j .

NASA

Technical

Paper

2439

July 1985

Iq/ A

Nacelle/Pylon/Wing Integration

on a Transport Model With a

Natural Laminar Flow Nacelle

Milton Lamb,

William K. Abeyounis,

and James C. Patterson, Jr.

NASA

Technical

Paper

2439

1985

RIASA

Nalional Aeronauhcs

and Space Adm,n_stration

Scientific and Technical

Information Branch

Nacelle/Pylon/Wing Integration

on a Transport Model With a

Natural Laminar Flow Nacelle

Milton Lamb,

William K. Abeyounis,

and James c. Patterson, Jr.

Langley Research Center

Hampton, Virginia

Introduction

Extensive efforts are underway to develop efficient transport aircraft with reduced direct operating costs (ref. 1). Much of this effort is focused on

providing extensive laminar boundary layers on the

various aerodynamic surfaces of the aircraft (ref. 2).

Laminar boundary layers can be maintained naturally by favorable pressure gradients. Pressure gradients favorable for laminar boundary layers are not

favorable for lifting surfaces because of the rearward position of minimum pressure coupled with the

trailing-edge pressure recovery. In addition, many of

the lifting surfaces on transport aircraft have significant leading-edge sweep angles that generate spanwise flows which are unfavorable to maintaining laminar boundary layers. Since nacelles are not required

to carry lift and have negligible leading-edge sweep,

they may be designed to have favorable pressure gradients for maintaining laminar boundary layers on a

large portion of their surface. Nacelles could, therefore, be a candidate for natural laminar flow. As

shown in reference 3, a reduction on the order of lV2

to 2 percent of total aircraft drag at cruise may be

achieved with little or no weight penalty by maintaining laminar flow on the nacelle.

An earlier study (ref. 3) was conducted in the

Langley 16-Foot Transonic Tunnel to determine the

extent of laminar flow obtainable on an isolated nacelle designed to have a favorable pressure distribution over approximately 70 percent of the fan

cowl length. Integration tests conducted on a fullspan model would involve nacelle sizes considerably

smaller than the isolated nacelle previously tested.

The smaller nacelle with its lower Reynolds number

is expected to have natural laminar flow over a larger

portion of the nacelle length.

The main purposes of this study were twofold:

First, to determine if the philosophy used in the

design of this smaller nacelle, as well as the isolated

nacelle, would result in laminar flow over that portion

of the nacelle for which it was designed; and second,

to determine if integration concepts had any adverse

effect on the extent of laminar flow on the nacelle. In

addition, this study determined the level of nacelle

drag reductions that may be obtained with flowthrough laminar flow nacelles installed on a highwing transonic transport configuration. The effects

of fixed and free transition as well as longitudinal

position and pylon contouring were also investigated.

Extensive static-pressure measurements were taken

on the nacelle and wing. The investigation was

conducted in the Langley 16-Foot Transonic Tunnel

at free-stream Mach numbers from 0.70 to 0.82 and

angles of attack from -2.5 ° to 4.0 ° .

Symbols and Abbreviations

A area, in 2

Are f wing reference area, 529.590 in 2

BL butt line of model (lateral dimension), in.

b wing span, 63.121 in.

CD drag coefficient,

qvc_ref

AC D installed drag, CD,WBNP -- CD,WB

CL lift coefficient, Lift

Cm pitching-moment coefficient, Pitching moment

qoocAref

Cp pressure coefficient, (p - pc_)/qoo

c chord measure in wing reference plane, in.

c W chord of wing at intersection with pylon,

9.972 in.

mean geometric chord, 9.107 in.

Fin nacelle internal drag, lb

FS fuselage station (axial dimension measured

from model nose), in.

k constant in figure l(c)

l nacelle length, 12.770 in.

M Mach number

m mass, slugs

NBL nacelle butt line, in. (fig. 3)

NWL nacelle water line, in. (fig. 3)

NS nacelle station (axial dimension measured

from nacelle lip), in.

p local static pressure, lb/in 2

Poo free-stream static pressure, lb/in 2

qoo free-stream dynamic pressure, lb/in 2

r nacelle radius, in.

V velocity, ft/sec

WL fuselage water line, in.

WRP wing reference plane (fig. l(a))

x local axial dimension, in.

XLE axial distance from pylon leading edge for

defining cap leading-edge section shape, in.

(fig. 4(a))

leading edge of nacelle, in. XLE,N
XTE axial distance from pylon trailing edge for

defining strut trailing-edge section shape,

in.

y local lateral dimension, in.

z local vertical dimension, in.

c_ angle of attack, deg

rl y/(b/2) for wing pressure locations (fig.

l(c))

¢ circumferential angular measurements for

nacelle orifice locations, deg (fig. 2)

Subscripts:

exit nacelle exit

o nacelle inlet

Model components:

B body

N nacelle

P pylon

W wing

Experimental Apparatus and Procedure

Wind Tunnel

The experimental investigation was conducted in

the Langley 16-Foot Transonic Tunnel. This tunnel

is an atmospheric, single-return wind tunnel with

continuous air exchange and is capable of operating

at Mach numbers from 0.20 to 1.30. A detailed

description of the tunnel is presented in references 4

and 5.

Model and Support System

The 1/24-scale model, representative of a widebody transport, is shown in figure l(a), and a photograph of the model with laminar flow nacelles in

the forward position is shown in figure l(b). The

model, which had a high wing with supercritical airfoil sections, was mounted on a sting-supported sixcomponent strain-gage balance. Details of the fuselage, wing, and wing pressure orifice locations can he

found in references 6 and 7. For these tests, the wing

of reference 6 was modified to reduce the curvature

of the wing lower surface in the region around the nacelle pylon (see fig. l(c)) to reduce the flow velocities

in that region.

A low-order panel method (ref. 8) in conjunction

with a specific compressibility correction technique

(ref. 9) was used in designing the long-duct flowthrough laminar flow nacelle to solve for the basic

installed flow field, with no consideration for laminar

boundary-layer stability theory. The desired nacelle

shape was analytically modeled into the wing-body

flow field, and the resulting pressure gradient of the

nacelle was checked for separation as well as for peak

surface Mach number. If either the pressure gradient or peak Mach number proved unsatisfactory, the

nacelle contour was modified and the entire process

was repeated until a satisfactory shape was obtained.

The final nacelle contour, shown in figure 2, had a

highly polished surface and a favorable pressure gradient over a significant portion (60 percent) of its

length. As shown in figure 3, the nacelle was installed

in a forward position (top sketch) or a rearward position (bottom sketch) on the symmetrical pylon and

in a forward position (middle sketch) on a contoured

pylon.

Details of the symmetrical pylons are shown in

figure 4(a) and details of the contoured pylon are

shown in figure 4(b). Because flow acceleration

around pylons is a major contributor to the high

velocity peaks on the wing lower surface, the pylon

was contoured in an attempt to reduce these high

velocities. The low-order panel method discussed

earlier was used to design the contoured pylon using

the wing/body/nacelle flow field obtained for the

laminar flow nacelle in the forward position.

Instrumentation and Data Reduction

The model aerodynamic force and moment data

were obtained by an internally mounted six-component strain-gage balance. The model surface static

pressures were measured by scanning electrical straingage transducers located in the model nose to reduce

the lag time required between data points. Sting cavity pressures, measured by remotely located straingage transducers, were used to correct the cavity

pressure to the free-stream static pressure. Based

on repeated runs, the accuracy was equivalent to 1

count of drag (C D of 0.0001).

All wind-tunnel parameters and model data were

recorded simultaneously on magnetic tape. Except

for scanning valve pressures, averaged values were

used to compute all parameters. The model angle of

attack was computed by correcting the support strut

angle, both for sting deflections based on balance

loads and for tunnel upflow determined from inverted

model runs in a previous tunnel entry (ref. 6). Stingcavity pressures were used to correct the longitudinal

balance components for pressure forces in the sting

cavity. Nacelle internal drag corrections were made

by using internal static pressures to determine the

mass flow for a one-dimensional flow calculation; that
is,

Fin -- mVexit - rnVo Jr (Pexit - po)Aexit

Forces and moments were transferred to the model

moment center, the quarter-chord point of the mean

geometric chord on the model water line.

The turbulent skin-friction drag for the nacelles

was calculated by the method of Frankl and Voishel

for compressible turbulent flow over a flat plate.

Laminar skin-friction drag for the nacelles was calculated by the method of Blasius. No corrections were

made for grit drag.

Tests

This experimental wind-tunnel investigation was

conducted in the Langley 16-Foot Transonic Tunnel at free-stream Mach numbers from 0.70 to 0.82

and Reynolds number from approximately 2.5 x 106

to 3.0 x 106 based on the mean geometric chord of

the wing. The model angle of attack was varied

from -2.5 ° to 4.0 ° . Boundary-layer transition on

the model was fixed using a grit-transition-strip procedure (ref. 10). A 0.1-in-wide strip of No. 100 carborundum grit was attached 1.0-in. behind the nose

of the fuselage. Strips of No. 90 and No. 80 grit

were applied on the upper and lower wing surfaces

(see fig. 11 of ref. 7) in a rearward position in order

to match the boundary-layer thickness at the trailing

edge of the wing (ref. 11). In those cases where it was

desirable to fix transition on the nacelles, a 0.10-in.

strip of No. 120 grit was placed 1.05 in. (x/l = 0.08)

and 0.375 in. (x/l = 0.03) aft of the nacelle lip on

the external and internal surfaces, respectively. The

text matrix is shown in the table below.

Angle of

Mach attack, Nacelle

number deg position Pylon

0.70 to 0.82 -2.5 to 4.0 Forward Symmetrical

0.70 to 0.82 -2.5 to 4.0 Forward Contoured

0.70 to 0.82 -2.5 to 4.0 Rearward Symmetrical

Nacelle

transition

Free & fixed

Free & fixed

Free & fixed

Results and Discussion

The basic aerodynamic characteristics of the different nacelle configurations are presented in figures 5

to 7. The wing-body configuration with original wing

leading edge is presented to show the effect of nacelle

installation with wing leading-edge modification and

the effect of free and fixed nacelle transition. The

addition of the nacelles reduced the lift coefficient

and increased the drag coefficient. Fixing the nacelle

transition caused a significant increase in drag coefficient while lift and pitching-moment coefficients were

affected only slightly, if at all.

Laminar Flow Design

The purpose of this study, as discussed earlier,

was aimed at determining if a nacelle designed to

achieve laminar flow over a large portion of the nacelle length could be integrated into a transport (nacelle/pylon/wing integration) configuration and preserve the extent of laminar flow achieved on the isolated nacelle. For the configuration with the nacelle

in the forward position (position which should produce the least interference), the drag reduction due

to laminar flow (increment between fixed and free

transition) is approximately 9 counts. (See fig. 5(e).)

Calculations of the skin-friction drag for the nacelle

with fixed and free transition (indication of the extent of laminar flow achieved on the nacelle) indicate

that the drag reduction expected for a nacelle of this

size with 60-percent laminar flow (observed in sublimation runs) is approximately 9 counts. Thus, it

would seem that the ability to design a nacelle to

achieve laminar flow has not been significantly altered by integration.

Effect of Nacelle Position

The effect of longitudinal position of the nacelles

on the longitudinal aerodynamic characteristics is

shown in figure 8. The nacelles were installed on

symmetrical pylons such that the ratio XLE,N/CW

was --1.20 or --0.90. (See fig. 3.) A large reduction in

lift and increase in drag is associated with installing

the nacelle in the rearward position (XLE,N/C W =

--0.90). This increase in drag and the associated loss

in lift are revealed by the wing pressure distributions

presented in figure 9. While there is some decrease

in upper surface pressure coefficients near the wing

leading edge (see fig. 9), there is a large decrease

in the lower surface pressure coefficients. This large

decrease is possibly due to flow acceleration between

the nacelle/pylon/wing and fuselage as the nacelle

is moved rearward, thus resulting in shock-induced

boundary-layer separation and the increase in drag

shown in figure 8.

Pylon Contouring

The effect of pylon contouring on the longitudinal aerodynamic characteristics is shown in figure 10.

This contoured-pylon configuration exhibits less drag

than the configuration with symmetrical pylons. The

increment in drag between the symmetrical and contoured pylons increases with increasing Mach number, not unusual since the contoured pylon was designed for M = 0.80. The wing pressure data inboard
of the pylon (fig. 11) show that the contoured pylon

allows a better pressure recovery on the upper surface and reduces the lower surface velocities, thereby

producing lift and reducing drag as noted in figure 10.

Installed Drag

In figure 12, the installed drag coefficients (AC D =

CD,WBNP -- CD,WB) are presented across the Mach

number range for the configuration cruise condition

(C L = 0.45). The curves were obtained from interpolated values of CD at CL = 0.45. Configurations with the nacelles installed in the forward position (figs. 12(a) and (b)) show only slight changes

in installed drag over the Mach number range. The

contoured-pylon configuration with free transition

(fig. 12(b)) exhibits less drag than the symmetricalpylon configuration with free transition (figs. 12(a)

and (c)). The installed drag for the nacelle-forward

symmetrical-pylon configuration with fixed transition remains essentially constant over the Mach

number range, whereas the installed drag for the

contoured-pylon configuration with fixed transition

decreases as Mach number increases. For the configuration with rearward-installed nacelles (fig. 12(c)),

installed drag increases rapidly with Mach number

and the increment between free and fixed transition

decreases rapidly with Mach number.

Nacelle Pressures

Figures 13 to 15 show the nacelle pressurecoefficient distributions at C L _ 0.45 over the Mach

number range. These pressure distributions will

probably explain the extreme differences noted for

the installed drags. For the configurations with the

nacelles in the forward position (figs. 13 and 14), the

pressure-coefficient distribution over the Mach number range is similar, with only a slight increase in

peak pressure coefficient with increase in Mach number. However, for the configuration with the nacelles

in the rearward position (fig. 15), peak pressure coefficient increases rapidly with Mach number, particularly on the inboard side in the region close to the

nacelle-pylon juncture. This high peak pressure is

essentially the same for free and fixed transition and

tends to override the pressure differences (i.e., a local disturbance due to transition strip) resulting from

fixed transition at low values of x/l. This probably

explains why the free and fixed transition installed

drag approaches the same value at the higher Mach

numbers.

Comparison of Installed Drag at M = 0.80

The installed drag coefficients at M = 0.80 and

C L = 0.45 are compared in figure 16, with the drag

broken down into skin-friction drag and interference

drag. For free transition, laminar flow was assumed

to exist over 60 percent of the nacelle length (based

on design and on sublimation pictures). Since only

the nacelles were allowed to have free transition, the

differences between the same nacelle-pylon configurations with free and fixed transition should be essentially equal to the skin-friction differences between

turbulent and laminar flow. Such is the case except for the contoured-pylon configuration of about

4 counts higher. This increase in drag is attributed

to interference drag.

Conclusions

An experimental investigation has been conducted in the Langley 16-Foot Transonic Tunnel to

determine the effects of installing flow-through laminar flow nacelles on a high-wing transonic transport

configuration. The effects of fixed and free transition

as well as longitudinal position and pylon contouring were obtained. The results of this investigation

indicate the following:

1. The ability to achieve laminar flow on the nacelle was not significantly altered by nacelle/pylon/

wing integration for the nacelles on symmetrical pylons.

2. The increment of drag between free and fixed

transition was essentially the calculated differences

between turbulent and laminar flow on the nacelles.

3. With the nacelles in a forward position, the

installed drag for the contoured pylon was less than

that for the symmetrical pylons.

4. The contoured pylon had increased interference drag with fixed nacelle transition.

5. The installed drag for the nacelles in a rearward position was greater than that for the nacelles

in a forward position.

6. The increment between free and fixed transition for the rearward nacelle decreased with an increase in Mach number, probably due to the high

peak pressures which overrode the pressure differences resulting from fixed transition.

NASA Langley Research Center

Hampton, VA 23665

April l, 1985

References

1. James, Robert L., Jr.; and Maddalon, DaJ V.: Airframe

Technology for Aircraft Energy E_ciency. NASA TM85749, 1984.

2. Wagner, Richard D.; Maddalon, Dal V.; and Fischer, Michael C.: TechnotogyDevelopments for Laminar Houndary Layer Control on Subsonic Transport Air-
the Fluid Dynamics Panel Symposium on Improvement

of Aerodynamic Performance Through Boundary Layer

Control and High Lift Systems (Brussels, Belgium), May

1984.

3. Younghans, J. L.; and Lahti, D. J.: Analytical and Experimental Studies on Natural Laminar Flow Nacelles.

AIAA-84-0034, Jan. 1984.

4. Corson, Blake W., Jr.; Runckel, Jack F.; and Igoe,

William B.: Calibration of the Langley 16-Foot Transonic

Tunnel With Test Section Air Removal. NASA TR R423, 1974.

5. Peddrew, Kathryn H., compiler: A User's Guide to the

Langley 16-Foot Transonic Tunnel. NASA TM-83186,

1981.

6. Lee, Edwin E., Jr.; and Pendergraft, Odis C., Jr.: Installation Effects of Long-Duct, Pylon-Mounted Nacelles on

a Twin-Jet Transport Model With a Swept, Supercritical

Wing. NASA TP-2457, 1985.

7. Abeyounis, William K.; and Patterson, James C., Jr.:

Effect of Underwing Aft-Mounted Nacelles on the Longitudinal Aerodynamic Characteristics of a High- Wing Transport Airplane. NASA TP-2447, 1985.

8. Maskew, Brian: Prediction of Subsonic Aerodynamic

Characteristics--A Case for Low-Order Panel Methods.

AIAA-81-0252, Jan. 1981.

9. Dietrich, D. A.; Oehler, S. L.; and Stockman, N. O.:

Compressible Flow Analysis About Three-Dimensional

Wing Surfaces Using a Combination Technique. AIAA83-0183, Jan. 1983.

10. Braslow, Albert L.; Hicks, Raymond M.; and Harris,

Roy V., Jr.: Use of Grit-Type Boundary-Layer-Transition

Trips on Wind-Tunnel Models. NASA TN D-3579, 1966.

11. Blackweil, James A., Jr.: Preliminary Study Effects of

Reynolds Number and Boundary-Layer Transition Location on Shock-Induced Separation. NASA TN D-5003,

1969.
J\,,,_!

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i

N

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d

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tm_'

0 ,'_

6
-0o
f[__. ____._led Modi _ ,_

IOrigin:/-/- __ _-x/c- 0. 28-,.I

r x/c- 0.56

Typical chord section

WRP

'.°I

k

0 • 154 .352 .440

At any local station: (z/C)modified• (z/C)original- (_ z/c) (10

Negative sign on AZ/C values indicate undercut

Positive sign onAz/c values indicate material added

xlc Az/c x/c Adc xlc ,_dc

0.0(0) 0.0(0)0 0.120 -0.0064 0.300 0.0013

O.002 -0. 0044 O.140 -0. 0055 O.380 O.0013

0.005 -0. 0064 0.160 -0.0045 0. 400 0. 00l 3

O.010 -0.OOlg O.180 -0.0035 O.420 O.0012

0.020 -0.0092 0.200 -0.0027 0.440 O.OOll

O.050 -0.0095 O.220 -0.0019 O.460 O.0008

O.040 -0.0094 O.240 -0.0012 O.480 O.0006

O.050 -0.0092 O.260 -0.0005 O.500 O.0004

O.060 -0.0089 O.280 O.0000 O.520 O.0003

O.070 -0. 0085 O.300 O.0004 O.540 O.0001

0. 080 -0. 0081 0. 320 0. 0007 0. 560 0. 0000

O.100 -0.0072 O.340 O.0011

(c) Wing leading-edge modifications.

Figure 1. Concluded.
§

CD

_• II /

o

Z

Z

CD

"I 2

r--

r-

• l

r_

I

°_

o

I

1

-.L

q,Q

F_b .L

_7

" _0_

0

• °

rv_

r-

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o

m_

w _

8

rf-

• o . • • ..... • o °

7_

.... o ........

e-

_2

q_

c_

.o

r_

o

c,i

f--4

g
FS 30._84 1.00o

l

t--l_ NBL 0.000

/

P ion FS 30.284 Wing 0.5°

. Nacelle y FS 26.5z8 I / FS 36.500,_

,_./_.--L_o / I / .....

.---_ NWL 1.547

Symmetrical pylon

Nacelle lorward

--NWL 0,0_

Contoured pylon __

Nacelle lorward

XL£,N/cw - -t.ZO

NS 0.0(]O

FS 26.528. FS _.284

_---_ /" 0.5°_).2_

- ---_NWL ).541 WRP

t NWL 0 00O

J

NS }.2.770

Symmetrical pylon

Nacelle rearward

FS 30.284

FS 26.528

FS _5.500

I ".... _ .. Nwl4

XLE.N/CW - -0.90 . _ O.S WRP

_-_-- ---_-- - NWt 1 547

t i NWL O.

NS 0.000

NS 12.770

Figure 3. Nacelle test matrix. Linear dimensions are in inches.

10
Cap

Base

strut

Tangent point r

0.800

t.ypicat oyroq sedio,_

FS 26,528 /

CaI_l) NWt 2.250 -- "

Ba_l}j _ ........ "--Strut

NS 12,170

FS 26.528

X1

FS 6.528

6.150

Rad.

NS 8.{_

V

_4.710_ _ X2

I

I C_nlL2r'7t iOtnat c h .... lie

NBL

0.000

0.48

Ellipseof O.e_O minor axis

0.4_ major axis

Basell) I 4.230

/_/ L Basel2)

j.:-

"Contour to wing lower surface coordinates

Coordinates are referred Irom FS 26.528 for x = 0.000 and WL 3.250 for z = 0.000

FS 26.528 /-

! _ _s_._

l.sso[ --_

Pylon leading-ed_ section

XLE Y

0.000 0.000

0.020 0.060

0.030 0.073

0.050 0.091

O. 100 O. I24

0,200 0.173

0.300 0.211

0.400 0.243

0.600 0,294

0.800 0.332

l.O00 0.361

1.200 0.382

1.400 0.395

1.600 0.400

L

Pylon traifinQ-'edge section

XTE Y

0.000 0.008

O.501 O.072

1.003 O.129

1.504 0.200

2,005 O, 2_

2._7 O. 334

2.841 O. 375

3.108 O. 393

3.342 0,_

wi _a-gcoordi nates

0.000 i -0.038

0._'0 [ -O.OO9

0.050 I -0.131

O,lO0 , -0.179

o.I_i -o.2_

0 399 -( 18

O. 598 -{ ,5

O.798 -{ 2

O. _l -0. 453

I. 396 -0. 5lg

I.795 -0. 573

2.194 i -0.613

l Pylon configuration Co nents -

Nacelle forv,erd I Cap411* Basett) + Strut

Nacetle rearv,,ard _L Cap_2) + Base(2) + Strut

(a) Symmetrical pylon.

Figure 4. Details of pylons. Linear dimensions are in inches.

U
I rllx_rd

fs 30.284

1.O_ BLI1.067

L J_

\

NBL 0.000

NWt 2.2_--

rs2i.sn /j -

I

NS 8.060

FS 26.528

NS ll,g_

I

_L 3._ I_

NS 12.770

Conlour to match

laminar flo'_ nacelle

Contour to win9 Iox_er surface (:_ordin_les

._ given in table ol figure _il)

O._io / NWI_ 4.000

._ - 3/

" ' _-_dWL 3. ;'_

FS 36.5O0

f

1// NWL 1.

!s 33.053

r -, -

i Coerdimles in aircrall reference system at Wl

! In I OU_ in in : out OLII i tn "in olal , Jo_

i zo,6_ n.328i 2o._33t n._ I 22.z_[ eL_F]f _.2_ H.7_I 2_42_i l_ _ i 26_l IU_

23.2_I i 11,226 ! 23.295 . 24 844 ii 232 24 B7 12.D_, 26.812 }.316 26._2 )2.119

llg 

25, J1_1 1134,t 25_ 12.45 26._ 11.3_i26/'_ I 2 92 28.60'5 113_" P&614 _ 121_

26.631 i I1452 i 26.6_ I 12-2581 28.066 I 11.467 I 28.096 I 12.267 I 29.79[ 11._ I zg._[_,i 12.1,17

27.3_I _1.s[6I 27._81,2._12I zs.zo+I 11.4+21_1+[ t2.z6zl _0._ ]1.315 i _.3_ / 12 108

27 g98 1 549 21 9_ 12.MO _.382 11.454 28.367 12.241 )0,985 11,?q7 30.977 12 06)

28.6/0111.568 28.681 12.325 13_.017 111.437 30.011 I 12.1g_131.577 ll.2'_D 31.5_2 I 12 gJ]

_,032 11.579 I _,_0 12.204 31,296 11,435 31,327 12.047 / 32765 11,317 I 32.777 _ 11 896

672 11.5_/ _0.652 12 12'+ 31.9)8 11.451 31.960 11,_7Z 33,7s59 I|.SM I )3.3_ 11:84'5

31.313 11.60/ I 31.]134 12.032 32,6_0 iI.88_ 32.592 11,898 33.9_1 11.409 ' 33.993 11.796

sz._ 11.6_832.01611.9_ ss.zT) n._ 33.zn _3._4 _._ n,_s _.ss_ 1L_

32.645! I168]l 32.6_oI n._ [3_.9121i],6oI.13_._ I n.t6_135.177n._ 135._70I n.m

(b) Contoured pylon.

Figure 4. Concluded.

12
-.10

C m

-.15

-.20

0

-.25

CD

Noco,,o mrons,tioo __0 .060 0 Off

[] On Free

{_ On Fixed ,,,,,,,_

d,

•050 .o/o,o"

.040 0 / iX -1

°/ /°n< -3

0s0 .0>" o'

0_-" o"

•025 _1 _/,

.020 .C)/C/t)

4

5

2

a,deg

1

0

.015

--.1 0 .1 .2 .3 .4 .5 .6 .7 .8 .9

CL

(a) M = 0.70.

Figure 5. Longitudinal aerodynamic characteristics with nacelles in forward position on symmetrical pylons.

13
C m

CD

-.10

-.15

-.20

-.25

.060

.O55

.O5O

.045

.040

.035

.O5O

.O25

.O2O

.015

-,1

Nacelle Transition

0 off

[] On Free

On Fixed

I

iQJ (

D

.5 .4 '

CL

(b) M = 0.Ts.

Figure 5. Continued.

O

/

_,deg

14
-.10

C m

-.15

-.20

O_

_0

-.25

CD

.060

.055

.O50

.O45

.040

.035

,050

.0251

.020

Nacelle Transition

0 Off

[] On Free

On Fixed

.0

_J

X

0

0

-1

-2

-5

_,deg

.015i

--,1 .2 .3 .4

CL

(c) M = 0.78.

Figure 5. Continued.

.5 ,6 .7 .8 ,9

15
-.10

C m

-.15

-.20

-.25

_O_

_O_

k"-'-"-- _ -- ,"hA A

,.,. ^-%

_O

5

CD

.060

.055

.050

.045

.040

.035

.030

.025

.020

Nacelle

0 off

[] On

On

Transition

0

3

2

o_,deg

1

0

-1

-2

-3

.015

--.1 .1 .2 .3 .4 .5 .6

CI_

(d) M = 0.80.

Figure 5. Continued.

.7 .8 .9

16
-.10

C m

-.15

-.20

-.25

()_-_ O_

-°----_'_--°__- -_--I__._ _

_o

CD

.060

.055

.050

.045

.040

.055

.050

.025

.020

Nacelle Transition

0 Off

[] On Free

On Fixed

_,i.

/"

(

( )_-0- _-----

/c

/0

d

0

-1

-2

-3

a,deg

.015

--.1 0 .1 .2 .3 .4

Cl.

(e) M = 0.82.

Figure 5. Concluded.

.5 .6 .7 .8 .9

17
-.10 r

C m

-.15

-.20

O_

-O--C 10

!

CD

-.25

i

i

i

I

.O2O

0

/

!

.o15 ]

-.1 0 .1 .2 .5 .4 .5 .6 .7 .8 .9

CL

(a) M = 0.70.

Figure 6. Longitudinal aerodynamic characteristics with nacelles in forward position on contoured pylons,

18
-.10

C m

-.15

-.20

_@----__q

_ 0--------

-.25

CD

.O6O

.055

.050

•045

.O40

.035

.030

.025

.020

0

[]

0

Nacelle

Off

On

On

Transition

Free

Fixed

_z

In_ n/ o/

)/Oj

J

/v

0

/

4

3

2

-1

-2

-3

_,deg

0 .1 .2 .3 .4

CL

(b) M = 0.75.

Figure 6. Continued.

.5 .6 .7 .8 .9

19
-.10

C m

-.15

-.20

-.25

- o_ c___c;_ -_:,_._ _r_D_ %__--__

_0

CD

.O6O

.055

.050

.045

.040

.055

.050

.025

.O2O

0

[]

o

Nacelle

Off

On

On

Transition

Free

Fixed

"_, 0

<

o

--I /_-

.0//0

/o

0

-1

-2

-3

ot,deg

.015

--.1 .1 .2 .5 .4

Ct.

(¢)M = 0.78.

Figure 6. Continued.

.5 6 .7 .8 .9

2O
-.10

C m

-.15

-.20

-.25 5

CD

.O6O

.055

.050

.045

.040

.035

.030

.025

.O2O

Nacelle Transition

0 Off ,_ _0

[] On Free

On Fixed /,,_//

¢o

q_ U "

_ _ 0

3

2

e,deg

1

0

-1

-2

-3

.015

--,1 .1 .2 .3 .4 .5 .6 .7 .8

C L

(d) M = 0.S0.

Figure 6. Continued.

.9

21
-.10

C m

-.15

-.20

-.25

()------ O-

-°---__'_ _ _'_ _---_z _o

5

CD

.O60

.055

.050

.045

.040

•O35

.050

.025

.020

Nacelle Transition

0 Off

[] On Free

On Fixed

//

/

/

./

/u

C

C)_ O- '-----_

#l

_0/()/

3

2

a,deg

1

-1

-2

-3

.015

--.1 0 .1 .2 .3 .4 .5

Ct_

(e) M = 0.82.

Figure 6. Concluded.

.6 .7 .8 .9

22
-.10

C m

-.15

-.20

-.25

CD

.060

.O55

.050

J_ /'h

Nacelle Transition <

0 Off {

[] On Free /

_ On Fixed

£

/

.o4-o 0 jC

-1

/

o. _Y // _.

oj <!_o/

.oso < >_/_%15 -3

]/o"

.025 <:_: _<:_ I D

.020 -o/C/0/

2

_,deg

I

.015

-.1 0 .1 .2 .3 .4 .5 .6 .7 .8 .g

CI.

(a) M = 0.70.

Figure 7. Longitudinal aerodynamic characteristics with nacelles in rearward position on symmetrical pylons.

23
-.10

C m

CD

-.15

-.20

-.25

.O6O

.055

.050

.O45

.040

.O35

.030

.025

.020

O_

( _

m

Nacelle Transition

0 off

[] On Free

<_ On Fixed

--0

/0

0

/

3

-1

-2

-3

c¢,deg

.015

--.1 .1 .2 .3 .4

CL

(b) M = 0.75.

Figure 7. Continued.

.5 .6 .7 .8 .9

24,
-.10

C m

-.15

-.20

-.25

_0

5

C D

.060

.055

.050

.045

.040

.035

•030

.025

.020

Nacelle Transition

0 Off

[] On Free

On Fixed

J

.o j

../

0 _C }/u

0

d

/

0

/

2

_,deg

1

-1

-2

-3

.1 .2 .3 .4

CL

{c} M = 0.78.

Figure 7. Continued.

.5 .6 .7 .8 .9

25
-.10

C m

CD

-.15

-.20

-.25

.06O

,055

.050

.045

.040

.035

.030

.025

,020

_0

Nacelle Transition

0 Off

[] On Free

<> On Fixed

/c

%) w

/{j' /U

/b. °

/ oa"_.p° o

D"O 2

/;

Y d

3

2

1

0

-1

-2

-3

_,deg

.015

--.1 .2 .3 .4

C L

(d) M = 0.80.

Figure 7. Continued.

,5 .6 ,7 .8 .9

26
-.10

C

m

-.15

-.20

-.25

C)----- O_

5

C D

.060

.055

.050

.045

.040

.035

,050

.025

.020

Nacelle

0 off

[] On

0 On

--f

C)_ U

Transition

Free

Fixed

/0 ,/_C

__I_0/(

C)__ 0_.__/

/o "°

o//o / .

_Y /

M_s o

/o//

/0 /

/

3

2

1

0

"-1

-2

-3

cqdeg

0 .1 .2 .3 .4

Cl_

(e) M : 0.82.

Figure 7. Concluded.

.5 .6 .7 .8 .9

27
-.10

C m

-.15

-.20

-.25 5

C D

•060

.055

.050

•045

.040

.035

.030

.O25

.020

Nacelle

0 Forward

[] Forward

Rearward

/X Rearward

Transition

Free

Fixed

Free

Fixed ,_

oJ

S

i

i

I

i

I

i

!

i

.015 ',

-.1 0 .1 .2 .,3 .4 .5 .6 .7 .8 .9

4

-1

-2

-3

2

cqdeg

1

C L

(a) M = 0.70.

Figure 8. Effect of longitudinal position of nacelles on aerodynamic characteristics with nacelles installed on

symmetrical pylons.

28
-.10

C m

-.15

-.20

:::_C

-.25

CD

.060

.055

•050

.045

.040

.035

.030

.025

O

[]

O

A

Nacelle

Forward

Forward

Rearward

Rearward

Transition

Free

Fixed

Free

Fixed

AJ /

E

/

2

ot,deg

1

0

-2

-3

.O2O

.2 .5 .4

CL

(b) M = 0.75.

Figure 8. Continued.

.5 .6 .7 .8 .9

2g
C m

CD

-.10

-.15

-.20

-.25

.06O

.055

.050

.045

.040

.055

.050

.025

.O2O

0

[]

<b

z_

i

Nacelle

Forward

Forward

Rearward

Rearward

d

.I

Transition

Free

Fixed

Free

Fixed

v

_ 4x-7'

S

L

E

f

....... ,z.VX'

/

i

i

I 5

; 4

5

1

0

-1

-2

-5

.2 ..3 .4 .5 6 .7 .8 .9

CL

(c) M = 0.78.

Figure 8. Continued.

2

_,deg

3O
-.10

C m

-.15

-.20

-.25

CD

.06O

.055

.050

.045

.040

.035

.030

.025

Nacelle

0 Forward

[] Forward

Rearward

/X Rearward

Transition

Free

Fixed

Free

Fixed

,, VT'I_

0

/

0

-1

-2

-3

a,deg

.020

.1 .2 .3 .4 .5

CL

(d) M = 0.80.

Figure 8. Continued.

.6 .7 .8 .9

31
-.10

C m

-,15

-.20

-.25

[]

CD

.060

.055

.050

.045

.040

.055

.030

0

[]

0

A

Nacelle

Forward

Forward

Rearward

Rearward

Transition

Free

Fixed

Free

Fixed

Sby

<4_ u

//

FI/

/

0

.,t ,g

j j

-1

-2

-3

2

_,deg

.025

.O2O

.015

.2 .5 ,4 .5

CL

(e) M = 0.82.

Figure 8. Concluded.

.6 .7 .8 .9

32
12

-1 0

-.8

-.6

0

};@q

Wing upper surface

.2

.4

.6

o

[3

0

Nacelle

Off

Forward

Rearward

] <

<

/

0 >KIE] "_

! .2 --

.4

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0

Wing lower surface

x/c

(a) Free transition.

Figure 9. Effect of nacelle position on wing chordwise pressure distributions inboard of pylon (rl = 0.328).

Nacelles on symmetrical pylons, M = 0.80, and C L ,_ 0.45.

33
-1.2

1.0 I_ L

'/ O_

-.6 D "

f

)

-.4 i

Cp -.2

0

.2

.4 --

,6 ---

8 __

1.2 --

-1.0 ---

-.8

-.6

Cp -.4

-.2

r _

1/" 5 t_

r_

ol/_j

[]

0

....

I

k Jl(

E

i --

i ....

[

Nacelle

0 Off

[] Forward

Rearward

-1 /

£ -

.1 2 .5 .4

0 j

.O-I

"(}_1

.5 .6 .7 8 9

×/c

(b) Fixed transition.

Figure 9. Concluded.

Wing upper surface

Wing lower surface

1.0

34
-.10

C m

-.15

-.20

-.25

C D

Pylon Transition /

.060 0 Symmetrical Free 5

[] Symmetrical Fixed

Contoured Free /

A Contoured Fixed _ 2

.055 /_,

J

1

•050 _,

.040 -1

.030

.O25

-5

e,deg

.020

.015

-,1 0 ,1 ,2 ,5 .4 ,5 ,6 .7 .8 .9

CL

(a) M = 0.70.

Figure 10. Effect of pylon contouring on longitudinal aerodynamic characteristics with nacelles in forward

position.

35
-.10

C m

-.15

-.20

_=:-_=---4_J _ _ _-_-4_-"'_>

-.25 5

CD

.O6O

.O55

.O5O

.045

.040

•O35

•030

.025

0

[]

0

A

Pylon

Symmetrical

Symmetrical

Contoured

Contoured

Transition

Free

Fixed

Free

Fixed

-1

-2

-3

2

_,deg

.O2O

.1 .2 .3 .4 .5

CL

(b) M = 0.75.

Figure 10. Continued.

.6 .7 .8 .9

36
-.10

C m

-.15

-.20

-.25

CD

.060

.055

.050

.045

.040

.035

.050

.O25

O

[]

A

Pylon

Symmetrical

Symme_.r icol

Contoured

Contoured

Transition

Free

Fixed

Free

Fixed

zl_"k

2

_,deg

1

-1

-2

-5

.020

.015

--,1 .1 .2 .5 .4 .5

CL

(c) M = 0.78.

Figure 10. Continued.

.6 .7 .8 .9

3_
C m

-.10

-.15

-.20

CD

-.25

.O6O

.O55

.050

.045

.040

.035

.030

.025

Pylon

0 Symmetrical

[] Symmetrical

Contoured

/k Contoured

Transition

Free

Fixed

Free

Fixed

/,

o

-1

-2

-3

o_.deg

.020

.015

--,1 .1 .2 .5 .4

CL

(d) M = O.SO.

Figure 10. Continued.

.5 .6 .7 .8 .9

38
-.10

Crn

-.15

-.20

-.25

CD

.06O

.055

.050

.045

.040

.035

.050

.025

0

[]

0

/,,

Pylon

Symmetrical

Symmetrical

Contoured

Contoured

Tr an sition

Free

Fixed

Free

Fixed

./

/f

+ yj

/,

5

2

o_,deg

1

0

-1

-2

-5

.O2O

.015

--.1 .1 .2 .5 .4 .5

CL

(e) M = 0.82.

Figure 10. Concluded.

.6 .7 .8 .9

39
-1.2

-1.0

-.8

-.6

.4

Cp -.2

i

i

i

Wing upper surface

.2

.4--

.6

-1.2

-1,0

.8

0

[]

Pybon

Off

Symmetric@

Con tour ecl

i

I

I

i

i

i

-.6

Cp -.4

-.2

0

[]

.4

0

I

I i

.1 .2 .5 .4 .5 .6 .7 .8 .J 1.0

x/c

Wing lower surface

(a) Free transition.

Figure 11. Effect of pylon contouring on wing chordwise pressure distributions inboard of pylon (r/-- 0.328).

Nacelles in forward position, M -- 0.80, and CL ,_ 0.45.

40
1.2

-1.0

-.8

-.6

[]

Wing upper surface

.2

.4--

.6

Pylon

0 Off

[] Symmetrical

Contoured

-1.2

-1.0

.8

-.6

Cp .4

.2

.4 3 j

0 .1 .2 .3

/r]\

Wing lower surface

1.0

(b) Fixed transition.

Figure 11. Concluded.

41
-- Free transition

.0040 ----- Fixed transition

.0036

• OO32

.0024

.OO2O

.00_

• OO36

• OO32

.0_8

_C D

.00_4

•0020

.0016

• 0056

• 0057

.004_

.00_

.OO4O

AC D

• 0036

. OO32

t

(a) Nacelle-forward, symmetrical-pylon

configuration.

//"

I I I I L._ .... _ f

x

(b) Nacelle-forward, contoured-pylon

configuration.

I t

(c) Nacelle-rearward, symmetricalpylon configuration.

/ i

/

/

/ /

,' / /

/ ,t

/

/

/

!

/

/ ,

/ /

/ ,

/

z//

,68 .70 ,?2 .7d .76 .78 ._ ._ .M

M

Figure 12. Variation of installed drag with Mach number for C L = 0.45.

42
Cp

-.2

_,_- -- ,,_ Transition

O"

/ / 0 Free

-.1 / / [] Fixed

0 ¢ = 0°

0 ¢ = 300

o2\o. _ - fo-..,, _,o[] _ /,

ob / \

0 .C _ ¢ = 180°

0 ) ¢ = 270°

°" _

0 _ ¢ = 330 °

U

.2

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0

×/_

(a) M = 0.70.

Figure 13. Nacelle pressure-coefficient distribution at C L _ 0.45 of nacelles in forward position on symmetrical

pylons.

43
Cp

-.2

n_ -.1

0 _ F-I_

/

F"

0

0

o

[]

.2

0

fC

fc

_jc

./Jq /c

jcJ _

.1 .2 .3 .4 .5 .6

x/Z

I

I

Transition i

I

i 0 Free

I [] Fixed

i

1

_o__..o, _ 3_ B "o

N

\

:g

.7 .8 .9 1.0

(b) M = 0.75.

Figure 13. Continued.

_ = 0 °

¢_ = 30 °

¢_ = 90 °

= 180 °

= 270 °

@ = 330 °
Cp

-.2

fE

O" -.1 I

o i_ -m- _C

o _

-_. _,

o -n" je

/ O_ / c

°ii osJ "_

o_/ ,// _c

0 I_ IU

.1

.2!

0 .1 .2 .3

-o.,,,,,

L

.4 .5 .6

x/Z

(c) M = 0.78.

Figure 13. Continued.

O

[]

3

\o

Transition

Free

Fixed

0

!

;/

:) "0

"0

'B

.7 .8 .9

_ = 0 °

= 30 °

¢_ = 90 °

¢, = 180 °

¢_ = 270 °

¢_ = 330 °

1.0

45
Cp

-.2

-.1

/o'I

0 _/E]_

0

o"

0 /D /

d

oP

.1

.2

0

.U

J/

/

C

2

_E

C

/

,.3

Transition

C) Free

E] Fixed

/

"0

.4 .5 .6

x/Z

(d) M = O.8O.

Figure 13. Continued.

\

7 .8 .9 1.0

¢ = 0 °

¢ = 30 °

¢ = 180 °

¢ = 2700

¢ = 3300

46
Cp

-.2

-.1 O/

/

0 I-]//0"i-I_

o o"

0 /D/

/ ./

} ©

o}o,

o!/.°:/_/

0

.1

.2

0 .1 .2 .3

_C

/C"

0

[]

Transition

Free

Fixed

-©

.4 .5 .6 .7 .8 .9 1.0

X/Z

(e) M = 0.82.

Figure 13. Concluded.

_ = 0 °

¢_ = 30 °

¢_ = 90 °

= 1800

¢_ = 270 °

¢_ = 330 °

47
Op

-.2

-.1

_C

0 _ _ 0_

U

0

/o-/

oA /c

o/d /_

[] o19_

D /

°;

Transition

0 Free

[] Fixed

P

-o

0

"0

"0

-o

"0

i :B .2

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0

= 0 °

= 30 °

¢ = 900

¢ = 180 °

¢ = 2700

¢ = 330 o

x/Z

(a) M = 0.70.

Figure 14. Nacelle pressure-coefficient distribution at C L _ 0.45 of nacelles in forward position on contoured

pylons.

48
Cp

-.2

-,I 0 7

•o //D/

o _©-

oy

t /

o/

'/

o!/-

.1

.2

0

7

/

/

/

.1 .2 .3

_E f-c

_jC

SG"

c

f

C

C

C

-°\

\

\

.4 .5 .6

x/Z

(b) M = 0.75.

Figure 14. Continued.

O

[]

q

3

<\

k

[]

Transition

Free

Fixed

0

I

"'0

¢ = 0°

¢ = 30 °

¢ = 90 °

¢ = 180 o

¢ = 2700

¢ = 3300

4@
Cp

-.2

//( t Transition

.._z _: t 0 Free

-.I 0 / ! [] Fixed

/ /

I_/- I-I; _,- r -...

o_/ ___ _ __,,, ]

[] ./_ \

0 -/ _ _

_ d" l c \ "_:8

0 -

i

[3

.1 0

.2

O .1 .2 .,3 .4 .5 .6 .7 .8 .g

(c) M = 0.78.

Figure 14. Continued.

= 0 °

¢ = 30°

¢ = 900

¢ = 1800

¢ = 270 °

¢ = 330 °

1.0

50
Cp

-.2

-.1

Transition

0 Free

[] Fixed

x/Z

(d) M = 0.80.

Figure 14. Continued.

_ = 0 °

¢ = 50 °

¢ = 90 °

¢ = 1800

¢ = 270 o

¢_ = 550 °

1.0

51
Cp

-.2

-.1

.1

.2

0

Transition

0 Free

[] Fixed

0

/

-{3

×/_

(e) M = 0.82.

Figure 14. Concluded.

¢ = 0 °

¢ = 30 °

¢_ = 90 °

¢ = 180 0

¢_ = 270 °

¢ = 3300

1.0

52
-.7

-.6

-5

C

P

-.4

--,3 --

-.2

Transition

0 Free

[] Fixed

.1

o _- _c

o j

0 _D_-_EF---Zy

o // ,//_

o ._e"/{_

/

0

.,_

I

\

Y

b

_ = 0 °

f_ ¢ = 50 °

X_. ¢ = 90 °

",El

0

:\

R, °_,o

t

¢ = 180 °

¢ = 270 °

¢ = 350 °

0 .1 .2 .3 .4 .5 ,6 .7 .8 .9 1.0

x/Z

(a) M = 0.70.

Figure 15. Nacelle pressure-coefficient distribution at GL ,_ 0.45 of nacelles in rearward position on symmetrical

pylons.

53
Cp

i

Transition

0 Free

[] Fixed

x//£

(b) M = 0.75.

Figure 15. Continued.

¢ = 0 °

¢ = 30 °

¢ = 90 °

¢ = 180 °

¢ = 270 °

¢ = 330 °

1.0

54
-.7

C

P

-.6

-.5

-.4 ....

-.3 --

-.2

Transition

0 Free

[] Fixed

-.1

o _o_

0 _E]'

o_°_

o°!ii_ °_

¢

0 [3"

.2

0 .1 .2 .3

_E t---____Cj

j:

c/°-

_,/

V _ /0-

_C,/c

/

-o\\ %

0

"\ \

.4 .5 .6

x/Z

(c) M = 0.78.

Figure 15. Continued.

,7 .8 ,9 1.0

¢ = 0 °

¢ = 50 °

¢ = 90 °

¢ = 1800

: 270 °

¢ = 5300

51i
C

P

.7

.6

-.5

-.4

-.3 --

-.2

-.1

0

]ronsition

0 Free

[] Fixed

o - e

.1

.2 .5

•

oli'

L

!

.4 5 .6 .7 8 .9

x/t

(d) M = 0.80.

Figure 15. Continued.

¢ = 0 °

¢ = 30 °

¢ = 90 °

¢ = 1800

¢ = 270 °

¢ = 330 °

1.0

56
Cp

-.7

-.6

-.5

-.4

-,3 --

Transition

0 Free

[] Fixed

-.2

-.1 o71_=_ c /

0 Iq" //

J

o / 7- '

[] 0"

O / \

o C

/

0

.l(/

0

.2

0 .1 .2 .3 .4 .5 .6 .7

x/Z

(e) M = 0.82.

Figure 15. Concluded.

C

\

.8 .9 1.0

(_ = 0 °

¢ = 30 °

¢ = 90 °

¢ = 180 °

¢ = 270 °

= 550 °

57
Symmetrical pylon

Nacelle rearward

Contoured pylon

Nacelle forward

Symmetrical pylon

Nacelle forward

i

o

oJ

.rq

Symmetrical pylon

Nacelle rearward

H

Contoured pylon

Nacelle forward

_/_ S_etrical pylon Nacelle forward

o

.2

0

A

II

H

=

0

=

o

"C.

o

,-,,,I

L I I I I

o o o o o

o o o o o

o o o o cD

I

o

58

1. Report No. I 2. Government Accession No. 3. Recipient's Catalog No.

NASA TP-2439 i

4. Title and Subtitle

Nacelle/Pylon/Wing Integration on a Transport Model With a

Natural Laminar Flow Nacelle

7. Author(s)

Milton Lamb, William K. Abeyounis, and James C. Patterson, Jr.

5. Report Date

July 1985

6. Performing Organization Code

505-40-90-01

8. Performing Organization Report No.

L- 15907

10. Work Unit No.

11. Contract or Grant No.

13. Type of Report and Period Covered

Technical Paper

14. Sponsoring Agency Code

9. Performing Organization Name and Address

NASA Langley Research Center

Hampton, VA 23665

12. Sponsoring Agency Name and Address

National Aeronautics and Space Administration

Washington, DC 20546

i15. Supplementary Notes

16. Abstract

Tests were conducted in the Langley 16-Foot Transonic Tunnel at free-stream Mach numbers from 0.70

to 0.82 and angles of attack from -2.5 ° to 4.0 ° to determine if nacelle/pylon/wing integration affects the

achievement of natural laminar flow on a long-duct flow-through nacelle for a high-wing transonic transport

configuration. In order to fully assess the integration effect on a nacelle designed to achieve laminar flow,

the effects of fixed and free nacelle transition as well as nacelle longitudinal position and pylon contouring

were obtained. The results indicate that the ability to achieve laminar flow on the nacelle is not significantly

altered by nacelle/pylon/wing integration. The increment in installed drag between free and fixed transition

for the nacelles on symmetrical pylons is essentially the calculated differences between turbulent and laminar

flow on the nacelles. The installed drag of the contoured pylon is less than that of the symmetrical pylon.

The installed drag for the nacelles in a rearward position is greater than that for the nacelles in a forward

position.

17. Key Words (Suggested by Authors(s))

Nacelle/pylon/wing integration

Natural laminar flow nacelle

Nacelle position

Pylon contouring

Transonic transport

18. Distribution Statement

Unclassified-- Unlimited

Subject Category 02

19. Security Classif.(of this report)

Unclassified ]20. Security Classif.(of this page) ]21. No. of Pages 22. Price Unclassified 59 A04

For sale by the National Technical Information Service, Springfield, Virginia 22161

NASA-Langley, 1985