NASA/TM--1999-209632 AIAA-2000-0545

An Experimental Study of Turbulent Skin

Friction Reduction in Supersonic Flow

Using a Microblowing Technique

Danny P. Hwang

Glenn Research Center, Cleveland, Ohio

December 1999

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f
NASA/TM--1999-209632 AIAA-2000-0545

An Experimental Study of Turbulent Skin

Friction Reduction in Supersonic Flow

Using a Microblowing Technique

Danny P. Hwang

Glenn Research Center, Cleveland, Ohio

Prepared for the

38th Aerospace Sciences Meeting and Exhibit

sponsored by the American Institute of Aeronautics and Astronautics

Reno, Nevada, January 10-13, 2000

National Aeronautics and

Space Administration

Glenn Research Center

December 1999
Acknowledgments

This work was funded by NASA Glenn Research Center's Fast Quiet Engine (FQE) program.

Special thanks are due to Mr. DavidArend who prepared the CFD design of the supersonic tunnel, and

Mr. Tony Herrmann who prepared the engineering design of the supersonic tunnel. The following individuals

are also acknowledged for their contributions during the experiments: Mrs. Gwynn Severt, Mr. Scott Panco,

Mr. Carlos Gomez, and Mr. Ron Abel. Finally, the author would like to express sincere thanks

to Dr. Greg Tillman of the United Technologies Research Center for reviewing the final draft

and providing valuable comments.

NASA Center for Aerospace Information

7121 Standard Drive

Hanover, MD 21076

Price Code: A03

Available from

National Technical Information Sen, ice

5285 Port Royal Road

Springfield, VA 22100

Price Code: A03
AN EXPERIMENTAL STUDY OF TURBULENT SKIN FRICTION

REDUCTION IN SUPERSONIC FLOW USING

A MICROBLOWING TECHNIQUE*

Danny P. Hwang+

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Abstract

A new turbulent skin friction reduction technology,

called the microblowing technique, has been tested in

supersonic flow (Mach number of 1.9) on specially

designed porous plates with microholes. The skin friction

was measured directly by a force balance and the boundary

layer development was measured by a total pressure rake

at the tailing edge of a test plate. The free stream Reynolds

number was 1.0(10) 6per meter. The turbulent skin friction

coefficient ratios (C_IC m) of seven porous plates are given

in this report. Test results showed that the microblowing

technique could reduce the turbulent skin friction in

supersonic flow (up to 90 percent below a solid flat plate

value, which was even greater than in subsonic flow).

A area of test plate

C_ total skin friction coefficient, (skin friction

force)/(1/2 p_ u_2A)

C_ total skin friction coefficient of a nonporous

plain flat plate

Cf,,.0 total skin friction coefficient of a porous plate

with no blowing

D diameter of blowing holes, mm

F blowing fraction. (Pt, vt,)/(P= u_)

H shape factor (ratio of displacement thickness

to momentum thickness)

T thickness of a plate

u velocity component parallel to surface, m/sec

v velocity component perpendicular to surface,

m/sec

u free stream velocity, m/sec

v_ blowing air velocity, m/sec

y vertical distance from a surface, cm

p= free stream density

Pb blowing air density

*patent no. 5,803,410

_'Seniormember, AIAA

Introduction

One of the challenges associated with the development

of a high-speed civil aircraft is the reduction of high skin

friction. One technique that has received much attention

for subsonic flow is laminar flow control. Its application

in supersonic flow is difficult. Many other skin friction

reduction techniques (summarized in Ref. 1) cannot provide

significant reduction in real world applications.

One proposed solution to skin friction reduction,

blowing (or surface mass injection), has been abandoned

because it had problems with flow separation. Many

researchers during the past 30 years have used the

parameters, 2F/Cfw0 versus C/Cf,vo, for their results and

have shown that blowing could achieve significant skin

friction reduction? -3 However, it was not clear whether

blowing could result in a reduction in skin friction below

a solid flat plate value. As indicated in Ref. 4, the unblown

skin friction coefficients, C_,_,of some porous plates were

more than twice the value of a solid flat plate and it was

impossible to reduce the skin friction of these plates below

that of a solid flat plate.

The microblowing technique (MBT) for reducing

turbulent skin friction was invented in 1994 by the author

and a patent was issued in 1998. 5In this unique concept,

an extremely small amount of air is blown vertically

through a specially designed porous plate with microholes.

The reduction in the surface roughness and the gradient of

the flow velocity profile on the surface decreases skin

friction.

Several experiments have been completed for subsonic

flow 5-8 and results showed that a turbulent skin friction

reduction of up to 50 to 70 percent below the value of a

solid flat plate was possible for subsonic flow if proper

porous plates were used.

A proof of concept MBT experiment for reducing

skin friction in supersonic flow at a Mach number of 1.9

was conducted in test facility CE229 at the NASA Glenn

Research Center in 1999. Seven test skins (Table 1) were

NASAfI'M--1999-209632 1
selectedfromtheprevioussubsonictests.Theinnerskin

thatwasusedformanysubsonictestswaseliminated. The

preliminaryresultshowedthattheMBTcouldreduce

skinfrictionupto90percent belowasolidflatplatevalue

insupersonicflow.

Inthispaper,thedetaileddescriptionftheexperiment,

theresultsof testdataandthediscussionof thisnew

technologywiIlbegiven.

Skin Tested

Seven test skins (Table l) were selected from the

previous subsonic tests. The inner skin that was used for

many subsonic tests was eliminated because it was believed

that the skin friction reduction was not affected by the

inner skin and also because the pressure loss across the

inner plate could be eliminated. Laser drilled microholes

(0.1651 mm in diameter) are shown in Fig. I and the crosssection view of 15-degree, slanted laser-drilled holes is

given in Fig. 2. Northrop Grumman GAC1897 has

hourglass-shaped holes with_0-percent porosity based on

an open hole area and only has 4 percent porosity based on

the small neck area as shown in Fig. 3. The schematic

diagram of hole pattern showing staggered holes and

aligned holes is illustrated in Fig. 4.

The size of the test plates is 12.36 by 25.06 cm.

Test Facility

The Advanced Nozzle and Engine Components Test

Facility (CE22) at NASA Glenn 9 was modified for this

experiment. A shock-free converging-diverging supersonic

duct 109.22 cm long is connected to a facility transition

duct with a rectangular cross-section exit of 10.8 by

20.32 cm. The facility provided a stable Mach number

of 1.9 at Reynolds number per meter of 1.0(10). 6 The back

pressure (exit pressure) was maintained as low as possible

to avoid keeping the shock inside the duct for too long,

which would damage the force balance.

Apparal_ _and Instrumentation

The force balance used in previous subsonic

experiments was used for this experiment? Since larger

skin friction and blowing air requirements were expected,

a 4800-gram load cell and a 1500 SLM flow meter were

used. The accuracy of the load cell used in the balance was

+/- 0.25 percent. The balance was placed under the test

section inside a sealed compartment, as shown in Fig. 5.

The leading edge of the test plate is located in a constant

pressure region, 63.5 cm from the transition duct. There

were two total pressure rakes along the center line of the

duct. The measurement plane of rake I was located 3.2 cm

upstream of the leading edge of the test plate while the

measurement plane of rake 2 was 1. I cm upstream of the

trailing edge of the test plate. Rake 2 had a very small

vertical gap (less than 0.2 mm) between the test plate and

the nearest tube. The total pressure rakes were built with

tubing that had an outside diameter of 0.508 mm. Rake 1

was removed after the data was taken to minimize the

interference on the measurement of skin friction. Rake 2

was used to calculate the momentum thickness and the

velocity profiles inside the boundary layer. Facility total

pressure rakes, facility total temperature thermocouples,

and a static pressure at the leading edge of the test plate

were used to measure the tunnel Mach number.

Calibration

An identical load cell was used to calibrate the load

cell inside the balance. The calibration showed that the

friction of the balance was very small due to the frictionless

flexuraI pivots. The data were adjusted during the data

acquisition process for this small difference.

Results and Discussion

The seven porous plates selected from previous

subsonic tests (Table I) were tested in supersonic flow

(Mach number of 1.9) in the region of no pressure gradient.

The standard deviation of force measurement was kept

below 2 grams for 15 data scans. The test results are

presented herein.

The wall static pressure along the supersonic duct is

shown in Fig. 6. As expected, there were no shocks in the

duct because there was not an abrupt increase in pressure

in the duct. The test plate was located from x = 76.2 to

101.6 cm from the entrance of a transition duct. A fairly

constant pressure gradient region is shown in Fig. 6,

including a case with the highest blowing rate.

A solid stainless steel flat plate without microholes

was first tested and the skin friction coefficient (Cft ,) was

considered as a reference. The repeatability of the facility

is shown in Fig. 7 for the results taken on April 14, 1999

and April 22, 1999. The small difference could be due to

the different level of turbulence on these two days. Based

on Tillman' s report, 7the uncertainty of drag measurement

is about 5 to I 0 percent, therefore, 2 to 3 percent difference

in measurement for these experiments is considered

acceptable. The total skin friction coefficients obtained

from this test for a plain stainless steel flat plate were

compared with Ludwieg and Tillmann's empirical

formula 1C_and is shown in Fig. 8. Keep in mind that the

total skin friction coefficients from the tests being reported

were the average value based on a 12.36- by 25.06-cm

plate. The momentum thickness was based on the

NASA/TM--1999-209632 2
downstreamtotalpressurerake,whichwashigherthanthe

momentumthicknessupstreanaatthecenteroftheplate.

However, theresultswerereasonablyclose.

Thetotalskinfrictioncoefficient ofaporousflatplate

(C_w0) withoutblowingwasmeasured. Next,theblowing

aircontrolvalvewasopenedtoallowfordifferentblowing

airflowrates.Theskinfrictioncoefficients(Cf)were

measuredattheMachnumberof 1.9.Jerominplotted

C/Cfw0against2F/Cfw0formanyresearchers' data,and

showedratherscattereddata3Manyresearchers alsoused

thesetwoparameters. Thereforethecurrentdatawerefirst

plottedinthesamewayasshowninFig.9. All skins

exceptGAC2054didverywellandfellwithinHefnerand

BushnelI'srange. 3TheGAC2054platewasselected

becauseithadaverylowunblownskinfrictioncoefficient

(Ct,,, _)insubsonicflow,butheskinfrictionreductionwas

notsignificantin supersonicflowbecauseof thelow

porosityof4percent andconical-shapedholes. Adrawback

ofthisfigureisthatonecannottellhowmuchskinfriction

reductionbelowthatofasolidflatplatehasbeenachieved.

Amoremeaningful waytopresent testresultsigiven

inFig.10,whereskinfrictionratios(C/Ct._) wereplotted

againsttheblowingfraction,F.All sevenporousplates,

whichwereselectedfromprevioussubsonictests,hadavery

lowunblownskinfriction.Thedatashowedthatskinfriction

ratioswereclosetoonewhenFwasequaltozero(i.e.,the

skinfrictionoftheseplateswithoutblowingwasverylow

andclosetoasolidflatplatevalue).Asexpected, theskin

frictionratiowasreducedbelowthatofaflatplatewithvery

lowblowingairandtheskinfrictioncontinuedtoreduce

technique. Theamountofreductionismuchlargerin

supersonicflowthaninsubsonicflow.Theporousplates

(exceptGAC2054), whichworkedverywellinsubsonic

flow,alsoworkedverywell in supersonicflow.The

inclinationandalignment ofholesdidnotseemimportant.

Thistechnologycanbeappliedtomanytypesofflows,

turbulentorlaminar,subsonicorsupersonic.

Aknowledgments

This work was funded by NASA Glenn Research Center's

Fast Quiet Engine (FQE) program. Special thanks are due to

Mr. David Arend who prepared the CFD design of the

supersonic tunnel, and Mr. Tony Herrmann who prepared

the engineering design of the supersonic tunnel. The following individuals are also acknowledged for their contributions during the experiments: Mrs. Gwynn Severt, Mr. Scott

Panco, Mr. Carlos Gomez, and Mr. Ron Abel. Finally, the

author would like to express sincere thanks to Dr. Greg

Tillman of the United Technologies Research Center for

reviewing the final draft and providing valuable comments.

References

1. Bushnell, D.M. and Hefner, J.N.: Viscous Drag Reduction

in Boundary Layers. Progress in Astronautics and

Aeronautics, volume 123,American Institute of Aeronautics

and Astronautics, Inc., Washington, DC, 1990.

2. Jeromin, L.O.F.: An Experimental Investigation of the

Compressible Turbulent Boundary Layer With Air

Injection, ARC Reports and Memoranda No. 3526,

until the limit of the flow meter was reached. The turbulent Aeronautical Research Council, London, England, 1968.

skin friction reductions of 80-90 percent below a flat plate

value were achieved at the highest flow rate, F = 0.004.

In Fig. 10, a very small difference was shown between

PN23 (straight vertical holes) and PN5 (15-degree slanted

holes) and also between PN5 (15-degree staggered holes)

and PN7 (15-degree aligned holes). This means that

additional effort to drill slanted holes is not necessary and

holes need not be aligned for real application.

The velocity profile obtained from the downstream

boundary layer rake is shown in Fig. 11. It was very clear

that blowing did reduce the flow velocity near the surface

such that it reduced the viscous shear stress. The magnitude

of reduction in flow velocity close to the surface is much

more than that of subsonic flow. _ This could be a reason

why the MBT reduced skin friction more for supersonic

flow than subsonic flow.

Cgn¢luding Remarks

A very significant turbulent skin friction r_-ffcti-0n

(up to 90 percent) below a solid flat plate value has been

achieved in supersonic flow using a microblowing

3. Hefner, J.N. and Bushnell, DM.: Viscous Drag Reduction

Via Surface Mass Injection. Viscous Drag Reduction in

Boundary Layers, vol. 123, Progress in Astronautics and

: Aeronautics. pp. 457-476, 1990.

4. Hwang, D.P.: A Proof of Concept Experiment for Reducing

Skin Friction by Using a Micro-Blowing Technique,

A1AA Paper No. 97-0546, 1997.

5. Hwang, D.P.: Skin Friction Reduction by Micro Blowing

Technique, U.S. patent no. 5,803,410, December 1995.

6. Hwang,D.P. and Biesiadny, TJ.: ExperimentalEvaluation

of the Penalty Associated With Micro-Blowing for

Reducing Skin Friction, AIAA Paper No. 98-0677, 1998.

7. Tillman, T.G.: Drag Reduction on a Large-Scale Nacelle

Using Micro-Blowing Technique, AIAA Paper No.

99-0130, 1999.

81 Hwang, D.P.: Skin Friction Reduction by a Mica'o-Blowing

Technique, AIAA J., vol.36, no. 3,March 1998,pp.480--481.

9. Beltran, L.R., Del Roso, R.L., and Del Rosario, R.: Advanced

Nozzle and Engine Components Test Facility. NASA

TM-1036M, 1992.

10. Ludwieg, H. and Tillmann, W.: Investigations of the

Wall Shearing Stress in Turbulent Boundary Layers.

NACA TM-1285, 1950.

NASA/TM--1999-209632 3
Platename

FlatPlate

GAC1897

GAC2054

Porosity,

percent

PN5

5O

Table1.--Specificationsoftestplates.

Hole

diameter,

mm

0.06

0.13

Skin

thickness

0.305

0.23

Aspect

ratio,

L/D

5.08

1.769

PN23 23 0.1651 1.016 6.15

PN4 23 0.1651 1.016 6.15

23 0.1651 1.016 6.15

PN6

PN7

23 1.016

1 23 1.016

0.1651

0.1651

6.15

6.15

Description/Material

StainlessSteel

NorthropGrumman'splate,nickel

NorthropGrumman'splate,nickel

Laserdrilledstraightholes,stainlesssteel

5-de_reean_,leholes,sta_ered,stainlesssteel

15-de_reean_leholes,sta_ered,stainlesssteel

5-de_,reean_leholes,aligned,stainlesssteel

15-degreean_leholes,aligned,stainlesssteel

Figurel.--Stainless-steel-laser-drilled microholes.

Figure2.--Cross-section viewof 15-degreeslanted-laser-drilled holes.

Figure3.--NorthropGrummanGAC1897hourglass-shaped microholes.

NASA/TM--1999-209632 4
000 Lo 00 000

00 000 000

000000 000

0 0 C 000 0 O0

k

000

000

000

000

Staggered holes Aligned holes

Figure 4.--Hole pattern.

I Test Plate

Force Balance

Figure 5.--Supersonic tunnel and balance.

90C(J0 -

i

70000

' D - (3 - F=0.00449

i

0 _

1 TestPlate I _

D-'I_, i

i

0 25 50 75 1130 125

x(cm)

Figure 6.--Wall static pressure along supersonic duct.

NASA/TM--1999-209632 5
1.2 -" i

1 ......... + GACi897,4/14/99 _

i -_ ---_,- GAC1897,4/22/99

| ,.

0.6 1

i

i " : i

0.4

_2

O

d

.o_=

LL

tv'

03 °2i

0 ..... i

O.0000 O.0010 0.0020 O.0030 O.0040

Blowing Fraction, F

Figure 7.--Repeatability of test facility.

0.004

0.0050

0.0035

0.003

0.0025

0.002...... _o

"_ 0.0015 .... ::': ..... ' ..... " ...................................................

o3

0.001. -- ......... Ludwlegand_llmann[ref. il], H=1.5' :

' -Ludwteg and 1_ mann, H=1.4 _ 1_

0.0005. _ ' 0 _:DATA ! : H=1.4 _ _. :

[

10000 10500 11000 11500 12000

Momentum _ickness Reynolds Number

Figure &--Skin friction coefficient of flat plate.

12500

1.20 ........

• GAC1897

PN23 !

1.00 _ x PN4

080 "-"'_, ,l,.,,q_ __ • PN6 i

' "'_". , _ "_.0 0 0 O 0 + PN7 i

"'-.. e_ _ ....... _-Lower Range of Hefr_r !

J_ ¢1_ "_'_-- _,_ _,_ and Bushnell data [10] ._

..... -. • _ _. -- Upper Range of Hefner i

"'_, _ •#_' _ andBushnelrdaa[10] [

;. - 3t _ # _ o GAC2054

-.. .o_ . _ i

/ ".. _• • __

0401 ...... _-C-_--_,, • __ q

I

"-,, • k: • i

0'0° t| " _ x"

0.013 0,50 1.00 1.50 2.00 2.50 3.00 3.50 4.00 4.50 5,00

2F/Q.,,.o

Figure 9.---Current data within range of Hefner and Bushnell data.

NASA/TM--1999-209632 6
1.2 .

0.(300 0.001 0.002 0.003 0.004 0.005

Blowing Fraction, F

Figure ]O.--Skin friction ratio of seven porous plates.

E

o

>-

1.8

1.6

1.4

1.2

1

0.8

0.6

0.4

0.2

0

2 _, ................................................................................................................................................................................................................................................

i

_F=O

i _-oooo_7 I

i _ 0.0007

-- ---o_ 0.00136 :-

_ +0.00204

---o--- 0.00277 /

i +°°- i _ ,

i

--!

11

i

!

I

0 1O0 200 300 400 500 600

U (m/sec)

Figure l l.--Boundary layer profile at the trailing edge of PN23.

NASAfI'M--1999-209632 7
REPORT DOCUMENTATION PAGE FormApproved

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1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

December 1999 Technical Memorandum

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

An Experimental Study of Turbulent Skin Friction Reduction in

Supersonic Flow Using a Microblowing Technique

6. AUTHOR(S)

Danny R Hwang

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

National Aeronautics and Space Administration

John H. Glenn Research Center at Lewis Field

Cleveland, Ohio 44135- 3191

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

National Aeronautics and Space Administration

Washington, DC 20546-0001

WU-523-36-13-00

8. PERFORMING ORGANIZATION

REPORT NUMBER

E-11959

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TM--1999-209632

AIAA-2000-0545

11. SUPPLEMENTARY NOTES

Prepared for the 38th Aerospace Sciences Meeting and Exhibit sponsored by the American Institute of Aeronautics and

Astronautics, Reno, Nevada, January 10-13, 2000. Responsible person, Danny P. Hwang, organization code 5850,

(216) 433-2187.

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Category: 02 Distribution: Nonstandard

This publication is available from the NASA Center for AeroSpace Information, (301) 621-0390.

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

A new turbulent skin friction reduction technology, called the microblowing technique has been tested in supersonic flow

(Mach number of 1.9) on specially designed porous plates with microholes. The skin friction was measured directly by a

force balance and the boundary layer development was measured by a total pressure rake at the tailing edge of a test plate.

The free stream Reynolds number was 1.0(10)" per meter. The turbulent skin friction coefficient ratios (C,/C,o) of seven

porous plates are given in this report. Test results showed that the microblowing technique could reduce the turbulent skin

friction in supersonic flow (up to 90 percent below a solid flat plate value, which was even greater than in subsonic flow).

14. SUBJECT TERMS

Supersonic flow; Friction drag

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