A large chord swept supercritical laminar flow control (LFC) airfoil was designed, constructed, and tested in the NASA Langley 8-ft Transonic Pressure Tunnel (TPT). The LFC airfoil experiment was established to provide basic information concerning the design and compatibility of high-performance supercritical airfoils with suction boundary layer control achieved through discrete fine slots or porous surface concepts. It was aimed at validating prediction techniques and establishing a technology base for future transport designs and drag reduction. Good agreement was obtained between measured and theoretically designed shockless pressure distributions. Suction laminarization was maintained over an extensive supercritical zone up to high Reynolds numbers before transition gradually moved forward. Full-chord laminar flow was maintained on the upper and lower surfaces at M sub infinity = 0.82 up to R sub c is less than or equal to 12 x 10 to the 6th power. When accounting for both the suction and wake drag, the total drag could be reducted by at least one-half of that for an equivalent turbulent airfoil. Specific objectives for the LFC experiment are given

Brooks, C. W.

Clukey, P. G.

Harris, C. D.

Harvey, W. D.

Stack, J. P.

English

NASA Technical Reports Server

DESIGN AND EXPERIMENTAL EVALUATION OF A SWEPT SUPERCRITICAL
LAMINAR FLOW CONTROL (LFC) AIRFOIL
W. D. Harvey, C. D. Harris, C. W. Brooks,
P. G. Clukey, and J. P. Stack
NASA Langley Research Center
Hampton, VA
/_P,
475
https://ntrs.nasa.gov/search.jsp?R=19880005569 2020-03-20T07:45:11+00:00Z
ABSTRACT
A large c Ird swel )ercritical laminar flow control (LFC) airfoil has been
designed, constructed, and tested in the NASA Langley/8-ft. Transonic Pressure Tun-
nel (TPT). The LFC airfoil experiment was established to provide basic informa-
tion concerning the design and compatibility of high-performance supercritical
airfoils with suction boundary layer control achieved through discrete fine slots
or porous surface concepts. It was aimed at validating prediction techniques and
establishing a technology base for future transport designs and drag reduction.
Good agreement was obtained between measured and theoretically designed shockless
pressure distributions. Suction laminarization was maintained over an extensive
supercritical zone up to high Reynolds numbers before transition gradually moved
forward. Full-chord laminar flow was maintained on the upper and lower surfaces at
Moo= 0.82 up to Rc _ 12xlO 6. When accounting for both the suction and wake
drag, the total drag could be reduced by at least one-half of that for an
equivalent turbulent airfoil. Specific objectives for the LFC experiment are given
in figure I.
LFC EXPERIMENT OBJECTIVE
Conduct basic aerodynamic and fluid dynamics research program
on a high-performance, swept supercritical, LFC airfoil to determine:
e Ability to laminarize over extensive supercritical region
• Ability of stability theories to predict transition and suction
laminarization requirements
• Relative merit of slottedand perforated suction surfaces for
LFC and HLFC
• Effects of surface conditions and boundary layer influences
on laminarization
Figure 1
476
TEST SETUP FOR LFC EXPERIMENT IN THE 8-FT. TPT
A schematic of the overall LFC experiment in the Langley 8-ft. tunnel is shown
in figure 2 along with tunnel modifications. The major compon_t was a large
chord, 23 ° swept supercritical LFC airfoil of aspect ratio near one which spanned
the full tunnel height. Laminar flow control by boundary layer removal was
achieved by suction through closely spaced fine slots extending spanwise on the
airfoil surface. After passing through the slots, the air passed through metering
holes located in plenums beneath each slot and was collected by spanwise ducts with
nozzles located at the ends. From the duct/nozzles, the air passed through airflow
system evacuation lines, through airflow control boxes which controlled the amount
of suction to each individual duct nozzle, and through sonic nozzles to a I0,000
ft3/min compressor which supplied the suction. All four walls of the tunnel were
contoured in order to produce a transonic wind tunnel flow which simulated
unbounded free air flow about an infinite yawed wing at model design conditions.
The contoured liner was shaped to conform to computed streamlines around the wing
and corrected for growth of the wall boundary layer. The success of the LFC
experiment depended to a large extent on the environmental disturbance levels in
the test section. Isolation of the test section from downstream disturbances was
achieved by an adjustable two wall-choke (sonic throat). Reduction of upstream
disturbances such as pressure and vorticity fluctuations was achieved by the
installation of a honeycomb and five screens in the settling chamber.
m_iner (4 walls)- 1
-" liJII _ /
7',:I',', _ | Suction flow
-_i',iii _ I _IL /- Sonic throat
__!,ill I _ II /
-!!i!i """ -" j_'_'---_. \ ,., \\ [I I Diffuser
_IIIiI
:::,,,, f \ ,,, _
_lllll
_-,,,,, / '/._, c,-, ,.,:..._/Duct-_ - (.,^,
_, ,,,i L 3101
::;: Plenum_ Netering holes
__:_eeey_:mb _ -Nozzle
Figure 2
477
PHOTOGRAPH OF MODEL INSTALLED I N  TUNNEL - UPSTREAM V I E W  
F i g u r e  3 i s  an upstream view o f  t h e  f i n i s h e d  l i n e r  and wing t r a i l i n g  edge as  
seen f rom t h e  t e s t  s e c t i o n  d i f f u s e r  ent rance where t h e  l i n e r  f a i r e d  i n t o  t h e  
o r i g i n a l  t unne l  l i n e s .  The LFC model extended f rom f l o o r  t o  c e i l i n g  and blended 
w i t h  t h e  l i n e r .  The o f f s e t  o f  t h e  wing mean p lane f rom t h e  tunne l  c e n t e r l i n e  may 
be seen as w e l l  as t h e  development o f  t h e  l i n e r  f l o o r  and c e i l i n g  s tep  which 
r e s u l t e d  f rom t h e  d i f f e r e n t i a l  spanwise f l o w  displacement i n  t h e  t u n n e l  channels 
"above" and "below" t h e  wing sur faces.  The dark v e r t i c a l  area on t h e  l e f t  o f  t h e  
photograph and downstream o f  model t r a i l i n g  edge i s  t h e  edge o f  t h e  t e s t  s e c t i o n  
access door. The dark r e c t a n g u l a r  area ahead o f  t h e  model i s  t h e  tunne l  
c o n t r a c t i o n  t h r o a t  region. 
F i g u r e  3 
478 
ORIGINAL PAGE rs 
OF POOR QUALITY 
PHOTOGRAPH OF MODEL INSTALLED I N  TUNNEL - DOHNSTREAM V I E W  
F igure  4 i s  a downstream view o f  t h e  upper sur face  o f  t h e  model taken f rom 
immediately upstream o f  t h e  model. Th is  f i g u r e  i l l u s t r a t e s  t h e  smooth s t reaml ine  
contour  o f  t h e  l i n e r  and how i t  blended w i t h  t h e  model. 
and bottom l i n e r  model j u n c t u r e  reg ions  a r e  s u c t i o n  panels i n  t h e  t i c o l l a r "  around 
t h e  ends o f  t h e  model t o  c o n t r o l  t h e  growth o f  t h e  boundary l a y e r  i n  these 
regions. The dark area on t h e  l e f t  v e r t i c a l  w a l l  and downstream o f  t h e  model i s  
one o f  t h e  f l e x i b l e  two-wal l  chokes (son ic  t h r o a t ) .  The choke p l a t e  on t h e  
oppos i te  w a l l  i s  h idden behind t h e  model. The dark area immediately i n  back o f  t h e  
model i s  the  tunne l  t e s t  s e c t i o n  access door f o l l o w e d  by t h e  downstream h i g h  speed 
d i  f f user. 
The dark areas a t  t h e  t o p  
F i g u r e  4 
479 
MEASURED AND DESIGN PRESSURE DISTRIBUTIONS
Measured and design chordwise pressure distributions on the upper and lower
surfaces of the LFC model are shown in figure 5 for two chord Reynolds numbers at
the design Mach number of 0.82. In general, these representative results indicate
measured pressure distributions very close to design. Shockfree flow is shown for
10-million Reynolds number and essentially shockfree flow for 20-million. The
slightly overall higher velocities on the upper surface and the chordwise deviation
from the design pressure distribution were attributed to classical problems
associated with wind tunnel testing, wall interference and model deformation under
design air loads. The velocity field between the upper surface and tunnel wall
(supersonic bubble zone) was slightly higher than predicted due to the liner
contour and inability to completely account for boundary layer displacement effects
in the design analysis. Coordinate deviations from design over the LFC model
forward upper surface at midspan were measured under simulated air load to be about
O.O03-inches and produced local surface contour deviations and irregularities in
the pressure distribution. As Reynolds number increased above lO-million,
transition moved rapidly forward on the lower surface and the flow became unable
to sustain the adverse pressure gradient leading into the trailing-edge cusp and
separation occurred at about 80-percent chord. This separated flow changed the
local effective area distribution of the test section resulting in a slightly
higher freestream Mach number and increased upper surface shock strength at
20-million Reynolds number.
-1.2
-0.8
-0.4
Cp 0.0
0.4
0.8
1.2
0.0 0.2
M = 0.820 M = 0.823
Rc= 10x 106 C{=.64 Rc= 20x 106 C[=.56
1%c= .82 106o //--Design RC 20 x o
0
 OOOo..... .55 i...... %,iooooo _oo_o____o_.%op_o_.
-6 a LOO6_.L. _V. A
...................... v--- •I....... _ o,
.... 1
+
%
%
\
%
L.... _
#
%
0.4 0.6 0.8 1.0 0.0 0.2 0.4 0.6 0.8
x/c x/c
1.0
Figure 5
480
MEASURED AND THEORETICAL SUCTION DISTRIBUTIONS
ORIGINAL PAGE IS
OF POOR QUALITY
The measured chordwise suction coefficient (CQ) distribution required to
maintain full chord laminar flow over both surfaces at the design Mach number of
0.82 and lO-million chord Reynolds number is shown in figure 6 compared to the
theoretical suction distribution. The required suction level was higher than the
theory over most of the upper and lower surfaces. About two-thirds of the
predicted or measured total suction contribution for both surfaces is necessary for
control of the lower surface geometry alone. The higher suction requirements were
due to the overvelocities and the surface pressure irregularities, as well as
higher suction control required to overcome the cross flow instabilities associated
with the steep pressure gradients on the upper and lower surfaces and the
minimization of centrifugal Taylor-G_rtler type boundary layer instabilities and
interactions in the concave regions of the lower surface. The overall higher
suction levels are also influenced by tunnel disturbance levels which are
inherently higher than free-air turbulence levels expected in flight.
CQ
0
-10
-20,
-30
M =0.82 RC=10x106
x 10-4
,, Upper surface 1L , , I l
0 _ F'--'T-"_ _ l:heoryl
! _,_." --,
-,o ,.
_/ Experiment /---!/ ' il
i,
-60 :_
!
Lower surface
.2 .4 .#, .8 1.o
XlC
Figure 6
481
SUMMARYOF TRANSITION VARIATION WITH REYNOLDS NUMBER
The data presented in figure 7 show the chordwise extent of laminar flow
achieved on the upper surface for several Mach numbers up to the design Mach number
of 0.82, as determined by a grid of flush mounted surface thin film gages. At
Rc = lO-million, full chord laminar flow could be maintained over the upper and
lower surfaces for all Mach numbers. As Reynolds number was increased for constant
Mach number, transition moved gradually forward on the upper surface. The Reynolds
number at which this forward movement began was dependent on Mach number and
occurred at progressively lower Reynolds numbers as Mach number increased. For the
design Mach number of 0.82, the forward movement began between II- and 12-million
and reached about 65-percent chord at Rc = 20-million. Transition on the lower
surface moved more rapidly than on the upper surface and occurred near the leading
edge for M : 0.82 and Rc = 20,million. It was concluded that suction
laminarization over a large supercritical zone is feasible to high chord Reynolds
numbers even under non-ideal surface conditions on a swept LFC airfoil at
high lift.
AA
Upper surface
0.4
0.2
O, L i _ t _ i J 106I0 12 14 16 18 20 22 24 x
RC
Figure 7
482
SUMMARY OF DRAG WITH M®
The total drag at M_ = 0.40 and 0.82 and Rc = 10 million with full chord
laminar flow is seen in figure 8 to be equal to about 31 counts (c d = 0.0031).
This represents an approximate 60-percent drag reduction as compared to an
equivalent conventional turbulent airfoil drag level of about 80 counts. Total
drag is the sum of measured wake drag from a wake rake at midspan and the suction
drag penalty required to maintain full chord laminar flow. The suction required to
maintain full chord laminar flow was somewhat higher than anticipated and the
contribution to the total suction drag was approximately 40-percent from the upper
surface and 60-percent from the lower surface. The increase in wake drag for Mach
numbers just below the design Mach number of 0.82 was associated with the formation
of a weak shock wave near the leading edge as the supersonic bubble began to
develop. As the bubble developed (0.78 < Moo < 0.80) full chord laminar flow was
still present but periodic turbulent bursts occurring over the upper surface caused
an increase in the wake drag. As the Mach number increased to 0.82, the supersonic
zone spread rearward to approximately the 80-percent chord, the turbulent bursts
over the upper surface disappeared, and the wake drag returned to near its subsonic
level.
Rc = I0 x 106
0.006 -
CD
0.004
0.002
Total drag -_..__
c o
D q:b[]
Suction drag
I I I I I
0 0.2 0.4 0.6 0.8 1.0
M
oO
Figure 8
483
MEASURED DRAG ON AIRFOILS WITH/WITHOUT SUCTION CONTROL
A summary of the measured drag on airfoils with and without suction control,
developed by the Airfoil Aerodynamics Branch of the Transonic Aerodynamics Division
over the past several years, is shown in figure 9. The most recent design concepts
with Natural Laminar Flow (NLF) or Laminar Flow Control (LFC) are identified as
NLF(1)-O414F, HSNLF(1)-0213, and SCLFC(1)-O513F. Performance evaluation of all the
concepts shown was conducted in NASA Langley facilities that have been rehabilitated
or modified for improved flow quality and low drag testing except the 6- x 28-inch
Transonic Tunnel (TT) which has not been modified. The total drag of the swept
supercritical LFC airfoil with suction slots includes the suction drag penalty
required to maintain full chord laminar flow. The solid symbols represent drag
levels obtained with the maximum extent of laminar flow at the design lift
coefficient and Reynolds number. The open symbols indicate drag levels obtained
with the same airfoils with fully turbulent, attached flow tripped at the leading
edge. In general, the results indicate about 60% drag reduction achieved with
laminar flow over the speed range, with or without suction control or sweep, when
compared with a turbulent drag level of about 80 counts. Of further major
importance is the fact that both the NLF(1)-O414F and HSNLF(1)-0213 airfoils showed
no degradation of lift performance or pitching moment characteristics when fully
turbulent.
CD T =
(CDs + CDwl
Data Facility Airfoil __A" (!JD)ma x
O LTPT
[] LTPT
_7 LTPT
LTPT
6 x 28 l-i"
'_ 8-It TPT
0 8-ft TPT
.0100 -
() /x
.0080[}- _7 []
.0060 --
.0040 V&
•0020 1 I
i .2
Long endur 0r 218
NLF(1)-0414F 0° 160
NLF(1)-0215F Or 175
HSNLF(1)-0213 0° 53
SCLFC(1)-0513F 23* 183
SC Turb 0° 75
O
Z_
Solid symbols for laminar flow
• • A A
n
f "_ I I I
.3 .4 .5 .6 .l
M.L= Moo COS.,_.
I
.8
Figure 9
484


Design and experimental evaluation of a swept supercritical Laminar Flow Control (LFC) airfoil

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