Correlation of full-scale drag predictions with flight measurements on the C-141A aircraft : Phase 2: Wind tunnel tests, analysis, and prediction techniques. Volume 2: Wind tunnel test and basic data

Document

F L I G H T MEASUREME TEST, ANALYSIS ANI) BASIC DATA 8. Performing Organization Report No.

l a G . MacWilkinson, W. T . Blackerby and J. H. Paterson LG73ER0058 10. Work Unit No.

tion Name and Address 501-06-09-01 Lockheed-Georgia Company 11. Contract or Grant No.

flaiiariettrz, Georgia NAS1-10045 13. Type of Report and Period Covered Contractor Report Vational Aeronautics and Space Administration 14. Sponsoring Agency Code dashington, D . C .

I T h i s i s one of two f i n a l r e p o r t s .

A research program has been conducted t o determine t h e degree of c r u i s e drag c o r r e l a t i o n on t h e Volume C-lhlA a i r c r a f t between predictions based on wind-tunnel test d a t a , and f l i g h t test results.

2 contains information on t h e wind-tunnel test program and b a s i c aerodynamic d a t a on t h e C - 1 4 U wind- tunnel model used i n t h e c o r r e l a t i o n studies described i n Volume 1. The model w a s t e s t e d i n t h e NASA Langley 8-foot transonic wind tunnel.

18. Distribution Statement 17.

I

C - 1 4 l - h Wind-Tunnel Test Langley 8-Foot Transonic Tunnel Unclassified-Unlimited 20. Security Classif. (of this page) Domestic, $5.2: * For sale by the National Technical Information Service, Springfield, Virginia 22151 . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

T . . . . . . . . . . . . . . . . . . . . . . . . . . .

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a . Test617 . . . . . . . . . . . . . . . . . 6

ata . Test617 . . . . . . . . . . . . . . . 6

. Test591 . . . . . . . . . . . . . . . . . . . . . . . . . . 6

617 . . . . . . . . . . . . . . . . . . . . 6

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I I I

IN TRODUCTI 0 N

s information on the wind tunnel test program and basic aerodynamic

configuration used in the correlation studies described in Volume 1 . The

ny was contracted to prepare an existing C-141A high-speed wind

on and test i n the NASA g I ey 8-f oot t ra

wo principal phases, and to scheduling acquisition and analysis of data from both tests was e 1971 to March 1973.

WlND TUNNEL TEST

Facility

141A model was made i n the NASA Langley 8-f

ngle-return ciosed-circuit tunnel. The test sec

e sides measure 7.1 feet (2.16m). The upper wer wal I s are

,20 to 1.30, with

i t a variable test section Mach number from a locka e. The total pressure can be varied from a minimum

.2 x 1 8 N/m2) at all test Mach numbers, to a maximum

IO5 N / m q at transonic Mach numbers. The stagnation

cally controlled and i s usually held constant at720"F ntil the dew point temperature in the test section is reduced tion effects. The resulting maximum Reynolds number available t (19.7 x lo6 per meter) for this program.

Model and Instrumentation 275-scale C-141A model was used for the wind tunnel investigation.

model was supported accommodate a ne pport system, i n which the

ng blade from the I e and attached to a support sting below

ed and fabricated to locate the NASA adaptor blocks were

e, and two fuselage afterbody fairings of identical geometries were system and the quirements for testi

the same as

development program ctured and tested during iIs of the model dimen ch Center provided a 2.0 inch (5.08 c om total model loads. In addition, a sec

balance was used to measure blade loads. T standard fuse lag

scanivalves for measuring pressures from 67 s ic orifices on th

e fuselage. The layout of these orifices is she\

ive configuration with a dorsal s

atic pressures on the afterbody, and 18 internal pressures

e afterbody and one inside a wheelwell fairing.

e model angie-of-attack was fott servo acceler tion from which ndevco servo-ac el Support Configurations

etails of the principal model configurations are given in figures 2 t

2 shows the model mounted on the live blade and sting combination

model. Configurations 2, but with the blade removed.

In order to measure the load on the blade separately on configur R e model was attached directly to the dorsal strut and the blade attac r

balance. In addition to the blade lo e interference of the model on t

included i n this measurement.

tioned in a cavit Tests were conducted with different s

(f igure 6 ) , was

offset fairing o rogram. in config ine, and the bullet fairing was therefore not required.

~ t ~ o i ~ a l configuration was included during the test program, and designated

CIS obtained from configuration 8 by removing the blade so that measurements

the lower sting only in position.

Test Conditions and

s fixed on a l l model surfaces by 0.05-inch (0.125 cm) bands of ballotini

i n layer of lacquer. The choice of bead size was based on previous

eaience on transition fixing for high speed drag evaluation during the C-5A

ASA Langley has shown that i t i s possible to obtain

ue to the transition strip providin narrow, sparsely d istri-

. The roughness size for a given eynolds number i s deter-

2 600, the c ~ ~ t i c a l roughness ynolds number based on

t the top of the roughness, uk, and the kinematic visco-

e 2 , showed that for a eynolds number 2 3 x lo6

range of Rx values from 0.2 x 106 to

to 0.0045 inches (0.0114 cm)) no measurable as obtained with transition of the boundary layer occuring immediately nditions. These results suggested that for the C-141A

from 3 x 106 per foot (TO x 106 per meter) to 5 x 106 per

Rx based on ten percent wing MAC, a

chosen value of m) was required. Further, it was ncluded that no

ing was necessary as a result of t C-5A data. Transi-

ode1 components, except the fuselage and wheel well

e inches (1.855 cm) back from the leading-edges. The

well fairings were placed 2 * (5,08 cm) and

he nose of those components.

tests were completed with transition located further aft on the wing upper ly ten percent local chord ahead of the wing shock wave at cruise

ess size was 0.0054 inches (0.0137 cm). This technique was included

simulation of the interaction of shock and boundary layer at high Mach ions, where forward transition i s known to induce a premature rear

Ids numbers for this wing design. Flow visualization tests, wing upper

were conducted to locate the main shock position for this investigation.

eral, covered the ach number range from M = 0.600 to 0.825. In most

mber was 5.0 x 106/foot (16.4 x 106/mete or 3.05 x 106/MAC.

Scale effects were obtained on some configurations by testing also at 3.

foot (10 x lo6 and 13.12 x 106/meter). Six-component force m

at zero yaw over an angle-of-attack range from - 4 O (-0.0619

Selected configurations also included measurements of the fuselage afte cavity pressures.

A summary of the test program is given in tab1

tions, Nos. 4 and 8, showing the installation in the

photographs of figures 7 and 8.

Data Reduction and Corrections

Force balance data was reduced to coefficient form in the stabili

a reference wing area of 2.44 square feet (0.227 square meters), and a

chord of 7,328 inches (18.62 cm). Pitching moment was referred to a cog.

The accuracy of the strain gaged nce urements was b

ers quoted figure of &0.5% of the full-scale rating of the balance.

of the C-141A model force data based on maximum balance loads i s

esults, however, have sug ted somewhat better acc figures. For example, repeatability of drag in a given test series i s s

of 0.0001 to 0,0002 in CD. The accuracy of the angle-of-attack measur

to be - 1 5 0 . O 1 degrees (0 .OOO 1745 rad).

Corrections have been made to the ta for test section flow angufari~

(0.001 135 rad) downflow)) and wall constraint interference on lift, given

na = 0.07’4 CN, degrees

A set of empty tunnel buoyancy corrections were applied to the

results in volume 1 of this report. These had the effect of increasing t

from ACD = 0,00036 at M = 0.700 to 0.00032 at M = 0.825. All basic d

volume 2 of this report have been corrected for angularity and wall co

but not for bouyancy.

e effects were assumed to be negligible for this installation, based on experience

personnel at the Langley 8-foot facility, and consequently no corrections c) the test data.

BASIC AERODYNAMIC DATA

Genera 1

tation of basic aerodynamic data includes results from two principal test

was conducted i n June 1971 and Test 617 i n September 1972. A further

tests was completed i n March 1973, and designated as a continuation of the

ting in the 591 series, some doubt arose concerning the validity of the

m the basic tail-off configurations due to suspected model fouling.

ring the 617 series, when i t was found that the drag levels of the tail-on and

uration were approximately ACD ;= 0,0010 and 0 A006, respectively,

he 591 data. As a final check on the validity of the initial 617 data, the retest

il-off configuration during March 1973 confirmed these results. It should be

this volume indicate these discrepancies from both test

lasic data given in the subsequent analysis used only some limited component drag and trim

m the 591 test, and the main analysis was concentrated on the test 617

Flow Angularity. Test 617

ins moment and drag data for model upright and inverted runs are shown in and (c). These data have been corrected for the flow angularity and indi- ent over the full ach number ran: . C ;he tests.

Repeatability. Test 617

es of the degree of test data repeatability during the 617 series are given i n

In general, values of drag are repeated within ACD =0.0001 to 0.0002.

ows the repeatability of the afterbody pressures by a comparison of integrated 1 tests.

Complete Model Baseline ata. Test 697

Figures 11 (a-d) present some of the complete model baseline data w

the analysis and drag correlation studies in volume 1 . These data were obt

wing surface finish was prepared to a high degree of smoothness and a new applied prior to this main series of runs.

Comparison of these data indicate!

' change it-, drag level of the order of ACD = 0.0001 for this config~ration.

results for the model baseline is given i n figures 2 0 (a-d).

Support Tare and Interference

r configuration 7, the blade load measurements, are

12 (a) and ( b ) . Results from the tests with the dorsal sting arrangement, con

5 and 10 are given in figures 13 through 16. The sting interference eval

in figures 17 through 20, for configurations 1, 2, 3 and 4 respective

are given

for the central sting configuration, No. 9 are contained in figures 21 thro

Basic Data. Test 5

Selected results from the 591 series are presented i n figures 25 t

three horizontal tail settings are contained in figures 25, 26

Reynolds number range from 1.83 x 106/MAC to 3.05 x 10

through 30 for the complete model configuration, and those for tail-of

figures 31 and 32. The pylon-nacelle effects can be derived from the

33 and 34.

Aft Located Transition. Test 617

Results for the tests with wing upper surface transition moved back fror location to a position ten percent local chord ahead of the shock at cruise cion given in figures 35, 36 and 37.

Miscellaneous Data. Test 617

ta from which pylon-n ment are derived are given i n

cts of sting

ff, are given in figures

Afterbody Pressure Data

lar plots giving static pressure variation with fuselage station are shown in

54. These, in general, have been selected to support the discussion of

re drag effects, for the different model support configurations, given in

R E F E R E N C E S

1 . Braslow, A. L.; Hicks, R. M.; and Harris, R. V. Jr.: Use of Grit-

Layer-Transition Trips on Wind-Tunnel Models. NASA T N D-3579,

2. Palfery, J . G.: C-5A igh Speed Drag Reduction Study and lnvestigatio

Boundary Layer Transition on a 0.0226 Scale Model in the NASA-Ame

Tunnel. High Speed Series 4C and 4 0 . Lockheed- Georgia Company

LGlT6-1-40, April 1968.

TABLE I

MODEL DIMENSIONAL DATA

0.0275 Scale

43.656

4.675 axirnurn diameter, inches

9.362 W L 5.500

uiselage reference line

FS 6,336

airing, b

FS

FS 54.148

8.246 eness ratio ( I / D ~ ~ )

0.017 rea, square feet

FS 37.998

clot chord Peading edge

izontal Stabi I izer , H

0.365 16,554

n aerodynamic chord, inches

0 3 7 0

25.000 weep of 25% chord, degrees

23.885 ail length, inches

0.488 a i l volume coefficient

o i l section NACA 64A(010)010.5

MODEL DIMENSIONAL DATA (CONT.1

Nace.lle Pvlon, K 0.027 e

Inboard area, square feet inboard span, inches Inboard chord, inches Outboard area, square inches Outboard span! inches Outboard chord, inches Sweep of leading edge, degrees Nacelles.

Length, inches Inlet diameter, inches Exit diameter, inches

FS

Inboard Leading edge location

5L

WL

FS

Outboard leading edge location

W L

nboard toe-in, degrees oard toe-in, degrees

Vertical Stabilizer, V

Area, square feet ean aerodynamic chord, inches

ODEL DIMENSIONAL DATA (CONT.)

0.0275 Scale

Stabilizer (Cont.

__.__I__

1.208

0.617

35.000

21.065

0.049 ica I tai 1 volume coefficient

2.440

52.689 €:an aerodynamic chord, inches 7.328

7.900

0.373

23.734 of 25% chord degrees (inboard)

FS 25.378

?h mean aerodynamic chord location

WL 7.248

B 10.818

25.025 eep of 25% chord, degrees (outboard)

4.891 * 584

-5

0.941 ~ ~ e ~ r a l of 25% chord, degrees (inboard

1.195 ord degrees (outboard)

10.91 1

6.419 mk station chord, inches (inboard

6.487

WS 11.138

ws 11.734

station I oca tion (outboard)

MODEL DIMENSIONAL DATA (CONT.)

0.0275 Scale

Wing (Cbnt.)

-

Airfoi I section

Root NACA 0013.0-1.10-40/1,575 (MOD) a 0 = 0 . 8 (MOD) C1. =0.153

Inboard break NACA 001 1.2-1.10-40/1.575 (MOD) a . = 0.8 (MOD)

C1. = 0.194

NACA 001 1 .O-1.10-40/1.575 (MOD) a . = 0.8 (MOD)

Outboard break

c1. = 0.201

NACA 0010.0-2.20-40/1.575 (MOD) mean line 1/2 (NACA 66 at

Tip

C1. = 1 .O-NACA 230 at C1 = 1 .O C1. =0.452)

I I a1

Antenna Fairing, Z

Leading edge location FS 26.208

FS 33 .&8

Trailing edge location

G2 I

Wheelwe I I Fairing ,Z

11.110

Length, inches

2.922 Maximum frontal area per side, square inches

5.760

Fineness ratio

FS 21.835

Leading edge location

w 7

Wing-Fuselage Fillet, Z

Leading edge location FS 15.286

FS 30.970

Trailing edge location .

v) cy

"!

m .

v) .

\I

cv

c) .

v) L hl 9.

t-

Y

?

r ) * A

i

c '4 c '4 @ > . .

3 \ 3

+

* *

Y Y

U U I Y O 5 ;

-It-

/

h c a 3 I=

.-

c U L 0)

.-

cc f

v

c L Q Q m

TABLE 2

WIND TUNNEL TEST SCHEDULES

(A) TEST591

Description

9 12 W7 19 7 a1 G21

B W Z K N Z Z 2 1/8 7

5 . O

01 nu IO IS t o I! no (Inverted)

2 16 15

5 .O

9 12 W7 19 7 a1 G21 4 6 8 8

B W Z K N Z Z D V H . b 1 5 . O

-I II 11 II so 11 i n I o n I 11 nt I!

5 . O

nu II o n i n 11 IU i t II nn OB I!

5 . O

IO nu IO II I I 81 !I II I O II II

5 . O

9 12 W7 19 7 a1 G21 4 6 8 8

B W Z K N Z Z D V H b 4 5 . O 55 54 53

to ii ti It i t ii ii i! i n H8 ii

4 5 . O 62 61 t

4-1

69 68 67

9 12 W7 19 7 a1 G21 4 6 8 8

- s

B W Z K N Z Z D V H , b - 4 73 72 -

5.0

II n i n i i n II ii i t i n i n nt t i

- -

4 . 0 74 75 -

01 II !I 01 00 I1 (I It #I I! iI

- -

4 3 . 0 81 80 -

- -

9 12 W7 19 7 a1 G21 4

3.0 85 84 -

B W Z K N Z Z

nu nn i t O B II II st

- -

5.0 86 87 -

9 12 W7 a1 G21

93 92 I- . . m -

B W Z Z Z

4 5.0

B9z G2 1

97 96 I- - -

5.0 (B) TEST 617

S u pport R d F t

escri ption x Config.

Zw 7( I nverted) 2 5 . O - 45 44 43 - 42 -

2 5 . O - 51 50 49 48 47 46

'Zal ZG2' Zw7B lade Sea I 7

5 .O 7 6 5 4 3 2 1

7 5 . O 12 - 11 - 10 - 9

8 5 . O 19 18 17 16 15 14 13

6 5 . O 26 25 24 23 22 21 20

5 5 .O 33 32 31 30 29 28 27

5 5 . O - 35 - - 34 - 36

10 - 39 40 - 38 41 37

5 .O

7'Za 1 ZG21ZW7D4V6H8,b8

4 5 .O - 60 59 58 57 56 55

-

4 5 . O - 69 68 67 66 65 64

3 . O 73 72 71 70 - -

3 . O 77 76 - 75 74 - -

3 . O 82 81 - 80 79 - -

4 5 . O 8 6 8 5 - 8 4 8 3 - -

4 5 . O - 93 92 91 90 8 9 -

5 . O - 126 125 124 123 122 -

- 121 120119118117 -

5 .O

5 . O 1 - 1161151141131112

(5) TEST 617 CONTINUED

Run -

R d F t

Support x Description Config.

9 12 19 7 a1 G21 W7 4 6

- 99 98

B W K N Z Z Z D V 4 "5.0

-1 I I II I 1 11 11 11

- 10510

4 *5.0

I1 I I II II I I 11 I I I 1 II I1

- lll11011

4 *5.0

9 12 19 7 a1 G2lZW7

- 134133113 -

5 W K N Z Z 9-1 f 5 .O

- 1 II 11 11 11 11 I1 I I I1 I 1 II II

9-2 If 5 .O

I1 I t I f I I II I I II II I 1 I t 11

9-3 .t/ 5 .O

II 11 I t II 11 I1 I t II I f II II

5 .O

9-4 f

9 12 19 7 a1 G2lZW7

B W K N Z Z 1 5 .O

11 II II II II I 1 II

2 5 .O

11 11 I I I t I t I I I I

5 .O

II I I I I II I t II I1

5 .O

* Transition Strip Moved Back to 10% Local Chord

Ahead of Shock.

+ Central Sting -

c v) a , I - U a , c

t

> S w Eu -ct .

d

I h W 0 0 0 0

cv

9 -3- cv *

d

0 d

Q

0 0 0 9 CJ e l n i - * x L .

U (I) I=

.-

c S U 0 0 0 0 cu 9 -4- C Y

0 d 0

R = ~ . 7 0 0

0 Tes

Test Test

Comparison of Test 591 and

617 Afterbody Pressure

Figure 10. Concluded I1 C> a : , * .

C : , II .

c ; c I I

-5

( I . . . .

J S C> 0 - "r 3 II m .

C> II C> C> P . .

C> 0 0 0 0 0 00 \8 04 " .

0 0 I I W TCI a ) S

.-

.I- S W 0 0 0 0 hl 0 0 I I-.-

-

2 4

-4 -2 0

Degrees - a . Runs 1-7.

( r ~ ~ ~ igure 12. Concluded.

d" e a l i

2 E

0 d 0

-2 -1 9 , 3. Continued, t I I- d- c I S

.-

+ U L I ) U c U Q V

.-

v) U m -Fs

d

I I Figure 14. Continued.

L n c (1- c c

4 6

';"

cu VI I= Cd .

U I

n

U n at W

d

c - .

a I 0.

0: N I h N

'$

0- 0"

d d

N l n o % .

a , 0 0 a , U I P- C a, L W -3 h a , W v) a , a , L 0) a , .

n

h c c v) a , I--.

d

c c Lo cu

P

h U c c3 U

d

n

c Lo

.-

ac h E U S x 0 cs T I Lo I h L * -J

v

n u 0 U W b .

.L a , L m

.-

L .

d

A V I- N I v) a ,

-7

a3 CT) v) c oc: U I e U e , c

.-

L 0 c OI 0 U 0 0 0 0 cy n LL 0 0 CJ 0 CJ

P I t

N r; I- \Q t c l n a , I - - .

v) P a , a , S L 0)

.-

c et

n

E

.-

q_ S

u

0) S

.-

c U v,

d

I .

J Ir) U c co U h n

z

U 0 v Ir)

n

n .

d

.-

v) U m n I I r; c a , 0)

.-

LL (u .- I hl v) 0 I a , a , L m a , I

a

r; c- ctt L d-0 h l c o ....

o n

.d 0"

P

I I

z

U h M 9 c v) a , I- (u S P

.-

E

23)

.-

u- r : W 23) S c- a ,

f

23)

.-

d

I f CI U I * a , S --E S U

d

z

d

a , L 0)

.-

ILL OD c- I I Ir, OD R I I Q Q egrees e IFigure 19. Sting ~ o ~ ~ i g u ~ ~ ~ i ~ ~ 3. Test 617.

U e , S

.-

c C W a m a , I .

c \ I N

-e - ' -

'do"

I I

z

0.80 0.60

c

0.40

0.20 -0.2c

Figure 20. Sting Configuration 4. Test 617.

c I t :

do'

c U J S c 0 I ) W h W 0 0 hl

0 2

0 d

I I

z

OI 0 - I n a , 0) n?: LL

.-

In a u) S

.-

.I- v, u) S

.-

U LL 0 .I- M rc S Gd .

tl .

U c I -I

s

.-

n U v I= x U L

v

.-

In U no LL S * 0 0 0 W W egrees

- 8

Figure 22.

M = 0.700 0.775 0,85

Figure 22. 0nti S dc d V I c3

v

rrq U W a , L 0)

.-

u- Figure 23. Continued.

d

m c VL S CL L3 U n u W m cu e , L 0)

0 .-

0 hL h ll \8 egrees

cyF: RL

d

e c9

f

v) c cr:

.-

. a - S U Q)

f

.-

U e !

J Ln h Ir) R h .

Ir) h h -43 * t I- .

-a at I) S

.-

c S 0 0 0 0

0 : i d &

co 9 -4- c \ l 0 0 0 0 m n a , W .

a , TI

-

v

0 0 \cp

0 d

v) C oc: I .

T I a , C

.-

c C V 0 0 0 00 *

d

0 0 I c c v) a, I-

d d

Q Q .

N '5) .

U I U

-i

.-

J- r :

2 0

d

r: cv 1 ' 4

.d

h I) Q 0 'i;r cu Q 0 Q Q * c3 W I -I 0 W -4- 0 e W W

d

n

cu W L X i l R I- @ a,

f

.-

0 0 N O O h e .

0 0 0 0 03 N

0 0"

. .

0 0 J

d

S W a ,

f

.-

U 0 0

cu cu

ir .

I J 0 0 N O O R

d d

0 0 N O 0 9 f a * 0 0 0 0 0 0 4 -3 II

d

0 d

I Ti- I;

n

p;;r pu 0 Q 0 0 0 0 0 03 9 OI h( 0 0 0 .

J W I

d

c3 (u

d

FR

(a) CL - a . Runs 86-89,

ic Aero~ynamic ai I -off, Test 59

= 5 x 10 6 /foot

T I a , S

.-

c S

-

M a , a n c

0 0 0 0 t

Q FRL uta. raij.-Off. Test 591.

= 3 x 10 /foot

c

Figure 32. Continue

0 0 -0 0) S

.-

c S U a

.-

E

S x 7 s & 0 0 0 0 0 hl d cu,

0 d d

ua r_ cx)

0 n

0 0

= 0.700 = 0,785

-0 0 0 C

M = 0.600

0 775 0 785

s 70 - 73

R

* 0 0 c- "- I- * .u) U M 7s S 0 0 CU

0 d

9 '

U OI a , S

.-

.I- o \ S v) S W oc

ui

M .J a , W L I 3 u)

.-

P LL U h U W - . .

0 0 .

cu

(u0 d

I I cr) 0 0 0 9 d O d e- & L n c cu -b.

do"

N r_ * hl 0 0 0 d cu * 0 0 * U a , P- U S _.

OL S V -I U I c3 L3 a , L U n u)

.-

a , W .u

.-

c r Q) L v) != u) (I- L L cd II W v) S 0 0 N o

o n

00'

v) S 3 U rL a , S

.-

J c

. =

U I U

n

v

n U a , Y u)

.-

n

*

cu 0 0 0 0

hl 'd

hl 0 0 I

1 45

0 0 0 0 0 -4- cu a3 9 0 0 0 U egrees a

F

Sting Configuration 2.

m 0 4 -0

3 O

1 5

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