The potential of laminar flow nacelles for reducing installed engine/nacelle drag was studied. The purpose was twofold: to experimentally verify a method for designing laminar flow nacelles and to determine the effect of installation on the extent of laminar flow on the nacelle and on the nacelle pressure distributions. The results of the isolated nacelle tests illustrated that laminar flow could be maintained over the desired length. Installing the nacelles on wing/pylon did not alter the extent of laminar flow occurring on the nacelles. The results illustrated that a significant drag reduction was achieved with this laminar flow design. Further drag reduction could be obtained with proper nacelle location and pylon contouring

Abeyounis, William K.

Lamb, Milton

Patterson, James C., Jr.

Re, Richard J.

English

NASA Technical Reports Server

_ N88- 14949_
NATURAL LAMINAR FLOW NACELLE FOR TRANSPORT AIRCRAFT
/ iJ....o _-S
Milton Lamb
William K. Abeyounis
James C. Patterson, Jr.
Richard J. Re
NASA Langley Research Center
Hampton, Virginia
_m_L'__Q_ aura
P_ECED.JNG PAGE [_L._I_II(r_OT'_L_D
445
https://ntrs.nasa.gov/search.jsp?R=19880005567 2020-03-20T07:45:20+00:00Z
PURPOSE OF TEST
Conceptually, a nacelle designed to have laminar flow over a significant
portion of its length should provide a substantial reduction in the installed
drag of the engine/nacelle/pylon of transport airplanes. While many factors
may make the attainment of laminar flow on nacelles difficult, the magnitude
of the reduction in installed drag and the corresponding increase in airplane
performance makes research in this area of interest. Therefore, the
Propulsion Aerodynamics Branch at the NASA Langley Research Center has
undertaken, in cooperation with the General Electric Company, a research
program to study the potential of laminar flow nacelles for reducing installed
engine/nacelle drag.
The purpose of this research program was twofold: (1) to experimentally
verify a method for designing laminar flow nacelles and (2) to determine the
effects of installation on the extent of laminar flow on the nacelle and on
the nacelle pressure distributions. The procedures used to accomplish these
objectives were:
1). Analytical design a nacelle that had a pressure distribution
favorable for the maintenance of laminar flow.
2). Verify this nacelle design experimentally by measuring the pressure
distributions on an uninstalled flow-through-nacelle model and through flow
visualization techniques determine the location of transition.
3). Following successful completion of steps 1 and 2, analytically
design a nacelle and pylon installed on a high wing transport configuration.
The design criteria was that the installed nacelle have essentially the same
pressure distribution as the uninstalled nacelle.
4). Determine experimentally whether the installation of the nacelle on
the high wing transport produced the predicted pressure distributions and also
determine if the installation causes premature transition.
The analytical designs were developed by the General Electric Company.
The experimental investigations were conducted in the Langley 16-Foot
Transonic Tunnel. The tests were made at a freestream Mach number of 0.80 and
a unit Reynolds number of approximately 4 million per foot. At a freestream
Mach number of 0.8, boundary layer transition occurs in this facility on a 10
deg. polished cone at a Reynolds number of 5.4 million. The angle-of-attack
was varied from -2.5 deg. to 4 deg.
In order to assess the integration effects on the laminar flow nacelle
design, the effects of fixed and free transition, nacelle longitudinal
position, and pylon contouring were investigated using a high wing transport
model. The results indicated that nacelle/pylon/wing integration did not af-
fect the laminar flow on the nacelle. The increment in installed drag between
free and fixed transition for the nacelle on a symmetric pylon was essentially
the difference between the calculated turbulent and laminar skin friction drag
for the nacelle. Locating the nacelle in a forward position relative to the
wing reduced the compressibility effect on the wing lower surface thereby re-
ducing the installed drag. Contouring the pylon resulted in a further reduc-
tion in the installed drag.
446
ORIGINAL PAGE 13 
OF POOR QLPALI'W 
ISOLATED NACELLE I N  THE 16-FOOT TRANSONIC TUNNEL 
The i s o l a t e d  n a c e l l e  was designed by General E l e c t r i c  t o  have a f a v o r a b l e  
The n a c e l l e  shown was pressure d i s t r i b u t i o n  over  70 percent  o f  i t s  leng th .  
t h e  l a r g e s t  p o s s i b l e  model based on t h e  r e s t r a i n t s  of t h e  e x i s t i n g  i n l e t  t e s t  
r i g  o f  t h e  16-Foot Transonic Tunnel. The n a c e l l e  had a maximum diameter  of 21  
inches w i t h  a t h r o a t  d iameter  o f  15.6 inches and was approximately" 22 inches  
i n  length .  
pressure g r a d i e n t  a t  a Mach number o f  0.78 and design c r u i s e  mass f l o w  r a t i o  
o f  0.88. 
The n a c e l l e  was designed t o  have a maximum l e n g t h  o f  f a v o r a b l e  
To h e l p  m a i n t a i n  laminar  f low,  t h e  sur face  was h i g h l y  po l i shed.  
447 
COMPARISON OF MEASURED AND PREDICTED MACH NUMBER DISTRIBUTION
The theoretical Mach number distribution (solid line) was predicted by
the GE Streamtube Curvature Method which solves the Euler equations.
Considerable contour tailoring near the velocity peak was required to produce
an adverse pressure gradient that would initiate transition and yet not
separate the flow. Wind tunnel results show excellent agreement with pre-test
calculations,
M =0.78
0 2 4 6 8 I0 12 14 16 18 20 22 24
X, in.
448
PHOTOGRAPH OF MODEL WITH NACELLES INSTALLED 
The model used i n  t h i s  i n v e s t i g a t i o n  was a 1/24 sca le  t r a n s p o r t  model 
w i t h  a s u p e r c r i t i c a l  wing designed f o r  a Mach number o f  0.8. 
t r a n s p o r t  was considered d e s i r a b l e  s ince  i t  would p r o v i d e  more fuselage 
outwash than a t y p i c a l  low wing a i r c r a f t  and t h e r e f o r e  represent  a more severe 
t e s t  case. The f low-through laminar  f low n a c e l l e  was designed us ing  a l o w  
o r d e r  panel method i n  c o n j u n c t i o n  w i t h  a s p e c i f i c  c o m p r e s s i b i l i t y  c o r r e c t i o n  
technique t o  s o l v e  f o r  t h e  b a s i c  i n s t a l l e d  f l o w  f i e l d .  The d e s i r e d  n a c e l l e  
shape was a n a l y t i c a l l y  modeled i n t o  t h e  wing-body f l o w  f i e l d ,  and t h e  
r e s u l t i n g  n a c e l l e ' s  pressure g r a d i e n t  was checked f o r  separa t ion  as w e l l  as 
f o r  peak sur face  Mach number. I f  e i t h e r  t h e  pressure g r a d i e n t  o r  peak Mach 
number proved u n s a t i s f a c t o r y ,  t h e  n a c e l l e  contour  was m o d i f i e d  and t h e  e n t i r e  
process repeated u n t i l  a s a t i s f a c t o r y  shape was obtained. The f i n a l  n a c e l l e  
shape had a f a v o r a b l e  pressure g r a d i e n t  over 60 percent  o f  i t s  leng th .  The 
surfaces o f  these n a c e l l  es were h i g h l y  po l  i shed. 
The h igh  wing 
450 
ISOLATEU NACELLE FLOW VISUALIZATION RESULTS 
ORlGlMAL PAGEl 
OF POOR QUALITY 
Acenaphthene was u t i l i z e d  f o r  f l o w  v i s u a l i z a t i o n  t o  assess where 
t r a n s i t i o n  from laminar  t o  t u r b u l e n t  f l o w  occurred. Acenaphthene i s  a 
s u b l i m a t i n g  chemical w i t h  a s u b l i m a t i o n  r a t e  t h a t  i s  p r o p o r t i o n a l  t o  heat  
t r a n s f e r  r a t e  ( f a s t e r  i n  t u r b u l e n t  f l o w ) .  It i s  a d i s t i n c t  w h i t e  a f t e r  i t  i s  
sprayed on t h e  model and t h e  f u l l y  laminar  f l o w  reg ion  appears a b r i g h t  w h i t e  
a f t e r  t h e  wind tunne l  run. 
a l l  t h e  way back t o  t h e  d e s i r e d  t r a n s i t i o n  l o c a t i o n ,  w h i l e  o t h e r  p o r t i o n s  
devel  oped " t u r b u l e n t  wedges" p r i  o r  t o  t h e  des ign t r a n s i t i o n  1 o c a t i  on. The 
sources o f  these premature t r a n s i t i o n s  were a t t r i b u t a b l e  t o  two f a c t o r s :  
1. Surface roughness o c c u r r i n g  as a r e s u l t  o f  non-uni form p a r t i c l e  
d e p o s i t s  o f  acenephthene. 
2. Leading edge sur face  contaminat ion due t o  wind tunne l  p a r t i c l e s  
impingement and subsequent n i c k s  and dents. 
Some p o r t i o n s  o f  t h e  n a c e l l e  produced laminar  f l o w  
449 
TEST MATRIX
In order to fully assess the integration effect on a nacelle designed to
achieve laminar flow, the effects of fixed and free transition, nacelle
longitudinal position, and pylon contouring were investigated. The nacelle
was tested in a rearward and forward position on a symmetrical pylon and in a
forward position on a contoured pylon.
x/c w = -0. 90 _ Wing
_ Pylon
Symmetrical pylon __
nacelle rearward _ Nacelle
x/c w = -I. 20
Symmetrical pylon
nacelle forward
m
jf -'--
Contoured pylon-___-_- ____nacelle forward
451
EFFECT OF NACELLE AND NACELLE TRANSITION ON LIFT
The addition of the nacelle/pylon to the wing-body configuration reduced
the lift as would be expected. The addition of grit on the nacelle to trip
the flow from a 60 percent laminar flow condition to a fully turbulent
condition along the entire length of the nacelle, had no effect on the lift
characteristics.
CL
Moo= 0.80
0
Fl
I I I I
0 I 2 3 4
a, deg
Nacelle Transition
Off
On Fixed
On Free
452
EFFECT OF NACELLE AND NACELLE TRANSITION ON DRAG
The addition of the nacelle/pylon to the wing-body configuration resulted
in an increase in drag, The removal of the transition grit from the nacelle
resulted in a decrease in total airplane drag (approximately 9 counts at the
cruise CL of 0.45). The increment in the flat plate friction drag for these
nacelles-and pylons with fixed and free transition on the nacelle indicated
that the drag reduction for two nacelles of this size with 60 percent laminar
flow are approximately 9 counts. This would indicate that integration effects
do not significantly alter the extent of the nacelle laminar flow achieved,
.O55
.050
.045
• 040
CD . 035
.030
•025
•020
•015
-. I
- Moo= 0.80
Nacelle Transition
Off
On Fixed
On Free
I I I I I I I I I I
0 .1 .2 .3 .4 .5 .6 .7 .8 .9
CL
453
I N S T A L L E D  NACELLE FLOW V I S U A L I Z A T I O N  RESULTS 
Acenaphthene was again used f o r  t h e  f l o w  v i s u a l i z a t i o n .  As noted f o r  t h e  
i s 0 1  ated nace l le ,  t h e r e  are  p o r t i o n s  of " t u r b u l e n t  wedges;" however, t h e  e x t e n t  
o f  laminar  f l o w  can be c l e a r l y  seen. 
n a c e l l e s  i s  approx imate ly  60 percent  of t h e  n a c e l l e  length .  
The laminar  f l o w  on t h e  i n s t a l l e d  
ORIGINAL PAGE IS 
03 EOOR QUALITY 
454 
EFFECT OF NACELLE POSITION ON LIFT
An additional study was conducted in an attempt to reduce the installed
drag of the wing-body-nacelle-pylon configuration by proper nacelle
positioning. Moving the nacelle forward reduced the compressibility effect on
the wing lower surface, thereby resulting in a lower installed lift loss.
CL
.8
.7
.6
._ --
.4 --
0-
-3
M =0.80
O0
I I I I I I I
-2 -I 0 1 2 3 4
a, deg
Nacelle
Rearward
A Forward
Transition
Free
Free
455
EFFECT OF NACELLE POSITION ON DRAG
Moving the nacelle to a forward position reduces the compressibility and
nacelle interference effects on the wing lower surface. This results in a
drag reduction of approximately i0 counts at the cruise lift of 0.45.
CD
.060 -
055 -
.050 -
.045 -
• 040 -
.035 -
.030 -
•025 -
. 020 -
•015
-. I
Moo= 0.80
acelle
,, _ _ Rearward
A Forward
I I I I I I I I I I
0 .I .2 .3 .4 .5 .6 .7 .8 .9
CL
Transition
Free
Free
456
EFFECT OF PYLON CONTOUR ON LIFT
A study was also conducted to determine the reduction in installed drag
that might be obtained by contouring the pylon to conform with the local
streamlines. The contoured pylon had very little effect on the lift
coefficient.
CL
.8-
.7 --
.6--
,5 --
.4
o3 --
°2
oi
0-
-° I
-3
M =0.80
oo
Pylon
/x Symmetrical
i',, Contoured
I I I I I I I
-2 -1 0 1 2 3 4
a, deg
Transition
Free
Free
457
EFFECT OF PYLON CONTOUR ON DRAG
Since flow accelerations caused by pylons are a major contributor to high
velocity peaks on the wing lower surface, the pylons were contoured in an
attempt to reduce these high velocities. Contouring the pylon resulted in a
decrease in drag of approximately 7.5 counts.
CD
.060 -
.055 -
• 050 -
• 045 -
• 040 -
.035 -
• 030 -
.025 -
.020 -
.015
-. 1
M =0.80
Pylon
A Symmetrical
Ix, Contoured
I I I I I I I I I I
0 .I .2 .3 .Zl .5 .6 .7 .8 ,9
CL
Transition
Free
Free
458
INSTALLED DRAG FOR M : 0.80 AND CL : 0.45
The incremental drag results obtained from this investigation (wing-body-
nacelle/pylon - wing-body) is presented, including the calculated flat plate
skin friction drag, for each nacelle/pylon configuration. The difference in
installed drag for the nacelles in the rearward position, transition fixed and
free, is comparable to the difference in the calculated skin friction drag.
An additional reduction in installed drag was obtained by locating the nacelle
in the forward position• Pylon contouring further reduced the interference
drag (including wave and form drag) to a value just above skin friction•
AC
.005
•OO4
•003
D
•002
•001
0
Fixed
transition
ACD = CD,WBNP -CD,WB
Drag
Interference
I I Skin friction
Nacelle rearward Nacelle rearward Nacelle forward Nacelle forward
symmetrical symmetrical symmetrical contoured
pylon pylon pylon pylon
459
SUMMARY
The results of the isolated nacelle test illustrated that laminar flow
could be maintained over the desired length.
Installing the nacelles on wing/pylon did not alter the extent of laminar
flow occurring on the nacelle.
The results illustrated that a significant drag reduction was achieved
with this laminar flow design.
Further drag reduction could be obtained with proper nacelle location and
pylon contouring.
460


Natural laminar flow nacelle for transport aircraft

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