NASA Technical Memorandum 4762

SE

Wing Leading Edge

Joint Laminar Flow Tests

Aaron Drake, Russell V. Westphal, Fanny A. Zuniga, Robert A. Kennelly, Jr.,

and Dennis J. Koga

October 1996

National Aeronautics and

Space Administration
NASA Technical Memorandum 4762

Wing Leading Edge

Joint Laminar Flow Tests

Aaron Drake, Russell V. Westphal

Washington State University, Tri-Cities, Richland, Washington

Fanny A. Zuniga, Robert A. Kennelly, Jr., Dennis J. Koga

Ames Research Center, Moffett Field, California

October 1996

National Aeronautics and

Space Administration

Ames Research Center

Moffett Field, California 94035-1000
Wing Leading Edge Joint Laminar Flow Tests

AARON DRAKE,* RUSSELL V. WESTPHAL,* FANNY A. ZUNIGA, ROBERT A. KENNELLY, JR.

and DENNIS J. KOGA

Ames Research Center

Summary

An F-104G aircraft at NASA's Dryden Flight Research

Center has been equipped with a specially designed

and instrumented test fixture to simulate surface imperfections of the type likely to be present near the leading

edge on the wings of some laminar flow aircraft. The

simulated imperfections consisted of five combinations

of spanwise steps and gaps of various sizes. The

unswept fixture yielded a pressure distribution similar

to that of some laminar flow airfoils. The experiment

was conducted at cruise conditions typical for business

jets and light transports: Mach numbers were in the

range 0.5-0.8, and unit Reynolds numbers were 1.5-2.5

million per foot. Skin friction measurements indicated

that laminar flow was often maintained for some distance downstream of the surface imperfections. Further

work is needed to more precisely define transition location and to extend the experiments to swept-wing

conditions and a broader range of imperfection

geometries.

Nomenclature

a

c

Cf

c

f 1,2,4

Cp

FTP

HLFC

Moo

local sound speed (ft/s)

chord (ft)

local skin friction coefficient

(T/(0.5 *p *U2))

skin friction coefficient at gauges 1, 2

and 4, respectively

coefficient of pressure (

flight test fixture

Stanton gauge height (in)

altitude (ft)

hybrid laminar flow control

free-stream Mach number

NLF

P

POO

Po

MOO

Re/ft

U

Uc«

X

Mr

a

P

AP

Uoo

P

POO

T

Motivation

natural laminar flow

local static pressure (lb/ft2)

free-stream static pressure (lb/

total pressure (lb/ft2)

free-stream dynamic pressure (lb/ft2)

unit Reynolds number

(poo* Uoo /(Joe )

local velocity outside boundary layer

(ft/s)

free-stream velocity (ft/s)

chordwise distance from nose leading

edge (in)

chordwise distance from leading edge

to transition location (in)

spanwise distance from FTP root (in)

angle-of-attack (degrees)

side-slip angle (degrees)

Stanton gauge pressure difference

(lb/ft2)

freestream viscosity (si/ft *s)

local density (si/ ft^)

freestream density (si/ ft^)

skin friction (lb/ft2)

Washington State University, Tri-Cities,

Richland, WA 99352-1643

The competitive pressure to build more energy-efficient

aircraft has generated a considerable amount of

research aimed at achieving natural laminar flow (NLF)

and hybrid laminar flow control (HLFC) on transonic

aircraft wings to reduce skin friction drag. This

research has focused primarily on airfoil design and

understanding transition behavior with less concern for

the surface imperfections and manufacturing variations

inherent to production aircraft (ref. 1). Well-defined
criteria in such areas as allowable sizes of gaps and

steps that result from aircraft skin seams, as well as criteria for rivet and fastener sizes are essential if

NLF/HLFC technology is to be incorporated on massproduced aircraft.

Most business jet wings are constructed with a spanwise joint between the leading-edge-skin piece and the

main-wing skin approximately four to six inches downstream of the leading edge (see fig. 1). The leadingedge-skin piece on business jets is typically polished

aluminum and is intended to be removable for

(infrequent) structural inspections. A major obstacle to

constructing a wing which will achieve laminar flow is

the attachment of the leading-edge piece (see fig. 2(a)).

A proposed method which is very similar to current

business jet practices, involves milling out a strip on the

leading edge adjacent to the joint where the screws

would attach. This would then be covered by a plastic

or aluminum strip held in place by adhesive. This strip

would cover the joint gap and be nearly flush with the

wing surface (see fig. 2(b)). The cover strip could

easily be trimmed to minimize the resulting gaps and

step misalignment.

Before the proposed construction method can be used,

it is necessary to have an understanding of the allowable size of steps and gaps that are needed to permit

laminar flow. This experiment begins to answer the

question: how small do steps and gaps in the surface of

business jet type aircraft wings need to be to allow laminar flow? To do this, a simulated wing leading edge

was constructed that would allow configurations with

various size steps and gaps to be tested in the presence

of a wing-like pressure distribution.

Experiment

To address concerns over scale effects and maximize

the applicability of the results obtained, this experiment

was conducted in flight using a full scale test article. A

facility ideally suited to such experimentation is the

Dryden Flight Research Facility F-104 Flight Test

Fixture (ref. 2). The Flight Test Fixture (FTP) is a low

aspect ratio ventral fin mounted on an F-104G aircraft

(see fig. 3). The FTP has a span of 2 ft and an approximate chord of 7 ft. The size of the FTP and the operational envelope of the F-104 allow full scale testing

directly applicable to business jets and light transports.

Furthermore, the flight envelope of the F-104/FTF

allows speed, altitude, Mach number, Reynolds number, and any atmospheric effects to be duplicated precisely. However, the key advantage to conducting this

experiment in flight, as opposed to ground-based testing using a wind tunnel, is the ability to precisely match

the disturbance environment of the test to that which

would be encountered in practice.

In order to perform NLF experiments using the FTP, it

is necessary to have a noseshape whose pressure distribution is similar to that of an NLF airfoil. In this way,

the flow over the wing leading-edge region of the aircraft can be simulated on the noseshape of the FTP.

Since no suitable noseshape existed, a replacement was

designed (see fig. 4). The primary goal of this design

was to provide a pressure distribution similar to that

found on some recently designed laminar flow airfoils

(ref. 3). These airfoils are characterized by smoothly

accelerating pressure gradients over the upstream 25%

of chord. Additionally, it was desirable to minimize the

sensitivity of the design to small changes in angle of

attack. Angle of attack on the FTP corresponds to aircraft sideslip and cannot be assumed to be precisely

constant in flight. For this design, it was assumed that

sideslip could be controlled to within ±0.5°. To allow

an appropriate error margin, the pressure distribution on

the fixture was to remain well-behaved at up to 1.5°

angle of attack. The flow about the final shape, denoted

TFN-7525 (see table 1), was analyzed using the twodimensional full-potential CFD code for a range of

conditions (ref. 4). The results showed that TFN-7525

met all the design goals. A favorable pressure gradient

is maintained over the first 80% of the noseshape and

remains favorable throughout the range of angle-ofattack and Mach number examined. Figures 5 and 6

show and tabulate the pressure distributions actually

measured in flight.

To allow different configurations of surface discontinuities to be examined, several aluminum strips were

milled. These strips, or coverpieces, fit into the spanwise slot on the left side of the noseshape. By using

cover-pieces of different widths and thicknesses, a variety of possible surface configurations could be tested.

Five configurations were chosen as representative and

were denoted with letters A through E (see fig. 7).

Configuration A consisted of two 0.06 in. gaps and a

0.02 in. step. Configuration B consisted of a single

0.03 in. gap with no step. Configuration C was similar

to configuration A but without the step. Configuration

D consists of a single 0.13 in. gap with no step.

Configuration E was designed to be identical to D

except that the downstream lip of the gap was rounded

rather than square. Work by other researchers has indicated that rounded corners on surface discontinuities

decreases their tendency to trip the boundary layer

(ref. 5).
The noseshape was instrumented with a single

chordwise row of pressure taps located approximately

halfway between the fixture root and tip (see fig. 8).

This row of pressure taps wrapped around the leading

edge providing pressure distributions on both sides of

the nose. A comparison of these pressure distributions

was used to ensure that the fixture angle of attack

(aircraft side-slip) stayed near zero. To determine

boundary layer state, hot-film gauges and Stanton

gauges were used. The arrangement of these gauges

was chosen to unambiguously determine if the boundary layer was tripped at or near the surface discontinuities (rather than the precise location of the transition

front).

The Stanton gauge is a simple, effective method for

measuring skin friction. The Stanton gauges were

formed by placing segments of razor blades over a

static pressure taps to form a kind of flattened total

pressure tube (ref. 6). The sloped cutting edge of the

blade forms the upper wall of the tube, and the surface

forms the lower wall (see fig. 9). The blade faces

directly upstream and the tip was aligned directly over

the forward-most lip of the pressure tap. To obtain a

skin friction measurement with the Stanton gauge, the

difference between the gauge pressure and the local

static pressure must be obtained. This was accomplished by having a static tap located adjacent to the

Stanton gauge at the same chordwise location and just

far enough away to prevent interference in the measured static pressure from the razor blade. The flow

over the noseshape was assumed to be twodimensional.

Skin friction was determined by applying East's calibration to the pressure difference (AP) measured by the

Stanton gauge (ref. 7). East's calibration relates two

dimensionless quantities,

= \og<

AP-h 2- p \

where h is the gage height.

The first of these quantities (x*) can be determined by

knowing the Stanton-gauge-pressure difference, the

flow properties and the gauge geometry. Using the

calibration relation,

v*=-0.23 +0.618-/+0.0l65-(jc*

the value of the second dimensionless quantity (y*) can

be determined. Once the value of y* has been found,

the same flow properties and gauge geometry information can be used to solve for the skin friction (T).

Because the calibration relation uses the flow properties, the exact relation between skin friction and

Stanton-gauge-pressure difference is different for each

flight condition. The large changes in skin friction

associated with boundary layer transition result in large,

easily identifiable changes in the Stanton-gaugepressure difference.

Despite the fact that East's calibration was made for a

turbulent boundary layer, as long as the razor blade is

sized correctly, it should be valid with 20% accuracy in

laminar boundary layers (ref. 8). The razor blade was

sized so that it lay completely within the boundary

layer for both laminar and turbulent boundary layers.

East's calibration does require that the proportions of

the razor blade strictly adhere to those used in the initial

calibration. Four Stanton gauges were placed downstream of the coverpiece, each at one of two chordwise

locations. Gauges number 2 and 3 were placed immediately downstream of the coverpiece and gauges 1 and 4

were placed further downstream. A small trip strip was

placed just downstream of the coverpiece directly in

front of gauge 1 so that the reading from gauge 1 could

be used as a known turbulent signal.

Five hot-film gauges were also placed on the noseshape. These were located adjacent to the Stanton

gauges with the fifth one just upstream of the coverpiece, in front of the trip strip for gauge 1. The results

of the hot films are presented in Zuniga et al. (ref. 9).

A range of target test conditions was chosen based on

the typical cruise operating conditions of business jets

and light transports. Mach numbers were varied

between 0.50 and 0.80 in increments of 0.10 and two

Reynolds numbers were selected: 1.50 x 10 and

2.5 x 10 per foot. The actual test points varied somewhat from the target values, being affected by

operational limitations and weather conditions.
Results

During the first flight, it was discovered that gauge 3

was malfunctioning; it was not possible to correct this

problem. As a result only three working gauges could

be used: gauge 1 was located at x = 11.5 in. and

because of the presence of a trip strip, provided a

known turbulent signal. Gauge 2 was located at

x = 5.5 in. and measured the skin friction approximately half an inch downstream of the aft edge of the

coverpiece. Gauge 4, located at x = 11.5 in., provided a

measurement six-and-a-half inches downstream of the

coverpiece. By comparing the CfS measured at these

three locations it was possible to determine if the

boundary layer was laminar or turbulent at each location (see table 2 and figs. 10-15).

Because of the arrangement of the Stanton gauges, it

was not possible to determine the exact location of

transition. Rather, it could only be said that transition

occurred in one of three regions: 1) Xtr < 5.5 in.—if

both the forward and the downstream Stanton gauge

indicated turbulent flow the transition location was

either on the coverpiece or immediately downstream of

it, 2) 5.5 in. < Xtr < 11.5 in.—if the forward gauge indicated laminar flow and the downstream gauge indicated

turbulent, and 3) Xtr > 11.5 in.—if both gauges indicated laminar flow, transition must have occurred aft of

all the gauges.

Because of the limited number of flights, most configurations could only be tested on a single flight. As a

result, it was not possible to test more than twelve conditions for any one configuration, and for most configurations the number of conditions tested was less than

this. The boundary layer was observed to remain laminar downstream of the coverpiece at a total of 29 of the

46 different conditions tested; at many of these conditions the forward gauge identified a laminar boundary

layer and the downstream gauge measured turbulent.

The location of boundary layer transition was clearly

dependent on the coverpiece configuration.

Comparisons between configurations generally show

that as qualitatively expected, the smaller the discontinuities in the surface, the higher the Reynolds number

at which the boundary layer is observed to be turbulent

(see figs. 16-20).

From the data obtained, it is not clear that the transition

location was dependent on Mach number. However, for

configurations B and D the results show a possible

influence of Mach number. With configuration B, at

approximately Re = 2.0 x 10 /ft, the boundary layer

remained laminar past the aft gauge for all Mach numbers except the highest, M = 0.8. At this condition, the

aft gauge indicated a turbulent boundary layer.

However, it is not clear that this must be due to the

effects of Mach number since the Reynolds number at

this condition was also slightly higher than the others.

A similar situation occurred for configuration D at a

nominal Reynolds number of 1.5 x 10 /ft.

A comparison of the results of configurations D and E

show that there may be some beneficial effects from

rounding the edge, as in configuration E. However, due

to the sparseness of the data, any such benefits cannot

be determined absolutely.

Despite the fact that the data points examined are rather

sparse and the number of step and gap configurations

limited, a beginning has been made at understanding

the effects of steps and gaps in aircraft surfaces on

boundary layer transition. Further work should include

an attempt to more precisely measure the transition

location in order to better understand the effects of

changes in Mach and Reynolds numbers. Additionally,

all work done here was with an unswept leading edge,

whereas many aircraft of the type for which one might

consider NLF or HLFC designs have swept wings.

Many of the techniques used in this experiment could

be applied to a test with a swept leading edge. Such a

test could make use of Stanton gauges for skin friction

measurements with confidence. Most importantly, this

experiment is only a starting point for comprehensive

research aimed at not only quantifying the allowable

step and gap sizes, but at an understanding of the

phenomena that create these limits.
References

1. Drake, Aaron; and Kennelly, Robert A., Jr.:

Selected Experiments in Laminar Flow: An

Annotated Bibliography. NASA TM-103989,

December 1992.

2. Meyer, Robert R., Jr.: A Unique Flight Test

Facility: Description and Results. NASA

TM-84900, 1982.

3. Campbell, R. L.; Waggoner, E. G.; and Phillips,

P. S.: Design of a Natural Laminar Flow Wing

for a Transonic Corporate Transport. AIAA

Paper 86-0314, 1986.

4. Kennelly, Robert A., Jr.: Improved Method for

Transonic Airfoil Design-by-Optimization.

AIAA paper 83-1864, 1983.

5. Holmes, B. J.; and Obara, C. J.: Observations and

Implications of Natural Laminar Flow on

Practical Airplane Surfaces. NASA TP-2256,

1984.

6. Hool, J. N.: Measurement of Skin Friction Using

Surface Tubes. Aircraft Engineering, vol. 28,

1956.

7. East, L. F.: Measurement of Skin Friction at Low

Subsonic Speeds by the Razor-Blade

Technique. Aero. Res. Counc. London,

R&M3525, 1966.

8. Bradshaw, P.; and Gregory, N.: The Determination

of Local Turbulent Skin Friction from

Observation in the Viscous Sub-layer. Aero.

Res. Coun. London, R&M 3202, 1961.

9. Zuniga, Fanny A.; Drake, Aaron; Kennelly,

Robert A., Jr.; Koga, Dennis J.; and Westphal,

Russell V.: Transonic Flight Tests of a

Laminar Flow Leading Edge with Surface

Excrescences. NASA TM-4597, August 1994.

Also presented as AIAA Paper 94-2142,

Colorado Springs, Colorado, June 1994.
Table 1. Symmetric noseshape surface coordinates

x (in.)

0.000

0.152

0.304

0.456

0.608

0.760

0.912

1.064

1.217

1.369

1.521

1.673

1.825

1.977

2.129

2.281

2.433

2.585

2.737

2.889

3.041

3.193

3.346

3.498

3.650

3.802

3.954

4.106

4.258

4.410

4.562

4.714

4.866

5.018

5.170

5.322

5.475

5.627

y (in.)

0.000

0.408

0.577

0.706

0.815

0.912

0.999

1.079

1.153

1.223

1.289

1.352

1.412

1.468

1.523

1 .575

1.625

1.674

1.722

1.768

1.812

1 .856

1.898

1.939

1.978

2.017

2.054

2.091

2.127

2.162

2.196

2.229

2.262

2.294

2.325

2.356

2.386

2.415

x (in.)

7.756

7.908

8.060

8.212

8.364

8.516

8.668

8.820

8.972

9.124

9.276

9.428

9.580

9.733

9.885

10.037

10.189

10.341

10.493

10.645

10.797

10.949

11.101

11.253

1 1 .405

11.557

11.709

11.862

12.014

12.166

12.318

12.470

12.622

12.774

12.926

13.078

13.230

13.382

y (in.)

2.773

2.795

2.816

2.837

2.858

2.878

2.898

2.917

2.937

2.955

2.974

2.991

3.009

3.026

3.043

3.059

3.075

3.090

3.105

3.120

3.134

3.148

3.161

3.174

3.186

3.198

3.210

3.221

3.231

3.242

3.251

3.261

3.270

3.278

3.286

3.294

3.301

3.308
Table 1. Concluded

5.779

5.931

6.083

6.235

6.387

6.539

6.691

6.843

6.995

7.147

7.299

7.451

7.604

2.444

2.473

2.501

2.528

2.555

2.581

2.607

2.632

2.656

2.681

2.704

2.728

2.750

13.534

13.686

13.838

13.991

14.143

14.295

14.447

14.599

14.751

14.903

15.055

15.270

3.314

3.320

3.325

3.330

3.333

3.337

3.340

3.342

3.344

3.345

3.347

3.347

Table 2. Matrix of the 46 test conditions

Point

la

Ib

Ic

Id

le

If

lg

Ih

2a

2b

2c

2d

2e

2f

2g

2h

2i

3a

3b

3c

4a

4b

Config.

A

A

A

A

A

A

A

A

B

B

B

B

B

B

B

B

B

B

B

B

C

C

Ma

0.742

0.699

0.592

0.806

0.743

0.711

0.580

0.491

0.775

0.747

0.800

0.748

0.599

0.698

0.798

0.751

0.597

0.599

0.701

0.599

0.799

0.695

Re/ft

1.57x 10°

1.53x 10°

1.45x 10°

2.44 x 10°

2.40 x 10°

2.42 x 10°

2.33 x 10°

2.34x 10°

1.47x 10°

1.55x 10°

2.03 x 10°

1.99x 10°

1.52x 10°

1.99x 10°

2.50 x 10°

2.52 x 10°

1.98x 10°

1.52x 10°

2.51 x 10°

2.52 x 10°

1.57x 10°

1.51 x 10°

Cf2

0.00097

0.00106

0.00105

0.00433

0.00488

0.00461

0.00459

0.00450

0.00178

0.00162

0.00175

0.00158

0.00154

0.00149

0.00242

0.00169

0.00141

0.00142

0.00140

0.00134

0.00166

0.00145

Cfl

0.00294

0.00298

0.00299

0.00326

0.00315

0.00307

0.00305

0.00303

0.00324

0.00308

0.00320

0.00314

0.00319

0.00310

0.00324

0.00314

0.00312

0.00306

0.00315

0.00307

0.00333

0.00317

Cf4

0.00320

0.00325

0.00327

0.00360

0.00338

0.00327

0.00326

0.00322

0.00096

0.00126

0.00435

0.00209

0.00097

0.00231

0.00356

0.00363

0.00238

0.00165

0.00377

0.00396

0.00407

0.00397

x,r(in.)

5.5-11.5

5.5-11.5

5.5-11.5

0.0-5.5

0.0-5.5

0.0-5.5

0.0-5.5

0.0-5.5

11.5-

11.5-

5.5-11.5

11.5-

11.5-

11.5-

5.5-11.5

5.5-11.5

11.5-

11.5-

5.5-11.5

5.5-11.5

5.5-11.5

5.5-11.5
Table 2. Concluded

4c

4d

4e

4f

4g

4h

4i

4j

5a

5h

5c

5d

5e

5f

6a

6b

6c

6d

6e

6f

6g

6h

6i

6j

C

C

C

C

C

C

C

C

D

D

D

D

D

D

E

E

E

E

E

E

E

E

E

E

0.799

0.704

0.801

0.699

0.599

0.605

0.505

0.499

0.800

0.693

0.595

0.697

0.690

0.599

0.800

0.707

0.806

0.733

0.798

0.508

0.596

0.601

0.503

0.493

2.03 x 10°

2.00 x 10°

2.44 \ ]()"

2.52 x 10°

2.52x 10"

2.02 x 10°

2.03 x 10°

2.47 x 10°

1.54x 10°

1.46x 10°

1.51 x 10°

2.00 x 10°

2.48 x 10°

2.51 x 10°

1.49x 10°

1.53 x 10°

2.08 x 10°

2.08 x 10°

2.51 x 10°

1.54x 10°

1.97x 10°

2.54 x 10°

2.02 x 10°

2.46 x 10°

0.00185

0.00153

0.00330

0.00163

0.00142

0.00141

0.00166

0.00411

0.00533

0.00184

0.00296

0.00466

0.00462

0.00446

0.00259

0.00184

0.00522

0.00478

0.00512

0.00220

0.00433

0.00441

0.00466

0.00465

0.00353

0.00334

0.00355

0.00324

0.00316

0.00325

0.00323

0.00316

0.00330

0.00312

0.00307

0.00323

0.00322

0.00318

0.00367

0.00325

0.00340

0.00327

0.00347

0.00311

0.00319

0.00315

0.00318

0.00320

0.00391

0.00402

0.00371

0.00366

0.00362

0.00379

0.00393

0.00349

0.00308

0.00356

0.00350

0.00319

0.00320

0.00324

0.00364

0.00334

0.00325

0.00324

0.00328

0.00322

0.00331

0.00317

0.00330

0.00325

5.5-11.5

5.5-11.5

5.5-11.5

5.5-11.5

5.5-11.5

5.5-11.5

5.5-11.5

0.0-5.5

0.0-5.5

5.5-11.5

5.5-11.5

0.0-5.5

0.0-5.5

0.0-5.5

5.5-11.5

5.5-11.5

0.0-5.5

0.0-5.5

0.0-5.5

5.5-11.5

0.0-5.5

0.0-5.5

0.0-5.5

0.0-5.5
Figure 1. Typical business jet wing leading edge.
PAINT

SCREW

SEALANT

FLUSH RIVET

FRONT SPAR

Figure 2(a). Typical business jet wing leading edge assembly.

L.E. SKIN COVER PIECE PAINT

SCREW

SEALANT

FLUSH RIVET

FRONT SPAR

Figure 2(b). Proposed wing leading edge assembly.

10
Figure 3. F-104/FTF with Laminar Leading Edge noseshape.

I I
Figure 4. Laminar Leading Edge noseshape.

12
-0.0

.6

.4

.2

o °-2

.4

.b

.0

i i 1 •t

4

.2

K

i»"

*

I

t*

*

) 5

i * JUt

) 1

x (in.)

~ x x

1 » • .

0 1

+ M =0.51

• M^O.60

A M^O.70

X M^O.80

5

Figure 5(a). Noseshape pressure distributions for Re = 1.5 x 10 .

x(in)

0.00

0.50

1.00

1.75

2.50

3.50

5.50

6.50

7.75

9.00

10.25

11.50

12.75

14.00

M 00=0.51

0.972

0.051

-0.141

-0.294

-0.345

-0.396

-0.435

-0.422

-0.428

-0.409

-0.416

-0.436

-0.448

-0.435

M oo=0.60

1.000

0.109

-0.090

-0.256

-0.269

-0.397

-0.449

-0.450

-0.442

-0.423

-0.449

-0.474

-0.476

-0.455

M oo=0.70

1.057

0.191

-0.019

-0.203

-0.292

-0.381

-0.432

-0.433

-0.433

-0.432

-0.457

-0.470

-0.464

-0.445

M oo=0.80

1.121

0.255

0.025

-0.191

-0.293

-0.408

-0.586

-0.459

-0.497

-0.498

-0.535

-0.573

-0.637

-0.650

Figure 5(b). Noseshape pressure distributions for Re = 1.5 x 10 .

13
-u.a

.6

.4

.2

O- no

o °-z

.4

.6

.0

1

 1

19

 3

1.2

C

.*;

** ^ • ^ -!*

*

*

) E

i

*

*

 * * * T iL

* * * *

> K

x (in.)

24 * + * * 4

3 1!

•* M -0 49 oo

• M^O.60

A M^O.70

XM =0.80 OO

Figure 6(a). Noseshape pressure distributions for Re = 2.5 x 10

x(in)

0.00

0.50

1.00

1.75

2.50

3.50

5.50

6.50

7.75

9.00

10.25

11.50

12.75

14.00

M oo=0.49

0.955

0.411

0.177

0.000

-0.101

-0.190

-0.253

-0.266

-0.291

-0.304

-0.329

-0.348

-0.354

-0.342

M oo=0.60

1.056

0.135

-0.023

-0.176

-0.258

-0.329

-0.382

-0.376

-0.382

-0.399

-0.423

-0.434

-0.446

-0.424

M oo=0.70

1.105

0.235

-0.012

-0.182

-0.282

-0.365

-0.423

-0.417

-0.435

-0.441

-0.470

-0.476

-0.494

-0.482

M oo=0.80

1.132

0.289

0.063

-0.145

-0.252

-0.377

-0.484

-0.453

-0.484

-0.497

-0.522

-0.566

-0.641

-0.717

Figure 6(b). Noseshape pressure distributions for Re = 2.5 x 10 .

\ 4
FLOW

0.02" i

T

1/16"

1/32"

1/16"

1/8"

1/8"

.00"_

Configuration A

Configuration B

Configuration C

Configuration D

Configuration E

1/16"

Aluminum

•*

Adhesive

1/16"

Figure 7. Coverpiece configurations.

15
10.0

Trip

6.0

3.i

Pressure Taps

Stanton Probe

& Hot Film

 5

-

25

Slot for Cover-Pieces

#5

5.5

• • • »

5.5

* 4.0 *n .0

11.5

13.5

— • *1

^

11.5

1

15.27

21.75

23.88

• Hot Film gage

0

 Stanton gage (consists of blade-covered pres. tap and an uncovered adjacent pres. tap)

Pressure taps at specified x locations on both left and right side of nose.

Ten taps on right side, thirteen on left, one on leading edge.

ALL DIMENSIONS IN INCHES

Figure 8. Instrumentation layout.

16
FLOW

Razor Blade Segment

Pressure Tap

Adhesive

Figure 9(a). Stanton Gauge schematic.

17
Figure 9(b). Stanton Gauge on Nosepiece.

is
0.004

0.003-

0.002-

0.001 -

0.000

 J

forward aft

Gauge Location

Figure 10(a). Condition 1a, M^ = 0.74, Re/ft = 1.57x106.

0.005

0.004-

0.003-

o 0.002-

0.001 -

o.oooforward aft

Gauge Location

Figure 10(b). Condition 1b, Mx = 0.70, Re/ft =1.53 x W6.

0.005

0.004-

0.003-

o 0.002-

0.001 -

0.000 ^

forward aft

Gauge Location

0.005

0.004-

0.003-

0.002-

0.001 -

0.000

forward aft

Gauge Location

Figure 10(c). Condition 1c, M«, = 0.59, Re/ft = 1.45 x 106. Figure 10(d). Condition 1d, M^O.81, Re/ft = 2.44 x 106

19
0.005

0.004-

0.003-

0

 0.002-

0.001 -

0.000

0.005

0.004-

0.003-

0

 0.002-

0.001 -

0.000

forward aft

Gauge Location

forward aft

Gauge Location

Figure 10(e). Condition 1e, Mx = 0.74, Re/ft = 2.40 x 106. Figure 10(f). Condition 1f, Mx = 0.71, Re/ft = 2.42 x 10t

0.005

0.004-

0.003-

O 0.002-

0.001 -

0.000

0.005

0.004-

0.003-

0.002-

0.001 -

0.000

forward aft

Gauge Location

forward aft

Gauge Location

Figure 10(g). Condition 1g, Mx = 0.58, Re/ft = 2.33 x 106. Figure 10(h) Condition 1h, M^ = 0.49, Re/ft = 2.34 x 106.

20
0.005

0.004-

0.003-

O 0.002^

0.001 -

0.000

Turbulent reference

\

forward aft

Gauge Location

Figure 11(a). Condition 2a, Mx = 0.78, Re/ft = 1.47 x 106.

0.005

0.004-

0.003-

0.002-

0.001 -

Turbulent reference

\

o.oooforward aft

Gauge Location

Figure 11(b). Condition 2b, Mx = 0.75, Re/ft = 7.55 x 106.

0.005

0.004-

0.003-

0.002-

0.001 -

0.000

forward aft

Gauge Location

0.005

0.004-

0.003-

o

0.002-

0.001 -

0.000

Turbulent reference

\

forward aft

Gauge Location

Figure 11(c). Condition 2c, Mm = 0.80, Re/ft = 2.03 x 106. Figure 11(d). Condition 26, Mm = 0.75, Re/ft = 1.99 x 10b.

21
0.005

0.004-

0.003-

O 0.002 -

0.001 -

0.000

Turbulent reference

\

forward aft

Gauge Location

Figure 11(e). Condition 2e, M^ = 0.60, Re/ft = 1.52x10

0.005

0.004-

0.003-

o 0.002-

0.001 -

0.000

Turbulent reference

\

forward aft

6

Gauge Location

Figure 11(f). Condition 2f, Mm = 0.70, Re/ft = 1.99x10*

0.005

0.004-

0.003-

0.002-

0.001 -

0.000

0.005

forward aft

Gauge Location

0.004-

0.003-

O 0.002-

0.001 -

0.000

forward aft

Gauge Location

Figure 11(g). Condition 2g, M^, = 0.80, Re/ft = 2.50 x 10°. Figure 11(h). Condition2h, Mx = 0.75, Re/ft = 2.52 x 10t

22
0.005

0.004-

0.003-

O 0.002-

0.001 -

0.000-I

Turbulent reference

\

0.005

0.004-

0.003-

0.002-

0.001 -

0.000

Turbulent reference

\

forward aft

Gauge Location

forward aft

Gauge Location

Figure 11(i). Condition 2i, Mm = 0.60, Re/ft = 1.98 x 10b. Figure 12(a). Condition 3a, Mro = 0.60, Re/ft = 1.52 x

0.005

0.004-

0.003-

O 0.002-

0.001

 J

0.000

0.005

0.004-

0.003-

0

 0.002-

0.001 -

0.000

Turbulent reference

\

forward aft

Gauge Location

forward aft

Gauge Location

Figure 12(b). Condition 3b, Mx = 0.70, Re/ft = 2.51 x 10°. Figure 12(c). Condition 3c, Mx = 0.60, Re/ft = 2.52 x 10t

23
0.005

0.004-

0.003-

o

0.002-

0.001 -

0.000

0.005

0.004-

0.003-

0.002-

0.001 -

0.000

forward aft

Gauge Location

forward aft

Gauge Location

Figure 13(a). Condition 4a, Mx = 0.80, Re/ft = 1.57 x 10b. Figure 13(b). Condition 4b, Mx = 0.70, Re/ft = 1.51 x 70fc

0.005

Turbulent reference

0.004-

0.003-

o

0.002-

0.001 -

0.000

0.005

0.004-

0.003-

O 0.002-

0.001 -

0.000

forward aft

Gauge Location

forward aft

Gauge Location

Figure 13(c). Condition 4c, Mx = 0.80, Re/ft = 2.03 x 106. Figure 13(d). Condition 4d, Mx = 0.70, Re/ft = 2.00*10*

24
0.005

0.004-

0.003-

O 0.002-

0.001 -

o.ooo -*

forward aft

Gauge Location

Figure 13(e). Condition 4e, M^ = 0.80, He/ft = 2.44 x 106.

0.005

0.004-

0.003-

o 0.002-

0.001 -

0.000

forward aft

Gauge Location

Figure 13(f). Condition 4f, Mx = 0.70, Re/ft = 2.52x10t

0.005

0.004-

0.003-

o 0.002-

0.001 -

0.000

forward aft

Gauge Location

0.005

0.004-

0.003-

O 0.002-

0.001 -

0.000

forward aft

Gauge Location

Figure 13(g). Condition 4g, Mm = 0.60, Re/ft = 2.52 x W6. Figure 13(h). Condition 4h, Mx = 0.61, Re/ft = 2.02 x

25
0.005

0.004-

0.003-

o 0.002-

0.001 -

0.000

0.005

0.004-

0.003-

0

 0.002-

0.001 -

0.000

forward aft

Gauge Location

forward aft

Gauge Location

Figure 13(i). Condition 4i, /W«, = 0.51, Re/ft = 2.03 x 706. Figure 13(j). Condition 4j, M^ = 0.50, Re/ft = 2.47 x 10t

0.006

0.005-

0.004-

0.003-

o

0.002-

0.001 -

0.000

0.005

0.004-

0.003-

O 0.002-

0.001 -

0.000

forward aft

Gauge Location

forward aft

Gauge Location

Figure 14(a). Condition 5a, M^ = 0.80, Re/ft = 1.54 x 10b. Figure 14(b). Condition 5b, Mx = 0.69, Re/ft = 1.46 x 10t

26
0.005

0.004-

0.003-

0.002-

0.001 -

0.000

forward aft

Gauge Location

Figure 14(c). Condition 5c, Mx = 0.60, Re/ft = 1.51 x 106.

0.005

0.004-

0.003-

O 0.002-

0.001 -

o.oooforward aft

Gauge Location

Figure 14(d). Condition 5d, Mm =0.70, Re/ft = 2.00 xW6.

0.005

0.004-

0.003-

0.002-

0.001 -

o.oooforward aft

Gauge Location

Figure 14(e). Condition 5e, MM = 0.69, Re/ft = 2.48 x 106.

0.005

0.004-

0.003-

o 0.002 -

0.001 -

0.000

forward aft

Gauge Location

Figure 14(f). Condition 5f, M^ = 0.60, Re/ft = 2.51 x 106

21
0.005

0.004-

0.003-

0

 0.002-

0.001-

0.000

0.005

0.004-

0.003-

o 0.002-

0.001 -

0.000

forward aft

Gauge Location

forward aft

Gauge Location

Figure 15(a). Condition 6a, Mx = 0.80, Re/ft = 1.49 x 10b. Figure 15(b). Condition 6b, M^ = 0.71, Re/ft = 1.53 x 10t

0.006

0.005-

0.004-

0.003-

O

0.002 -

0.001 -

0.000

0.005

0.004-

0.003-

0

 0.002-

0.001 -

0.000

forward aft

Gauge Location

forward aft

Gauge Location

Figure 15(c). Condition 6c, M^ = 0.81, Re/ft = 2.08 x 106. Figure 15(d). Condition 6d, Mm = 0.73, Re/ft = 2.08 x 10e

28
0.006

0.005-

0.004-

0.003-

0.002-

0.001 -

0.000

0.005

0.004-

0.003^

0.002 -

0.001 -

0.000

forward aft

Gauge Location

forward aft

Gauge Location

Figure 15(e). Condition 6e, Mm = 0.80, Re/ft = 2.51 x 10b. Figure 15(f). Condition 6f, M«, = 0.51, Re/ft = 1.54 x 10t

0.005

0.004-

0.003-

o 0.002-

0.001 -

0.000

forward aft

Gauge Location

0.005

0.004-

0.003-

o 0.002-

0.001 -

0.000

forward aft

Gauge Location

Figure 15(g). Condition 6g, M^ = 0.60, Re/ft = 1.97x10°. Figure 15(h). Condition 6h, M^ = 0.60, Re/ft = 2.54x10*
0.005

0.004-

0.003-

O 0.002-

0.001 -

0.000

forward aft

Gauge Location

0.005

0.004-

0.003-

o

0.002-

0.001 -

o.oooTurbulent reference

forward aft

Gauge Location

Figure 15(i). Condition 6i, M^ = 0.50, Re/ft = 2.02 x 10°. Figure 15(j). Condition 6j, M«> = 0.49, Re/ft = 2.46 x Wt

30
Re/ft

(in millions)

.'H

? fi

? 4

2.2.

2 0 -

Configuration A

D Xtr<55'

A 5.5"<Xtr<11.5"

0.5 0.6 O.B 0.9

Mach

Figure 16. Summary of results for Cover-Piece Configuration A.

Re/ft 2.0 -|

(in millions)

1 HConfiguration B

5.5'<Xlr<11.5"

0.5 0.6 0.7

Mach

Figure 17. Summary of results for Cover-Piece Configuration B.

31
Configuration C

30

2 8

2 4

2 2

Re/ft 2.0

(in millions)

1 8

1.6

1 2

1 0

D Xlr<5.5-

4 5.5"<Xlr<11.5"

Mach

Figure 18. Summary of results for Cover-Piece Configuration C.

2 8

2.6

24 .

Re/ft

(in millions)

1.8

1.6

0 4

Configuration D

0.6 0.7

Mach

D Xtr<5.5"

4 5.5'<Xtr<11.5-

0.8 0.9

Figure 19. Summary of results for Cover-Piece Configuration D.

32
Configuration E

3.0 •

28 .

26 -

24 -

2 2 -

Re/ft 2.0 •

(in millions)

1 8-

1.6.

1.4 •

1.2-

1.0-

0

D Xtr<5.5"

4 5.5'<Xlr<11.5-

0

D D

D

a

4

4 0.5 0.6 0.7 0.8 0.

Mach

Figure 20. Summary of results for Cover-Piece Configuration E.

33
REPORT DOCUMENTATION PAGE

Form Approved

OMB No. 0704-0188

Public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources,

gathering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this

collection of information, including suggestions for reducing this burden, to Washington Headquarters Services. Directorate for information Operations and Reports. 1215 Jefferson

Davis Highway. Suite 1204. Arlington. VA 22202-4302, and to the Office of Management and Budget, Paperwork Reduction Project (0704-0188), Washington. DC 20503.

1. AGENCY USE ONLY (Leave blank) 12. REPORT DATE

October 1996

3. REPORT TYPE AND DATES COVERED

Technical Memorandum

4. TITLE AND SUBTITLE

Wing Leading Edge Joint Laminar Flow Tests

6. AUTHOR(S)

Aaron Drake, Russell V. Westphal, Fanny A. Zuniga,

Robert A. Kennelly, Jr., and Dennis J. Koga

5. FUNDING NUMBERS

505-59-20

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

Ames Research Center

Moffett Field, CA 94035-1000

8. PERFORMING ORGANIZATION

REPORT NUMBER

A-962704

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TM-4762

11. SUPPLEMENTARY NOTES

Point of Contact: Robert A. Kennelly, Ames Research Center, MS 227-6, Moffett Field, CA 94035-1000;

(415)604-5860

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified — Unlimited

Subject Category 05

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

An F-104G aircraft at NASA's Dryden Flight Research Center has been equipped with a specially designed

and instrumented test fixture to simulate surface imperfections of the type likely to be present near the leading

edge on the wings of some laminar flow aircraft. The simulated imperfections consisted of five combinations of

spanwise steps and gaps of various sizes. The unswept fixture yielded a pressure distribution similar to that of

some laminar flow airfoils. The experiment was conducted at cruise conditions typical for business jets and light

transports: Mach numbers were in the range 0.5-0.8, and unit Reynolds numbers were 1.5-2.5 million per foot.

Skin friction measurements indicated that laminar flow was often maintained for some distance downstream of

the surface imperfections. Further work is needed to more precisely define transition location and to extend the

experiments to swept-wing conditions and a broader range of imperfection geometries.

14. SUBJECT TERMS

Laminar flow, Skin friction, Transition

15. NUMBER OF PAGES

36

16. PRICE CODE

A03

17. SECURITY CLASSIFICATION

OF REPORT

Unclassified

18. SECURITY CLASSIFICATION

OF THIS PAGE

Unclassified

19. SECURITY CLASSIFICATION

OF ABSTRACT

20. LIMITATION OF ABSTRACT

NSN 7540-01-280-5500 Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std Z39-18