An experimental verification of a high performance natural laminar flow (NLF) airfoil for low speed and high Reynolds number applications was completed in the Langley Low Turbulence Pressure Tunnel (LTPT). Theoretical development allowed for the achievement of 0.70 chord laminar flow on both surfaces by the use of accelerated flow as long as tunnel turbulence did not cause upstream movement of transition with increasing chord Reynolds number. With such a rearward pressure recovery, a concave type deceleration was implemented. Two-dimensional theoretical analysis indicated that a minimum profile drag coefficient of 0.0026 was possible with the desired laminar flow at the design condition. With the three-foot chord two-dimensional model constructed for the LTPT experiment, a minimum profile drag coefficient of 0.0027 was measured at c sub l = 0.41 and Re sub c = 10 x 10 to the 6th power. The low drag bucket was shifted over a considerably large c sub l range by the use of the 12.5 percent chord trailing edge flap. A two-dimensional lift to drag ratio (L/D) was 245. Surprisingly high c sub l max values were obtained for an airfoil of this type. A 0.20 chort split flap with 60 deg deflection was also implemented to verify the airfoil's lift capabilities. A maximum lift coefficient of 2.70 was attained at Reynolds numbers of 3 and 6 million

Mcghee, Robert J.

Pfenninger, Werner

Viken, Jeffrey K.

English

NASA Technical Reports Server

N88 : 1494 7 J
//v' -,_ 2-----
/_/_.
ADVANCED NATURAL LAMINAR FLOW AIRFOIL
WITH HIGH LIFT TO DRAG RATIO
Jeffrey K. Viken and Werner Pfenninger
ESCON
Grafton, Virginia
Robert J. McGhee
NASA Langley Research Center
Hampton, Virginia
401
https://ntrs.nasa.gov/search.jsp?R=19880005565 2020-03-20T07:44:46+00:00Z
ABSTRACT
An experimental verification of a high performance natural laminar flow (NLF)
airfoil for low speed and high Reynolds number applications has been completed in the
Langley Low Turbulence Pressure Tunnel (LTPT). The airfoil was designed for a
c = 0.4-0.45 at a Reynolds number of I0 million and M_ < 0.4 with a thickness of
044 chord. Theoretical development allowed for the achievement of 0.70 chord lami-
nar flow on both surfaces by the use of accelerated flow as long as tunnel turbulence
did not cause upstream movement of transition with increasing chord Reynolds number.
With such a rearward pressure recovery, a concave type deceleration was implemented.
This type of recovery efficiently recovers pressure by decelerating most when the
boundary layer has the most energy, then continuously decreasing the gradient toward
the trailing edge. The airfoil's leading edge is moderately sharp representing a
compromise between a sharp nose needed for a wide low drag c range and a blunt
nose for better c . performance. A 0.125 chord simple fl_p is incorporated to
• Imax
substantially increase the low drag cl range by keeping the stagnation point at
the leading edge at different c_'s.
Two-dimensional theoretical analysis indicated that a minimum profile drag coef-
ficient (c d_) of 0.0026 was possible with the desired laminar flow at the design
condition. _ith the three-foot chord two-dimensional model constructed for the LTPT
experiment, a minimum profile drag coefficient of 0.0027 was measured at a c£ = 0.41
and Rec = I0 x I0 . The low drag bucket was shifted over a considerably large
range by the use of the 12.5% chrod trailing edge flap. At a Reynolds number of _
million and 6 = -I0 °, the lower end was shifted to c£0: O. With a positive flapdeflection of {2.5 ° , the upper end was shifted to c£ = .81. This yielded a two-
dimensional lift to drag ratio (L/D) of 245.
Suprisingly high C_max values were obtained for an_airfoil of this type. A
C_max of 1.83 was obtaiffed for 6f = 0° at Re. = I0 x 10 b and M = 0.12. The _(_axdecreases slowly with decreasing Reynolds number. A 0.20 chord split flap with
deflection was also implemented to verify the airfoil's high lift capabilities. A
maximum lift coefficient of 2.70 was attained at Reynolds numbers of 3 and 6 million.
402
NLF(1)-O414F DESIGN OBJECTIVES :_"
The first and primary objective of the design was to design a natural laminar
flow (NLF) airfoil, for low speed applications, that achieved significantly lower
profile drag coefficients at cruise than existing NLF airfoils but was still practi-
cal to use. This resulted in an exercise to design an airfoil with as extensive
favorable gradients (dp/dx < O) as seemed practical without making the far aft pres-
sure recoveries too severe. The airfoil was also designed for reasonably high chord
Reynolds numbers, approximately I0 million (fig. I).
To help lessen the severity of the far aft pressure recoveries with respect to
separation, concave type pressure recoveries were utilized. A concave pressure re-
covery decelerates the flow when the boundary layer has the most energy, tapering the
gradient of the deceleration downstream on the airfoil as the boundary layer loses
energy. For off-design conditions, the possibility of utilizing boundary layer re-
energizers or momentum redistributors was also examined as a means of alleviating the
problem of turbulent separation in the pressure recovery.
To improve C_max performance, a thicker leading edge was utilized than is nor-
mally considered for airfoils with such extensive laminar flow, operating at such
high chord Reynolds numbers. It was known that this thick leading edge would limit
the low drag ca range on the bare airfoil with premature negative pressure peaks
.
however, the chance of a leadlng edge type stall would be reduced. Also,
Pfenninger's earlier work (ref.l) showed that the use of a small chord simple trail-
ing edge flap could be used to regain a respectable low drag c_ range. Deflection
of this small chord flap, both positively and negatively, allows the conversion of
lift due to angel of attack into lift due to flap deflection. By changing the lift
at the design angle of attack, favorable gradients can be maintained on both surfaces
simultaneously for a relatively wide range of lift coefficients.
With the steep pressure recoveries that result on an airfoil of this kind, it
was known that there would be problems with laminar separation at lower chord
Reynolds numbers. In the far aft pressure rises this results in profile drag penal-
ties due to the formation of separation bubbles. In the leading-edge region this
could result in poor high lift performance. The use of boundary layer trips was
explored as a means of causing transition before the laminar separation point was
reached, thereby eliminating the problem.
Finally, when designing configurations for maximum cruise performance, one is
inevitably led to flying as close to (L/D)ma x as possible. This means increasing
the wing loading, and results in the need for greater maximum lift coefficients.
NLF(1)-O414F was designed with the intent of integrating it with a slotted Fowler
flap arrangement and possibly even a Kruger flap to achieve high maximum lift
coefficients.
403
NLF(1)-O414F DESIGN OBJECTIVES
• lOZ CHORDNATURALLAMINAR FLOW (NLF) ON BOTH SURFACESAT REc = 10 RILLION
• CORPRORISESORE LOWDRAGCZ RANGE (AT 6F = 0°) TO IMPROVE CZMAX
PERFORMANCEBY THICKENING THE LEADING EDGE
• INCREASE LOWDRAGCZ RANGEWITH A SRALL CHORDTRAILING-EDGE FLAP
• IMPLEMENTCONCAVEPRESSURERECOVERYTO REDUCETHE TURBULENTSEPARATIONPROBLERWHEN
TRANSITION OCCURSFAR FORWARDON THE AIRFOIL. ALSO, POSSIBLY USE SORE FORE OF
BOUNDARYLAYER RE-ENERG[ZATION OR ROMENTURREDISTRIBUTION
• USE OF BOUNDARYLAYER TRIPS (TAPE, GRIT, BLEED AIR, ETC.) TO ELIRINATE LARINAR
SEPARATIONAT LONER REYNOLDSNUMBERS,BOTH IN THE REAR PRESSURERECOVERYAND AT THE
LEADING EDGE AT HIGH ANGLESOF ATTACK
• IRPLERENTATION OF AN EFFICIENT HIGH LIFT SYSTEM: SLOTTEDFOWLERFLAPS AND POSSIBLY A
KRUGERFLAP
Figure 1
404
NLF(1)-O414F PROFILE WITH FLAP DEFLECTION
Shown in figure 2 is the profile of NLF(1)-O414F. It is characterized by a
moderately sharp leading edge. This leading edge is thicker than a normal high
Reynolds number NLF airfoil but sharper than a conventional turbulent flow airfoil.
Maximum thickness is 14.3% of the chord, being relatively far aft on the airfoil at
the 45% chord location. The trailing-edge region is sharp for low pressure drag
penalties. The 12.5% chord trailing-edge flap is illustrated at two deflection
angles: -10 ° and 12.5 °.
.2--
Y/C
.I
-.1 I I I I I
0 .1 .2 .3 .4 .5
x/c
6f, deg
f -I00
I I I I I
.6 .l .8 .9 1.0
Figure 2
405
EXPERIMENTAL/THEORETICAL PRESSURE DATA AT DESIGN CONDITIONS
A comparison of the_experimental pressure distribution of NLF(1)-O414F at
M® = 0.40, Rec = I0 x I0 b, and _ = -I ° is compared with theoretical pressure dis-
tribution calculated by the Korn-Garabedian (ref. 2) potential flow analysis
(fig. 3). There are favorable gradients on both surfaces up to the 70% chord loca-
tion. The steep concave pressure recoveries of NLF(1)-O414F are also illustrated.
There is a flat spot in the upper surface pressure distribution at x/c = 0.15. This
resulted from the addition of thickness in the leading-edge region to improve c, a-
performance. Results form the Tollmien-Schlichting boundary layer stab111ty ana_'sls
showed that this flat spot in the pressure distribution yielded a smaller disturbance
growth than with a continuous acceleration in this region.
-I.0 -
-.5
0
Cp
.5
1.0
1,5 --
M =0.40, R =lOxlO 6, a=-I °
Theory
Experiment o n
Figure 3
406
STALL CHARACTERISTICS AND REYNOLDS NUMBER EFFECTS
The section characteristics of NLF(1)-O414F are shown in figures 4 and 5 at
chord Reynolds numbers ranging from 2 to I0 million with no flap deflection. For a
chord Reynold number of I0 million (the design case) the minimum profile drag coef-
ficient, at _ = -I °, is 0.0027 with 70% chord laminar flow on both surfaces. This
represents a profile drag coefficient that is only 38% that of an unseparated fully
furbulent airfoil. The low drag bucket is very narrow at this high Reynolds number
in the wind tunnel experiment. The pitching moment coefficient about the quarter
chord point is -.079, at _ = -I °. At the design chord Reynolds number of i0 million
C_max was 1.83, at a = 18.0 ° , with a gentle stall behavior. If the chord Re nolds
ii:ii uiili  wi ;ii i;Ii iz!iiii:i iioii ii ! i  ii  v i:So3 dc T c _ 1 ses
indication of a leading e_ge type stall. As Reynolds number is decreased, the mini-
mum profile drag coefficient increases in a manner greater than unseparated airfoils.
Significant laminar separation bubbles start to occur at the beginning of the pres-
sure rise on each surface, resulting in profile drag penalties. Note the character
of the low drag bucket at chord Reynolds numbers of 2 and 3 million, at both ends,
when the leading edge negative pressure peak eliminates the laminar separation bubble
on one surface or the other, the profile drag coefficient is lower than in the middle
of the low drag bucket. In the middle of the low drag bucket, there are laminar sep-
aration bubbles on both surfaces.
I I I I I I I I I
-8 -4 0 4 8 12 16 20 24 0 .004.008.012.016.020.024-.5-.2 -.I
% deg Cd Cm
Figure 4
407
STALL CHARACTERISTICS AND REYNOLDS NUMBER EFFECTS
C[
2.4
2.0 I
1.6-
1.2-
.8-
.4-
0 I
"'4 l-, 8
-1.2
-8
0
,13
0
i
M R x 10-6
•I00 2.0
•148 3.0
•220 4.4
I
8f =0 °
I I 1 I i I I
-4 0 4 8 12 16 20 24
1 1 I i i i 1 I 1
0 .004. 008.012.016.020•024 -._ -.2 -.I
a, deg Cd C m
Figure 5
408
FLAP DEFLECTION EFFECTS
The drag polar at Re_ = 6.1 x 106 for various flap deflections ranging from -10 °
•
to 20 ° is shown in figure 6. Deflecton of this 12.5% chord simple trailing-edge flap
gave a c- rang with a low drag from c-= -.007 to nearly c = 1.06. The minimum
profile d#ag coefficient with 0° flap de1_lection at Rec = 6 x _06 is 0.0032 for a
c¢ range from 0.273 to 0.417. At the lower end of the drag bucket, with a flap
dCflection of -10 °, the minimum profile drag coeffficient is 0.0036 at c_ = .-007.
At the upper end of the low drag cQ range, with a flap deflection of 20 , the
minimum profile drag coefficient is_.0043 at co= 1.06. This yields a L/D of 247;
however, this is only at one c_ and it would Ee hard to fly at one design point•
Examining the 17.5 ° flap deflection, this same minimum profile drag coefficient,
0.0043, is realized at a section lift coefficient as high as 0.905, yielding a L/D of
210. For this flap deflection, there is a reasonable c. range to fly in. At the
design chord Reynolds number of 10 million, at the upper'end of the low drag range
(12•5 ° flap deflection), there was minimum profile drag coefficient of only 0.0033
with c_= 0.81. This gives a L/D of 245.
NI=.07,R =6.1 xlO6
Cd
•010 -
. OO8
.OO6
.OO4
•OO2
0
-o 4
o h,
o A I'{0
- ,o
,/<<b /J tC_i 1
\ iql
I I I I I I I
-.2 0 .2 .4 .6 .8 1.0
C[
Figure 6
I
1.2
6f, deg
0 -I0.0
[] -5.0
0 0.0
z_ 5.0
I0.0
h 12.5
0 15.0
0 17.5
0 20.0
409
ROUGHNESS EFFECTS
A NLF airfoil has to not only be designed to achieve low profile drag coeffi-
cients with extensive laminar flow but must also be able to perform well when the
flow is fully turbulent. This can happen when insects or rain cause transition far
forward on the airfoil. First, the C_ax performance should not be degraded.
Second, the airfoil should be designed_'o that the profile drag at cruise is not
usually high, resulting from separation in the far aft pressure recovery. Figure 7
shows that section characteristics of NLF(1)-0414_ with transition free and transi
tion fixed near the leading edge at Rec = I0 x I0 . The CQma× with the flow fulIy
turbulent is 1.81 as compared to 1.83 with free transition.--_Tth the flow fully
turbulent, the minimum profile drag coefficient is 0.0080, nearly three times that of
the extensively laminar value. However, this profile drag coefficient is comparable
to that of normal unseparated fully turbulent airfoils.
C_
2.4,-
2.0"
1.6-
1.2
.8
.4
0
-.4
-. 8 1
-1.2
-8
Transition
o Free
a Fixedat O.07.5c
o Fixedat O.010c
6f=0 °,R =lO.OxlO 6
m
m
/
a
n
I
-4 0
I I I I I
4 8 12 16 20
a,deg
I I I I I J I I
0 .004.008.012.016.020.024-.2 -.I
Cd Cm
Figure 7
!
q
410
SPLIT FLAP PERFORMANCE
To show the performance of NLF(1)-O414F under the conditions of a high lift
system, a split flap was tested in the wind tunnel experiment. A high lift system
causes very large negative pressure peaks and will aggravate the leading-edge stall
problem if their is one. This split flap was similar to those tested on the NACA 4
and 5 digit and the 6 series airfoils, a 20% chord plate deflected 60 ° from the lower
surface of the model. The section characteristics of NLF(1)-O414F with and without
the split flap are shown in figure 8. Drag values were not measured because of the
large unsteady wake behind the model. With the split flap installe_, c. ax
NLF(1)-O414F was increased to 2.73 at _ = 9.36 ° and Re c = 6.1 x I0 °. A_"seen in the
plot, the st_ll is very gently, showing no signs of a leading-edge type stall. At
Rec = 3 x I0 U with the split flap, C_nax was 2.66 with the same kind of stall
performance.
C_
..2 m
2.8-
2.4-
2.0-
1.6-
1.2-
.8-
.4-
0
-.4
-8
Of, deg
o 0.0
o 60. 0
M =0.07, R =6.1 xlO 6
_b
oo
I I
Q
0
o
o
o
o
I
o
I
Q
0
I
P
9
8
8
0
q
0
-4 0 4 8
o, deg
I
12 16 20
I I I I I
-.5 -.4 -.3 -.2 -.I
C
m
Figure 8
0
I
.I
411
TAPE TURBULATOR EFFECTS
As the Reynolds number is decreased, the boundary layer becomes increasingly
stable at the laminar separation point in the beginning of the steep pressure rises
on both surfaces. When the highly stable boundary layer reaches the laminar separa-
tion point, it separates and takes a considerable distance before transitioning and
reattaching back to the airfoil surface. Associated with this separated region are
large pressure drag penalties. One method of eliminating this separated region is
utilizing turbulators to trip the flow before the laminar separation point is
reached. Figure 9 illustrates the drag reduction realized by suppressing these lam-
inar separation regions on NLF(1)-O414F for a range of chord Reynolds numbers from 3
to i0 million. The type of tubulator used in this case was tape of 0.012" thick and
1/4" wide placed at 68% chord on the=upper surface and 66% chord on the lower sur-
face. At c o : 0 4 and Re_ = 3 x I0 U, with the turbulator tape installed, the
" ° o o_ • •
profile drag coefflclent is 0.0041. Thls Is 20% less than the profile drag coef-
ficient of 0.0051 measured on the clean airfoil. This benefit is reduced as the
Reynolds number increases and the boundary layer becomes naturally more unstable.
Once the separated region is eliminated naturally,=then there is a drag penalty from
the tape on the airfoil surface. At Rec = I0 × I0 U, with turbulator tape installed,
the profile drag coefficient is 0.0031, instead of the 0.0027 measured on the clean
airfoil.
cd .004
• 008 -
cd
•0O8
cd
• 004
_f- 0° M- 0.15
• 008 -
• 004 _ cd
R : 4. OxlO6
I I
0 .4 .8
cL
R - 3. OxlO6
I I
0 .4 .8 0
cL
Figure 9
•008 -
•004 -
R • 8. OxlO6
I I
0 .4 .8
c[
TAPE
o OFF
a ON Cd
•008
•004
R --6.OxlO6
I I
.4 .8
cL
n
R - lOxlO6
L I I
.4 .8
cL
412
COMPARISON OF AIRFOIL PERFORMANCE
The section characteristics of NLF(1)-O414F are compare_ with Somers' NASA air-
foil section NLF(1)-O215F (ref. 3) figure 10 at Rec = 6 x 10 . The NLF(1)-O215F was
designed for 40% chord laminar flow on the upper sOrface and 60% on the lower. With
the increased extent of laminar flow for NLF(1)-O414F, 70% of the chord on both sur-
faces, the minimum profile drag coefficient is 0.0032, where that of NLF(1)-O215F is
0.0045. Even though NLF(1)-O414F has less overall camber, at a chord Reynolds number
of 6 million NLF(1)-O414F has a C_a x of 1.82 at _ = 18.5 ° and NLF(1)-O215F has a
C_nax of 1.74 at _ = 13.2 °.
C[
2.4
2.0I
1.6-
1.2-
.8--
.4
0
(
-.4
"8 I
-1.2
-8
0
rl
M_< O.15,R = 6.0 x 106
NLF(1)-0215F, 6f=0 °
NLF(1)-0414F, 6f=0 °
m._D O_ 300
,jo °
%
I t I I I I
-4 0 4 $ 12 16 20
I I I I I I i I
0 .004.008.012.016.020.024 -.2 -.I
C d C ma, deg
0 i
o 8o
o @
b ob
o. a
o d
Figure 10
413
CONCLUSIONS
70Z CHORD NLF ACHIEVED ON BOTIt SURFACES AT REc = 10x106 IN LTPT (_/u = O.04Z),
CO = 0.0027
WIDE LOW DRAG CZ RANGE (C Z = 0-0 TO 0.81) ACHIEVED AT HIGH REYNOLDS NUHBERS BY
DEFLECTING A 0-125 CHORD TRAILING-EDGE FLAP, L/D = 245 AT CZ = 0-81
CZmAX PERFORMANCECONSIDERABLy HIGHER THAN EXPECTED, 1-83 WITH _F = O° AND
2-70 WITH 0.20 CHORD SPLIT FLAP (_F = 60°), WITH CORRECT DESIGN OF THE LEADING
EDGE AND STEEP PRESSURE RECOVERY
LOW DRAG PERFORMANCEAT REc = 3x106 IMPROVED ABOUT 20% BY ELIMINATING LARINAR
SEPARATION BUBBLES ON BOTH SURFACES WITH TAPE TURBULATORS
ADDITION OF ROUGHNESSNEAR THE LEADING EDGE REDUCED C_.AX BY ONLY 1Z AT
REc = 10xl0 6 AND 3Z AT RE C = 6x10 6
REFERENCES
i. Pfenninger , Wo: Investigations on Reductions of Friction on Wings, in Particular
by Means of Boundary Layer Suction. NACA TM-II81, August 1947.
2. Bauer, F., Garabedian, P., and Korn, D.: A Theory of Supercritical Wind Sections,
with Computer Programs and Examples. New York: Springer-Verlag, 1972.
3. Somers, Dan M.: Design and Experimental Results for a Flapped Natural-Laminar
Flow Airfoil for General Aviation Applications. NASA TP-1865, June 1981.
414


Advanced natural laminar flow airfoil with high lift to drag ratio

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