and after the suction peak. Figure 21 explains graphically this criterion. We used the described criterion 

to locate the shock location from the GRUMFOIL pressure distribution. This was necessary to be consistent 

with the shock location from the experimental data. 

We show in figures 22 and 23 the shock location as a function of the lift coefficient on the Canadian 

and French model, respectively. The agreement of both the experimental data and the GRUMFOIL results is 

excellent for the Canadian model. The root mean square deviation is 0.022 and 0.027 for the NASA and NAE 

data, respectively. The trends of the NASA and NAE data are also consistent with the stall. Because of the 

reason stated before, we could expect a disagreement of the shock location at higher lift coefficient. The 

agreement of the NASA data and the GRUMFOIL results is satisfactory for the French model. The root mean 

square deviation is 0.031. The trend of the NASA data is also consistent with the stall. The sparsity of the 

available data does not allow us to evaluate the ONERA data. 

6.2.6 Drag Curve 

We found the agreement between the NASA and NAE drag coefficients of the Canadian model is 

excellent over the whole range of Mach number, 0.70 c M, c 0.79 (see figure 24). Both sets of data show: 

1) practically &e same drag divergence Mach number. We could roughly include this value in 

- the range: 0.765~ M, c 0.775. 

2) the drag creep before the drag divergence. The drag creep is a typical phenomenon of some 

super-critical airfoils (Ref. 39). It is generated by a complicated interaction of several causes 

and is strongly related to the boundary layer condition and the airfoil fluid shape. The 

agreement of the experimental and theoretical results is also satisfactory up to M, = 0.765. 

The NASA drag coefficients of the French model are in good agreement with the theoretical results 

(see figure u). Even though the numerical results are systematically higher than the experimental data, the 

curves show the same trend over the whole range of the computed results (0.70 c M, c 0.77). The agreement 

of the ONERA drag coefficients with both the NASA and theoretical results is also reasonably good up to M, EJ 

0.73. The ONERA data are lower than the NASA data at higher Mach number (M, = 0.762, M, = 0.765). A 

shock weakening, due to a local decrease of the Mach number, could have produced the drop of the drag 

coefficients. The Mach number decrease, in turn, was probably due to stronger residual wall interferences. 

Unlike the drag curve of the Canadian model, both the NASA and ONERA data show no drag creep in the 

drag curve of the French model. The different test Reynolds number could explain the absence of this 

phenomenon. 

6.2.7 Lift Curve Slope 

We computed the slope of the linear part of the lift curve (Cia) by the finite difference of two adjacent 

lift coefficients: 

We show the NASA data with the experimental tolerance band. We computed the tolerance band from the 

22 
measurement accuracies of 20.005 for the lift coefficient and 20.01 deg for the angle of attack (Ref. 9). 

We show the slope of the linear part of the lift curve for the Canadian model as a function of the Mach 

number and Reynolds number, in figures 26 and 27, respectively. To show quantitatively the benefits of the wall 

streamlining in reducing the wall interference effects, we plotted the uncorrected Canadian data in the same 

figure. The agreement of the NASA and NAE data and the theoretical results are reasonably good on the 

whole interval of the Mach number. Both the Canadian data and the theoretical results practically lie within the 

tolerance band of the NASA data. The adding of numerical approximation error in computing the finite 

difference slope could explain the slight wavering of the theoretical results. The agreement of the NASA and 

NAE data is excellent on the whole interval of the Reynolds number (see figure 27). We show in the figure only 

the GRUMFOIL results computed at R, = 6x106 and R, = 20~10~. The lift curve slope computed at R, = 

10x106 and R, = 15x106 turned out to be physically unacceptable, even though the runs were successful from 

the numerical point of view. The runs showed, in fact, good convergence process and the residual error 

dropped four orders of magnitude. A possible cause of convergence toward a wrong solution, at the selected 

test conditions (C,, R,, M,), could be due to the inability of the code to match the potential and viscous solution 

because of the presence of a large pressure gradient close to the leading edge. 

We show the lift curve slope of the French model as a function of the Reynolds number on figure 28. 

The agreement of the NASA and the theoretical results are satisfactory in the trend. The discrepancy in C,, 

arises because the GRUMFOIL code predicts the super-lift phenomenon, while the experiments, as shown 

before (see figure 17), do not show this feature. The experimental lift coefficients C,, (C,, RI 0.4) used to 

compute C,, from equation 41 are beyond the linear range for the GRUMFOIL code. This is easily verified 

from figure 29, where we show a typical experimental lift curve of the CAST 10 airfoil from the NASA tunnel 

and the corresponding theoretical curve. We shifted the theoretical curve along the a-axis to match the angle of 

attack at C, = -0.019. The sparsity of the available data does not allow us to evaluate the ONERA data. 

Maximum Lift Coefficient and Two-Dimensionality of the Flow Field 

We assessed once again: 

the good agreement of the NASA and NAE data by comparing Clmax as a function of the 

Reynolds number (see figure 30) and the Mach number (see figure 31). The two sets of data 

show similar trends if we do not consider the anomalous lift coefficient at M, = 0.8 from the 

NASA tunnel. The difference is perhaps due to the different operations of BLC. The NAE 

data were obtained with the suction velocity V, = 0.85% U, while the NASA data were 

obtained without BLC suction (V, = 0). The matching of the results is even more surprising 

because the residual interference effects are opposite in sign in ventilated and in solid test 

section. The wall interference, in fact, increases the maximum lift coefficient in ventilated test 

section (Ref. 28) and decreases the maximum lift coefficient in solid wall test section (Ref. 21). 

The wall streamlining improves the two-dimensional flow in the test section. We easily 

verified this by comparing the profile of the NASA and NAE drag coefficient along the 

transverse dimension of the test section. The drag profiles were obtained at C, N 0.3 (see 

figure 32) and C, N 0.6 (see figure 33), respectively. The NASA data are even more remarkable 

because: i) what obtained without BLC (V, = 0), and ii) the wake rake of the NASA tunnel 

was located nearer to the trailing edge of the model (27.4 cm) (Ref. 30) than the wake rake of 

the Canadian tunnel (41 cm). 

23 
7. CONCLUDING REMARKS 

We have compared critically three of the best 2-D, transonic wind tunnels of the western world i) 

NASA 03m TCT which is fitted with an AWTS, ii) ONERA T-2 which is also fitted with an AWTS, and iii) 

NAE 5'16' which has a conventional, ventilated test section. We have discussed 

1) The general, technical features of the tunnel circuits and, in more detail, the operational and 

performance aspects of the test sections. 

2) The correction techniques: two wall adaptation strategies and one classical correction method 

(Mokry's method). 

3) Various sets of airfoil data. We validated these data by the comparison with the theoretical 

results of the GRUMFOIL code. 

The data were obtained on two models of the CAST 10-2/DOA-2 supercritical airfoil. The high 

sensitivity to the changes of the Mach number and of the Reynolds number and, moreover, to the wall 

interference, makes this airfoil a useful tool to evaluate the flow quality in a wind tunnel. 

We found the NASA and the NAE tunnel to have the same data accuracy. However, the NASA tunnel 

appears to have a better ratio of benefits to costs because: 

1) the smaller dimensions of both the circuit and the tests section involve less initial investment, 

maintenance, and operation costs. 

2) The operation of a cryogenic wind tunnel is cheaper and more effective for independent 

control of the Mach number and the Reynolds number. 

3) The AWTS: i) allows immediate interference free measurements, without needing to use any 

post-test correction method and ii) improves the general flow characteristics in the test section 

as, for instance, the two-dimensionality of the flow field. 

The above mentioned remarks are, from a general point of view, true also for the ONERA tunnel. The 

ONERA tunnel suffers, however, from the shortcoming of intermittent operation. This feature strongly limits 

the tunnel capability in performing systematic tests at severe flow conditions. This limitation did not allow us 

to: i) definitively validate the results because of a sparsity of available data and ii) obtain a precise insight into 

the capability of the wall adaptation process. In addition, the ONERA tunnel is not able to test at such high 

Reynolds numbers as routinely set in the NASA tunnel. 

We can affirm, on the basis of the objective considerations in this report, that the NASA tunnel is more 

powerful and effective for 2-D, transonic tests up to full scale Reynolds numbers. 

24 
8. REFERENCES 

1. Martindale, W. R.; Butler, R. W.; et al.: Study on Needs for a Magnetic Suspension System Operating with 

a Transonic Wind Tunnel. NASA CR 3900, May 1985. 

2. Sears, W. R.; and Erickson, J. C.: Adaptive Wind Tunnel. Ann. Rev. Fluid Mech., Vol. 20,1988. 

3. Wolf, S. W. D.; and Kilgore, R. A.: Adaptive Wall Test Section. Lecture Series on: "Advanced Experimental 

Techniques for Transonic Wind Tunnel" presented at The National Defense Academy Department of 

Aeronautical Engineering, Yokosuka, Oct. 1987. 

4. Ashill, P. R.; and Weeks, D. J.: A Method for Determining Wall Interference Corrections in Solid Wall 

Tunnel from Measurements of Static Pressure at the Walls. AGARD Conference Proceedings on: "Wall 

Interference in Wind Tunnel", N.335, Paper N.l, London, May 1982. 

5. Wolf, S. W. D.: The Design and Operational Development of Self Streamlining Two-Dimensional Flexible 

Walled Test Sections; NASA CR 172328, Mar. 1984. 

6. Wolf, S. W. D.; and Goodyer, M. J.: Predictive Strategy for Two-Dimensional Flexible Walled Adaptive 

Wind Tunnel. A detailed Description of the First One-Step Method. AASU Memo 85/12, Jan. 1986. 

7. Chevallier, J. P.; and Mignosi, A.; et al.: T-2 Wind Tunnel Adaptive Walk Design, Construction and Some 

Typical Results. La Recherche Aerospatiale (English Edition), N.4, Jul./Aug. 1983, pp 1-19. 

8. Ladson, C. L.; and Ray, E. J.: Evolution and Operation Characteristics of Two-Dimensional Test Section of 

the Langley 03-m Transonic Cryogenic Tunnel. NASA TP 2749, Sept. 1987. 

9. Design Library of the 0.3-m TCT Test Section, Experimental Techniques Branch, NASA LaRC, Hampton 

(VA) 

10. Gobert, J. L.: Data Acquisition and Process Control in The Self-correcting Cryogenic Wind-Tunnel T-2 at 

ONERA/CERT by Integration of Two Minicomputers. 11th International Congress on: "Instrumentation 

in Aerospace Simulation Facilities", Stanford, Aug. 1985. 

11. Michel, R.; and Mignosi, A.: Adaptation and First Cryogenic Operation of T-2 ONERA/CERT Wind 

Tunnel. La Recherche Aerospatiale (English Edition), N.2, Mar./Apr. 1982, pp 75-85. 

12. Archambaund, J. P.; Blanchard, A.; et al.: Instrumentation and Testing Techniques in T-2 Transonic 

Cryogenic Wind Tunnel at ONERA/CERT. 11th International Congress on: "Instrumentation in 

Aerospace Simulation Facilities", Stanford, Aug. 1985. 

13. Archambaud, J. P.; and Mignosi, A.: Two-Dimensional and Three-Dimensional Adaptation at the T-2 

Transonic Wind Tunnel of ONERA/CERT. AIAA 15th Aerodynamic Testing Conference, San Diego, May 

1988. 

14. Rallo, R. A.; Dress, D. A.; et al.: Operation Envelope Chart for The Langley 0.3-m Transonic Cryogenic 

Wind Tunnel. NASA TM 89008, Aug. 1986. 

25 
15. Dor, J. B.; Mignosi, A.; et al.: Qualification of the T-2 Wind Tunnel in Cryogenic Operation. B: Flow 

Fluctuation - Particle Detection and Qualification. NASA TM 77782, Dec. 1984. 

16. Igoe, W. B.: Private Communication 

17. Wolf, S. W. D.: Wall Adjustment Strategy Software for Use With the NASA Langley 0.3-m Transonic 

Cryogenic Tunnel Adaptive Test Section. NASA CR 181694, Nov. 1988. 

18. Wolf, S. W. D.: Private Discussion 

19. Johnson, C. B.; Murthy, A. V.; et al.: A Description of the Active and Passive Side Wall Boundary Layer 

Removal System of the 0.3-m Transonic Cryogenic Tunnel. NASA TM 87764, Nov. 1986. 

20. Wolf, S. W. D.; and Goodyer, M. J.: Predictive Wall Adjustment Strategy for Two-Dimensional Flexible 

Walled Adaptive Wind Tunnel - A Detailed Description of the First One-Step Method. NASA CR 181635, 

Jan. 1988. 

21. Wolf, S. W. D.: Evaluation of a Flexible Wall Testing Technique to Minimize Wall Interference in the 

NASA Langley 0.3-m Transonic Cryogenic Tunnel. AIAA 26th Aerospace Science Meeting, Reno, Jan. 

1988. 

22. Brown, D.: Information for Users of the National Research Council’s S’xY Blowdown Wind Tunnel at the 

National Aeronautical Establishment. Laboratory Technical Report HA-6, Ottawa, Sept. 1977. 

23. Ohman, L. H.; Brown, D.; et al.: Recent Improvements to the NAE 5’s Blowdown Wind Tunnel. NAE. 

Aeronautical Note N. 31, Ottawa, Aug. 1985. 

24. Chan, Y. Y.: Wind Tunnel Investigation of CAST 10-2/DOA-2 12% Super-critical Airfoil Model. 

Laboratory Technical Report HA-162, Ottawa, May 1986. 

25. Chan, Y. Y.: Wind Tunnel Investigation of CAST 10-2/DOA-2 12% Super-critical Airfoil Model, Phase 11. 

Laboratory Technical Report HA-170, Ottawa, June 1987. 

26. Mokry, M.; and Ohman, L. H.: Application of the Fast Fourier Transform to Two-Dimensional Wind 

Tunnel Wall Interference. J. Aircraft, Vol. 17, N. 6, June 1980, pp 402-408. 

27. Mokry, M.; and Chan, Y. Y.: Wall Interference Correction from Boundary Measurements. AGARD AG N. 

281 on Two-Dimensional Wind Tunnel Interference. Nov. 1983. 

28. Stanewsky, E.; Demurie, F.; et al.: High Reynolds Number Tests of the Cast 10-2/DOA-2 Transonic Airfoil 

at Ambient and Cryogenic Temperature Conditions. AGARD Conference Proceedings on Wind Tunnels 

and Testing Techniques. N. 348, Paper N. 10, Cesme, Sept. 1983. 

29Arcara, P. C.: CAST 10-2/DOA-2 Airfoil Models Tested in 0.3-m Transonic Cryogenic Tunnel. NASA 

Unpublished Report. 

30. Mineck, R. E.: Wall Interference Tests of a CAST 10-2/DOA-2 Airfoil in an Adaptive Wall Test Section ; 

NASA TM 4015, Dec. 1987. 

26 
31.Murthy, A. V.: Private Communication. 

32. Jenkins, R. V.: Private Communication. 

33. Mignosi, A.; Saraudie, A.; et al.: Premier Rapport d’Essais du Profd CAST 10 (Chorde = 180 mm) en 

Transition Naturelle, Effectues dans la Souffleire Transonique Cryogenique T-2 en Presence de Parois 

Auto-Adaptables ; ONERA/CERT R.T. OA N. 59/1685, Mars 1985. 

34. Seraudie, A.; Blanchard, A.; et al.: Rapport d’Essak du Profil CAST 10 en Transition Declenchee, 

Effectues dans la Souffleire Transonique Cryogenique T-2 en Presence de Parois Auto-Adaptables ; 

ONERA/CERT, R.T. OA N. 63/1685, Aout 1985. 

35. Green, L. L.; and Newman, P. A.: Transonic Wall Interference and Correction for Airfoil Data from the 

0.3-m TCT Adaptive Wall Test Section ; AIAA 19th Fluid-dynamics, Plasma Dynamics Conference, 

Honolulu, June 1987. 

36. Melnik, R. E.; Chow, R. R.; et al.: An Improved Viscid/Inviscid Interaction Procedure for Transonic Flow 

over Airfoils ; NASA CR 3805, Oct. 1985. 

37. Mead, H. R.; and Melnick, R. E.: GRUMFOIL A Computer Code for the Viscous Transonic Flow over 

Airfoil ; NASA CR 3806, Oct. 1985. 

38. McCroskey, W. J.: A Critical Assessment of the Wind Tunnel Results for the NACA 0012 Airfoil ; 

AGARD FDP Symposium on Aerodynamic Data Accuracy and Quality: Requirement and Capabilities in 

Wind Tunnel Testing, Paper N.1, Naples, Sept. 1987. 

39. Johnson, W. G.; Hill, A. S.; at al.: High Reynolds Number Tests of a BOEING BAC I Airfoil in the 

Langley 0.3-Meter Transonic Cryogenic Tunnel ; NASA TM 81922, Apr. 1982. 

9. ACKNOWLEDGMENT 

Although the conclusions presented in this paper were derived from an objective review by the author, 

the preparation of the basic information represented a team effort. My appreciation for this assistance goes to 

Mr. Clyde R. Gumbert of the Theoretical Aerodynamic Branch, and the Langley 0.3-m TCT Research Group. 

Hampton, Va. 

August 7,1989 

27 
$. 

tSET UP TUNNEL AND MODEL1 

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[CONDITIONS 

I I 

MEASURE TWO INDEPENDENT FLOW 

VARIABLES pr, qr ON IvSgt 

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BOUNDARY CONDITION q, ON tvSvt I 

BOUNDARY "S" FOR STEAMLINING + ADJUST S 

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Figure 1. - Flow chart of wall adaptation principle. 

28 
Imaginary fbwfiekls’ 

extending to infinity 

If 

Figure 2. - Artistic illustration of 2-D wall adaptation principle (from Ref. 5). 

29 
NASA 0.3-m TCT 

ONERA T-2 

Figure 3. - Schematics of NASA 0.3-m TCT and ONERA T-2. 

30 

2 R I0762 m) -I 
x 106 

360 

240 

Rm 

120 

NASA 0.3-m fCT 

0 

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Moo 

300 

240 

180 

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t loo 

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Figure 4. - Operating envelopes of NASA 0.3-m TCT and ONERA T-2 (from Ref. 11) at 

several stagnation pressures and temperatures. 

31 
NASA 0.3-m TCT 

THROAT ONERA T-2 

Figure 5. - Schematics of test sections of NASA 0.3-m TCT and ONERA T-2. 

32 
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34 
35 
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37 
Geometrical characteristics: 

Max. percentage thickness = 0.121 at x/c = 0.45 

Trailing edge percentage thickness = 0.005 

Aerodynamic design characteristics: 

M, = 0.76, C1 = 0.595, a = 0.3 deg. 

Figure 11. - Sketch of CAST 10-2/DOA-2 airfoil. 

38 
Flow Break r 

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61 
TABLE I1 

PHYSICAL AND OPERATING CHARACTERISTICS OF THE 2-D, TRANSONIC 

TEST SECTION OF THE NAE 5Ix5' WIND TUNNEL 

(All lengths are in cm) 

Test section walls: 2 ventilated (ceiling and floor) 

Test section dimensions (LxHxW): 358x152~38 

Ceiling and floor porosity: 8.4% 

Test gas: Filtered and dried air 

Turbulence level: 0.007 (M, = 0.8) 

Max. stagnation pressure: 5 bars 

Pressure transducer: Quartz crystal oscillator 

Pressure transducer resolution: 0.13 mbar 

2 solid (side) 

Mach number: 0.7 to 1.4 

Mach number accuracy: k0.002 

Reynolds number per meter: 1.6~10~ to 6.8~10~ (M, = 0.75) 

Test time: 47 to 4 sec (M, = 0.75) 

. 

62 
TABLE I11 

GEOMETRICAL PARAMETERS OF THE MODEL-WIND TUNNEL SYSTEMS 

w/c 

H/c 

French 

Model 

c = 18 cm 

NASA 

1.8 

1.8 

Canadian 

Model 

c = 23 cm 

ONERA 

2.2 

2.1 

NAE 

// 

// 

W/C 1 1.4 

H/c 1.4 // 6.7 

63 
Report Documentation Page 

1. Report No. 2. Government Accession No. 

NASA CR- 181871 

4. Title and Subtitle 

An Evaluation of Three Experimental Processes for TwoDimensional Transonic Tests 

7. Author(s1 

Genna ro Zuppa rd i 

3. Recipient's Catalog No. 

5. Report Date 

August 1989 I 

6. Performing Organization Code 

8. Performing Organization Report No. 

10. Work Unit No. 

Aversa (CE) Italy 

9. Performing Organization Name and Address 

Gennaro Zuppardi 

U. of Naples 

Via Ettore Corcione, 44 Aversa 

1 13. Type of Report and Period Covered 

11. Contract or Grant No. 

NCC1-47 

JIAFS Core Program 

.. - 

Contractor Report 12. Sponsoring Agency Name and Address 

National Aeronautics and SDace Administration 

19. Security Classif. (of this report) 20. Security Classif. (of this page) 21. No. of pages 

Uncl ass i f i ed Uncl assi f i ed 70 

Langley Research Center 

Hampton, VA 23665-5225 

22. Price 

A04 

I 14. Sponsoring hgency Code 

I 

15. Supplementary Notes 

Langley Technical Monitor: Edward J. Ray 

16. Abstract 

The aerodynamic measurements in conventional wind tunnels usually suffer from 

the interference effects of the sting supporting the model and the test section 

walls. These effects are particularly severe in the transonic regime. Sting 

interference effects can be overcome through the Magnetic Suspension technique. 

effects can be alleviated by; testing airfoils in conventional, ventilated tunnels 

at relatively small model to tunnel size ratios, treatment of the tunnel wall 

boundary layers,or by uti1 ization of the Adaptive Wall Test Section (AWTS) concept. 

This paper assesses the operating capabilities and results from two of the foremost 

two-dimensional, transonic, AWTS facilities in existence. These facilities are the 

NASA 0.3-Meter Transonic Cryogenic Tunnel and the ONERA T-2 facility located in 

Toulouse, France. In addition, this paper will evaluate the results derived from 

the well known conventional facility, the NAE 5'X5' Canadian wind tunnel. CASTlC!/DOA 

Airfoil results will be used in all of the evaluations. 

Wal 

17. Key Words (Suggested by Authork)) I 18. Distribution Statement 

Adaptive Wall Test Section 

Boundary Layer Unclassified - Unlimited