Wind-Tunnel - NASA Tunnel The Langley Low-Turbulence Pressure Tunnel (ref. 3) is a closed-throat,...
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Wind-Tunnel
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NASA Technical Memorandum 4003
Low-Speed Wind-Tunnel
Results for Symmetrical
NASA LS(1)-0013 Airfoil
James C. Ferris, Robert J. McGhee,
and Richard W. Barnwell
Langley Research Center
Hampton, Virginia
National Aeronauticsand Space Administration
Scientific and TechnicalInformation Office
1987
"l I i
Summary
A wind-tunnel investigation was conducted in the
Langley Low-Turbulence Pressure Tunnel to evaluate
the performance of a symmetrical NASA LS(1)-0013low-speed airfoil. The airfoil contour was obtained
from the thickness distribution of a 13-percent-
thick, high-performance airfoil developed for general
aviation airplanes and is intended for use on the
tail and other aerodynamic control surfaces. The
tests were conducted at Math numbers (M) from
0.10 to 0.37 over a Reynolds number (R) range fromabout 0.6 x 106 to 12.0 x 10 °. The angle of attackvaried fi'om about -8 ° to 20 °.
The results indicate that the aerodynamic char-
acteristics of the present airfoil are similar to those
of the NACA 0012 airfoil. The lift-curve slope at
small angles of attack for the present airfoil is slightly
larger than that for the NACA 0012, and the maxi-
nmm lift coefficient is also larger for Reynolds num-bers greater than about 4.0 x 10_. For smaller
Reynolds numbers, the maximum lift coefficients forthe two airfoils are about the same. The zero-lift drag
coefficient with roughness applied is slightly greater
for the present model for the test Reynolds numbers
(2.2 x 10_, .I.0 x 106, and 6.0 x 106). The opposite
trend was obtained for the airfoils without rough-
ness. The stall characteristics of the present airfoil
are of the tm'bulent or trailing-edge type. The stall
angle of the present airfoil is 1 ° or 2° larger thanthat of the NACA 0012 airfoil. It is shown experi-
mentally that about the same profile drag data areobtained with No. 80 and No. 90 grits for M = 0.15
and R = 4 x 106 and essentially the same zero-
lift drag coefficients are obtained for M = 0.15 andR = 6 × 10_ with No. 60, No. 80, No. 90, and No. 100
grits. The theoretical viscous analysis methods cor-
rectly predicted the lift, drag, and pitching-momentcoefl]cients and chordwise pressure distributions for
points at which the airfoil boundary layer is art ached.
Introduction
Research on advanced technology airfoils has
received considerable attention at the Langley Re-
search Center in the last several years. Refer-ence 1 reports the results for an initial thickness
family of airfoils developed for low-speed general
aviation application. These results show that the
13-percent-thick member of this airfoil family pro-
rides the best overall performance (maximum lift. and
lift-drag ratio). In this report, the basic low-speed
characteristics for a symmetrical version of this airfoil
are presented. This symmetrical airfoil is intendedfor use on the tail and other aerodynamic control
surfaces of general aviation airplanes.
The first series of low-speed airfoils is designated
in the form NASA LS(l)-xxxx, where LS(1) indicates
low speed (lst series), the next two digits give thedesign lift coefficient in tenths, and the last two
digits indicate the maximum thickness in percent
chord. Consequently, the present 13-percent-thick
symmetric airfoil is designated the NASA LS(1)-0013airfoil.
The investigation was performed in the LangleyLow-Turbulence Pressure "lSmnel at Math numbers
from 0.10 to 0.37. The chord Reynolds numbervaried from about 0.6 x 106 to 12.0 x 106 , and the
geometric angle of attack varied from -8 ° to 20 ° .
Symbols
c,,
C
Cc
Cd
Cl
CTrt
C71.
h
M
P
q
lg
t
a7
z
zt
o_
pressure coefficient,qeo
airfoil chord, in.
section chord-force coefficient,
f C,, d(-_)
section profile-drag coefficient,
point-drag coefficient (ref. 2)
section lift coefficient, c,_cosa-cc sin a
section pitching-moment coefficient
about quarter-chord point,
- f C_ (_, - 0.25) d(_) + f C,,_ d(_)
section normal-force coefficient,
- f Cp d(_)
vertical distance in wake profile, in.
free-stream Mach number
static pressure, lb/ft 2
dynamic pressure, lb/ft 2
Reynolds mmlber based on free-streamconditions and airfoil chord
airfoil thickness, in.
airfoil abscissa, in.
airfoil ordinate, in.
mean thickness, in.
geometric angle of attack, deg
Subscripts:
L local point on airfoil
max maximumo zerolift,
,_, undisturbedstream
Model, Apparatus, and Procedure
Model
Tile airfoil model consisted of a metal core with
plastic fill to form the basic contour. (See table I.)
Two thin layers of fiberglass were bonded to the plas-l ic to form the smooth final surface. Tile contour
of the airfoil is compared with that of the NACA
0012 airfoil in figure 1 and a comparison of thenondimensiona] thickness distributions of these air-
foils is shown in figure 2. The model had a chordof 24 inches and a span of 36 inches. The model
was equipped with both upper and lower surface ori-fices located 2 inches off the midspan. The airfoilsurface was sanded in the chordwise direction with
No..100 dry silicon carbide paper to provide a smooth
aerodynamic finish. The model contour accuracy was
generally within -t-0.004 inch.
Wind Tunnel
The Langley Low-Turbulence Pressure Tunnel
(ref. 3) is a closed-throat, single-return tunnel whichcan be operated at. stagnation pressures from 1 to 10
atmospheres with wind-tunnel empty-test-sectionMach numbers up to 0.42 and 0.22, respectively. The
maximum unit Reynolds number per foot. is about15.0 x 106 at a Mach number of about 0.22. The
immel test section is 3 feet wide by 7.5 feet high.
Hydraulically actuated circular plates provided
posit.ioning and attachment for the two-dimensionalmodel. Tile plates are 40 inches in diameter, ro-
t.ate with the airfoil, and are flush with the tunnel
wall. The airfoil ends were attached to rectangular
mode] attachment plates (fig. 3), and the airfoil wasmounted so that the center of rotation of the circu-
lar plates was at 0.25c on tile model reference line.
The air gaps at the tunnel walls between the rectan-
gular plates and tile circular plates were sealed withflexible sliding metal seals shown in figure 3. Tunnel
sidewall boundary-layer control was not available forthis test.
Wake Survey Rake
A fixed wake survey rake (fig. 4) at the modelmidspan was cantilever mounted from the tunnel
sidewall and located 1 chord behind the trailing edge
2
of the airfoil. The wake rake utilized total-pressure
tubes, 0.060 inch in diameter, and static-pressuretubes, 0.125 inch in diameter. The total-pressuretubes were flattened to 0.040 inch for 0.24 inch from
the tip of the tube. The static-pressure tubes each
had four flush orifices drilled 90 ° apart, longitudi-
nally located 8 tube diameters from the tip of the
tube, and radially located in and perpendicular to
the measurement plane of the total-pressure tubes.
Instrumentation
Measurements of the static pressures on the air-
foil surface and the wake rake pressures were made
simultaneously by an automatic pressure-scanning
system with variable-capacitance-type precisiontransducers. Basic tunnel pressures were measured
with precision quartz manometers. Angle of attack
was measured with a calibrated digital shaft encoder
operated by a pinion gear and rack attached to
the circular model attachment plates. Data were
obtained by a high-speed acquisition system and
recorded on magnetic tape.
Test and Methods
The airfoil was tested at. Mach numbers from 0.10
to 0.37 over an angle-of-attack range from about
-8 ° to 20 ° . Reynolds number based on tile airfoilchord was varied from about 0.6 x 106 to 12.0 x 106.
The airfoil was tested both smooth (natural transi-
tion) and with roughness located on both upper and
lower surfaces at. 0.075c. In general, the roughness
was sized for each Reynolds number according to ref-
erence 4; however, a limited roughness study was con-
ducted to verify the appropriate roughness sizes. Theroughness consisted of granular-type strips 0.05 inch
wide, sparsely distributed, and attached to the airfoil
surface with clear lacquer.
The static-pressure measurements at the airfoil
surface were reduced to standard pressure coef-
ficients and machine integrated to obtain sectionnormal-force and chord-force coefficients and sec-
tion pitching-moment coefficients about the quarter-chord. Tile section profile-drag coefficient was com-
puted from the wake-rake total and static pressures
by the method reported in reference 2.
An estimate of the standard low-speed, w'ind-
tunnel boundary corrections (ref. 5) amounted to a
maximum of about 2 percent of the measured coef-
ficients; therefore, these corrections have not been
applied to the data.
q | t7
Presentation of Results
The results of this investigation have beenreduced to coefficient form and are presented in the
following figures:
Figure
Effect of Reynolds nmnber on section
characteristics for NASA LS(1)-0013airfoil at M < 0.15 with model
smooth .............. 5
Effect of Reynolds munt)er on section
characteristics for NASA LS(1)-0013airfoil at ._I < 0.15 with transitionfixed at 0.075c ............ 6
Effect of roughness on sectioncharacteristics for NASA LS(1)-0013airfoil at M = 0.15 .......... 7
Effect of roughness size on zero-lift drag
coefficient for NASA LS(1)-0013airfoil at AI ----0.15 andR = 6.0 x 106 ............ 8
Effect of Math number on section
characteristics for NASA LS(1)-0013airfoil at R = 6.0 x 106 with
transition fixed at 0.075c ....... 9
Effect of angle of attack on pressure
distribution for NASA LS(1)-0013airfoil at M = 0.15 and
R = 2.2 x 10_ with transition
fixed at 0.075c ............ 10
Effect of angle of attack on pressuredistribution for NASA LS(1)-0013
airfoil at M = 0.15 andR = 4.0 x 10_ with transition
fixed at 0.075c ............ 11
Effect of angle of attack on pressure
distribution for NASA LS(1)-0013airfoil at M = 0.15 and
/i' = 6.0 x 106 with transition
fixed at 0.075c ............ 12
Comparison of experimental and
theoretical chordwise pressure
distributions for NASA LS(1)-0013airfoil at M = 0.15 andR = 6.0 x 106 with transition
fixed at 0.075c ............ 13
Comparison of experimental andtheoretical section characteristics
for NASA LS(1)-0013 airfoil at_hi = 0.15 and R = 4.0 x 106
with transition fixed at 0.075c ..... 14
Comparison of experimental andtheoretical section characteristics for
NASA LS(1)-0013 airfoil at M = 0.15and R = 6.0 x 106 with transition
fixed at 0.075c ............ 15
Comparison of section characteristicsfor NASA LS(1)-0013 and NACA 0012airfoils at M = 0.15 with transition
fixed at 0.075c ............ 16
Variation of maxinmm section lift
coefficient with Reynolds number for
NASA LS(1)-0013 and NACA 0012airfoils at M _< 0.15 ......... 17
Variation of maximum section lift
coefficient with Mach number for
NASA LS(1)-0013 and NACA 0012airfoils at R = 6.0 x 106
with transition fixed at 0.075c ..... 18
Variation of zero-lift drag coefficient
with Reynokts number for
NASA LS(1)-0013 and NACA 0012airfoils at M < 0.15 ......... 19
Variation of zero-lift drag coefficient with
Mach number for NASA LS(1)-0013and NACA 0012 airfoils atR = 6.0 x 106 with transition
fixed at 0.075c ............ 20
Discussion of Results
Experimental Results
Airfoil smooth. Figure 5 shows that with the
airfoil smooth (natural boundary-layer transition),
a lift-curve slope of about 0.12 per degree was ob-tained at small angles of attack for all test Reynoldsnumbers. Maxinmm lift coefficients increased fl'om
about 1.0 to about 1.8 as the Reynolds number wasincreased from about 0.6 x 106 to 9.0 x 106 , and
the angle of attack at which the maxinmm lift coef-ficient occurred increased from about 13° to about
19 ° over this Reynolds number range. No apprecia-
ble change is noted in the maximum lift coefficient
or the angle of attack at which maximum lift oc-curs when the Reynolds number is increased above
9.0 x 106. The stall characteristics of the airfoil
are of the turbulent or trailing-edge type as shown
by the lift data of figure 5 and the pressure data of
figures 10, 11, and 12.The pitching-moment coefficient data with the
airfoil smooth (fig. 5) are generally insensitive
to Reynolds numbers from about 2.0 x 106 to12.0 x 106. A decrease in Reynolds number below
about 2.0 x 106 causes a positive increment in Cm
for angles of attack greater than about 9° .
Theprofile-dragdataof figures5 and19showareductionin tile dragcoefficientin thelow •ar g rangeup to a Reynolds number,- of about 4 xl0 b . The
increase in drag at the higher Reynolds numbers is aresult of the decrease in the extent of laminar flow on
the model. The minimum drag coefficient measuredwas about 0.0056.
Roughness effect. The addition of roughness at
0.075(: (figs. 6 and 7) had no appreciable effect onthe lift curves for this airfoil for Reynolds numbers
of 2.2 x 106, ,I.0 x 106 , and 6.0 x 10 a at M = 0.15.
The maximum lift coefficient was unaffected by the
addition of roughness under these conditions. (See
fig. 17.) The addition of roughness causes no ap-preciable change in the pitching-moment coefficientfor M = 0.15 and Reynolds numbers of 2.2 x 106 ,
4.0 × 10a, and 6.0 x 10 r;. The addition of rough-ness forces transition to turbulent flow at the chord
station of 0.075c and thus increases the drag coeffi-
cient. This effect is particularly large in the rangeof lift. coefficient between -0.3 and 0.3 where the ex-
tent of laminar flow on the smooth model can be
appreciable.
An attempt has been made to determine to
what extent tile drag data is affected t)y rough-
hess size. Figure 7(b) shows that about the same
profile-drag data are obtained with No. 80 andNo. 90 grits for M = 0.15 and R = 4.0 x 106 .
It can be seen in figure 8 that essentially the
same zero-lift drag coefficients are obtained forM = 0.15 and R = 6.0 x 106 with No. 60, No. 80,
No. 90, and No. 100 grits. For most general avi-
ation applications, the drag data of most practi-
cal interest are thought to be those obtained with
roughness applied since transition is usually fixed
near the leading edge by" construction roughnessor insect remains gathered in flight. However,
new structural materials and fabrication techniqueswouht enable the use of laminar-flow airfoils on
general aviation airplanes.
Mach number effects. Figure 9 shows that at
a Reynolds number of 6.0 x 106 with transition
fixed at 0.075c, the lift characteristics do not vary
appreciably for Mach numbers below 0.20. The
maxinmm lift coefficient and the angle of attack atwhich it occurs vary smoothly with Much numbersfrom 0.20 to 0.37. This same Much number increase
results in a decrease of about 6° in the angle of attack
for stall and a decrease in Cl,ma x of about 0.50. For a
Reynolds number of 6.0 x 106 with transition fixed
at 0.075c (fig. 9), the pitching-moment coefficient isinsensitive to Much numt)er change from 0.10 to 0.20.An increase in the Mach numbers above 0.20 causes a
positive increment in cm for angles of attack greaterthan about 8 ° .
Figure 9 shows that the profile-drag data forR = 6.0 x 106 with transition fixed at 0.075c are not
affected appreciably by Math numbers from 0.10 to0.37 for lift coefficients between about -1 and 1. For
larger lift coefficients, the general effect of increasing
Much number is to increase tile drag coefficient.
Pressure distributions. The chordwise pressure
data of figures 10, 11, and 12 illustrate the effects of
angle of attack for M = 0.15 and Reynolds numbersof 2.2 x 106, 4.0 x 106, and 6.0 x 106, respectively,
with transition fixed at 0.075c. The symmetry of
the model is indicated by the coincidence of theupper and lower surface data for c_ = 0° and the
coincidence of the data fi'om opposite surfaces of theairfoil for o_ = 4.0 ° and -4.0 °. The airfoil stall is
of the turbulent or trailing-edge type and is reflected
somewhat in the pressure distributions at the highest
angles of attack.
Experimental and theoretical data predictions of
the chordwise pressure distributions by the methods
of references 6 and 7 are compared with the experi-mental data at M = 0.15 and R = 6.0 x 106 for an-
gles of attack of 4 ° and 8° in figure 13. Both methods
predict the pressure distribution well; however, there
are some slight differences over the aft 20-percentchord.
Predictions of the aerodynamic characteristics by
the viscous method of reference 6 are compared at
a Much number of 0.15 and Reynolds number of4.0 x 106 in figure 14 with experimental results for
angles of attack at which the flow is attached. TheReynolds number is increased to 6.0 x 106 in the
data in figure 15, and predictions from reference 7
are also included. The theoretical methods predict
the lift, drag, and pitching-moment data well over
the range of angle of attack to 12° . The method
of reference 7 predicts the drag coefficient somewhat
low in the angle-of-attack range from 8° to 12 ° .
Comparison of NASA LS(I)-0013 WithNACA 0012 Airfoil
The section characteristics of the NASA LS(1)-0013 and the NACA 0012 airfoils at M = 0.15 with
transition fixed at 0.075c are compared in figure 16
for the Reynolds numbers 2.0 x 106, 4.0 × 106, and
6.0 × t06. Both airfoils were tested in the Lang-
ley Low-Turbulence Pressure Tunnel. The lift-curveslope at small angles of attack for the NASA LS(1)-
0013 is slightly larger, and the stall angle of attack
for this airfoil is approximately 1° larger for Reynoldsnumbers of 2.0 × 106 and 4.0 × 106 and approxi-
mately 2 ° larger for the Reynolds number 6.0 × 106.
4
The maximuna lift coefficients for tile two airfoils
(fig. 17) are approximately the same for Reynoldsnmnl)ers of about 4.0 x 106 and less. For larger
Reynolds numbers, the maximum lift coefficient of
the NASA LS(I)-0013 airfoil is larger. For exam-pie, for M = 0.15 and R = 6.0 x 106 with transition
fixed at 0.075c, the values of Ct,max for the NASALS(1)-0013 and the NACA 0012 airfoils are about
1.75 and 1.65, respectively. The difference in C/,ma xdiminishes with increasing Mach number (fig. 18).
Figure 16 also shows that the pitching-moment coef-
ficient for the NASA LS(1)-0013 airfoil is less thanthat for the NACA 0012 airfoil for all positive lift
coefficients. In this range, the pitching-moment co-
efficient for the NASA LS(I)-0013 is zero or slightly
negative, whereas that for the NACA 0012 is zero or
slightly positive.The section drag characteristics for the NASA
LS(1)-0013 and the NACA 0012 airfoils at M = 0.15with transition fixed at 0.075c (fig. 16) are about thesame for Reynolds numbers of 2.0 x 106, 4.0 x 106,
and 6.0 x 106. Figure 19 shows that the drag coeffi-
cient at zero lift for the NASA LS(1)-0013 airfoil at
3I = 0.15 with roughness applied is slightly greaterthan that for the NACA 0012 airfoil under similar
conditions. The additional thickness of the LS(1)-0013 airfoil should cause the drag coefficient to be
slightly higher. The opposite trend applies for theairfoils without roughness. This is a result of more
laminar flow present on the LS(1)-0013 airfoil com-
pared with that on the NACA 0012 airfoil. The zero-
lift drag coefficients of both airfoils with roughness
applied are independent of Mach number over the
test range (fig. 20).
Concluding Remarks
A wind-tunnel test has been conducted in the
Langley Low-Turbulence Pressure Tunnel to evalu-ate the performance of a symmetrical NASA LS(1)-
001.3 airfoil. The contour of this symmetrical air-foil was obtained from the thickness distribution of a
13-percent-thick, high-performance airfoil developedfor general aviation airplanes. The tests were con-
ducted at Mach numbers (M) fl'om 0.10 to 0.37 over
a Reynolds number (R) range from about 0.6 x 106to 12.0 x 106. The angle of attack varied from about
-8 ° to 20 ° .
Tim results indicate that the aerodynamic char-
acteristics of the present airfoil are similar to thoseof the NACA 0012 airfoil. The lift-curve slope
for small angles of attack for the present airfoil is
slightly larger; and the maximum lift coefficient for
the present airfoil is larger for Reynolds numbers
greater than about 4.0 x 106. For smaller Reynolds
numbers, the maximum lift coefficients for the two
airfoils are about the same. The zero-lift drag co-
efficient with roughness applied is slightly greater
for the present, model for the test Reynolds num-bers (2.2 x 106 , 4.0 x 106 , and 6.0 x 106). This
trend is consistent with the fact that the present
airfoil is slightly thicker than the NACA 0012 air-foil. The opposite trend was obtained for the air-
foils without roughness. This is a result, of more
laminar flow present on the LS(1)-0013 airfoil com-pared with that on the NACA 0012 airfoil. Thestall characteristics of both airfoils are of the tur-
bulent or trailing-edge type. The stall angle of the
present airfoil is about 1° greater than that of theNACA 0012 airfoil for R = 2.2 x 106 and 4.0 x 106
and about 2° greater for R=6.0 x 106 . It was
shown experimentally that about the same profile-
drag data are obtained with No. 80 and No. 90 grit forM = 0.15 and R = 4 x 106 and essentially the same
zero-lift drag coefficients are obtained for M = 0.15and /? =6 x 106 with No. 60, No. 80, No. 90,
and No. 100 grits. The theoretical viscous analy-
sis methods correctly predicted the lift, drag, and
pitching-moment coefficients and chordwise pressure
distribution when the airfoil boundary layer isattached.
NASA Langley Research CenterHampton, VA 23665-5225June 24, 1987
References
1. McGhee, Robert J.; and Beasley, William D.: Effectsof Thickness on the Aerodynamic Characteristics of anInitial Low-Speed Family of Airfoils for General Aviation
Applications. NASA TM X-72843, 1976.2. Pankhurst, R. C.; and Holder, D. W.: Wind-Tunnel
Technique. Sir Isaac Pitman & Sons, Ltd. (London),1965.
3. McGhee, Robert J.; Beasley, William D.; and Fos-ter, Jean M.: Recent Modifications and Calibration ofthe Langley Low-Turbulence Pressure Tunnel. NASATP-2328, 1984.
4. Braslow, Albert L.; and Knox, Eugene C.: SimplifiedMethod for Determination of Critical Height of Distrib-uted Roughness Particles for Boundary-Layer Transitionat Mach Numbers From 0 to 5. NACA TN 4363, 1958.
5. Pope, Alan; and Harper, John J.: Low-Speed WindTunnel Testing. John Wiley & Sons, Inc., 1966.
6. Smetana, Frederick O.; Summey, Delbert C.; Smith,
Neill S.; and Carden, Ronald K.: Light Aircraft Lift,Drag, and Moment Predictions--A Review and Analysis.NASA CR-2523, 1975.
7. Bauer, Frances; Garabedian, Paul; Korn, David; andJameson, Antony: Supercritical Wing Sections II. Vol-ume 108 of Lecture Notes in Economics and Mathemat-
ical Systems, M. Beckmann and H. P. Kuenzi, eds.,
Springer-Verlag, 1975.
5
TableI. MeasuredCoordinatesfor NASALS(1)-0013Airfoil
Uppersurface Lowersurface
x/c z/c x/c z/c0.00000
.00624
.01255
.01771
.02470
.03737
.04998
.06260
.07535
.lOOl4
.15007
.19992
.25029
.30055
.34984
.40004
.45013
.50025
.55021
.60038
.65028
.70010
.75018
.80021
.85035
.90019
.95029
.98960
1.00000
0.00000
.01344
.01892
.02222
.02583
.03075
.03465
.03790
.04075
.04541
.05245
.05750
.06094
.06307
.06407
.06432
.06374
.06203
.05896
.05446
.04868
.04183
.03429
.02638
.01859
.01172
.00633
.00371
.00302
0.00000
.00708
.01301
.01843
.02505
.03729
.04993
.06245
.07498
.10003
.14914
.19979
.24982
.29964
.34958
.39969
.44988
.50082
.54984
.59961
.64938
.69962
.74964
.79934.84858
.90080
.94991
.99007
1.O0000
0.00000
-.01435
-.01938
-.02283
-.02612
-.03080
-.03462-.03788
-.04070
-.04543
-.05241
-.05755
.06102
-.06316
-.06428-.06453
-.06388
-.06209
-.05905
-.05459
-,04885
-.04196
-.03441
-.02654-.01891
-.01165
-.00618
-.00358
-.00302
6
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0
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c_
.<o'3
©
7
8
/
//
!
/
! z z!
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l
o.r=mg
\\\
NN
Tunnel side walls
_---- Diam.- 40
2\\\\ \_,\\ _'f///i/////////////////H///,]l
!
L---- N
Airflow A
II
II
Circular plateXtlIIIIIIIII'_ ' r ---
"z;'_ ill-- - -_--_ - - :" _I _ I I Airfoil positioning
attach ment
.Top view /--- Model
/Seal detail "Z" <
m i
-7Tunnel centerline
24
i_ I1/Z__'_%%,s.
._ 36
A
1
attachment
plate
- Zero incidencereference
End view, section A-A
Figure 3. Typical airfoil mounted in wind tunnel. All dimensions are in inches.
9
Statk-pressure probe
l.,_-_ 6.240 1.0_ J 1.032
-_ r ._:_ _-1.296
2
Static-pressure probe J --
Airflow
Tunnel {-
Total-pressure probe
(tubes flattened)
-f.264
.130
29.040
-y
4.704(typ.)
_t
I
Figure 4. Drawing of wake survey rake. All dimensions are in inches.
10
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Figure 10. Effect of angle of attack on pressure distribution for NASA LS(1)-0013 airfoil at AI = 0.15 and/? = 2.2 x 106 wit, h transition fixed at 0.075c.
19
(]p
-lq
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-12
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A Ik i_ _ • Lower surface
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1 ii
Cp
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-3.; Z_
-2 -8 ..................................
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L_
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'A
-1 _ _ "" _L
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o -q.O -. 2
rl 0.0 -.008
0 Ll-I ,El70
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•00_3 -.O00.i
.0096 -.0035
•0116 -.0123.
Upper surface
Lower surface
0 .t .P_ .3 .q .5 .6 .7 .8 .S l .O
z/c
(a) _ _ 8.1 °.
Figure 11. Effect of angle of attack on pressure distribution for NASA LS(1)-0013 airfoil at M = 0.15 andR = 4.0 x 106 with transition fixed at 0.075c.
21
-t;
c, cd cmIA 8.1 .91"8 .0116 -.0127
12.1 1.3Ll .0164 -.0126
b 14.1 1.482 .0201 -.0092
o 16.7 1.636 .0315 -.0054
o 18.7 1.093 -.1225
I z_ t,.r_ (_ o surPaceUppeT
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-10
-S
Gp
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-I
0'
3
I
• t .2 .3 .q .5 .6 .7 .8 .9 t .O
x/c
|
22
(b) (x >__8.1 °.
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1 Iii
Cp
0 .1 .2 .3 .½ .5 .6 .7 .8 .9 t.O
x/c
(a) _ _<8.0°
Figure I2. Effect of angle of attack on pressure distribution for NASA LS(I)-0013 airfoil at ._1 = 0.15 andR = 6.0 × 106 with transition fixed at 0.075c.
23
-1:
-1;
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12.1 1.338 .0138 -.015116.2 1.663 .02ql -.0067
o 18.6 1.76q .025_ -.0022
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-8
tl
I'
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_._ = )_.1 !m_,•2 .3 .q
..=
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x/c
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!
24
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; -q .2
E
=
CP
-q .0 .....
-3.8 --- --
-3.6 -- --
-3 .u,__ . .=
--Ore ExperimentalTheoretical (Ref. 6)Theoretical (Ref. 7)
-3.2
-3.0
-2.8
-2.6
-2.q
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-2.0
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-.6
-.q
-.2
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-Lowersurface _;_5_
• 1 .2 .3 .u, .5 ,6 .? .8 .9 ! .0
xlc
(a) = 4°.
Figure 13. Comparison of experimental and theoretical chordwise pressure distributions For NASA LS(1)-0013airfoil at M = 0.15 and R = 6.0 × l0 G with transition fixed at 0.075c.
25
2B
_t[ .2
-_ .0
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surface
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__k- i_ Z
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x/c
(b) _ = 8°.
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3[
w
=
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IH
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!.8
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1.2
1.0
Airfoil
0 NASA LS(I)-O013[] NACA 0012
Plain symbol, roughness off
Flagged symbol,
.8.4 .7 I 2 4 7 I 0 20 x I0"
R
Figure 17. Variation of maximum section lift coefficient with Reynolds number for NASA LS(1)-0013 andNACA 0012 airfoils at M <_ 0.15.
32
il !iiJ
c_
X
E
0
oo.
_o
c_
-- C'J E
_ _ ×
_ H
o .o _
L_
012
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O NASA LS(1)-0013r-I NACA 0012
Plain symbol, roughness offed symbol, roughness on
010
008
Cd,o 006
004
002
0.4 .7 I 2 4 7 10 20x 106
R
Figure 19. Variation of zero-lift drag coefficient with Reynolds number for NASA LS(1)-0013 and NACA 0012
airfoils at M _< 0.15.
i1lii
9!
lw
zz
©E3 0
I ! I ! l I
0
0
tiI
0
0
g
z
c_
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o X
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0
35
ii II
Report Documentation PageNal_onal Aeronauhcs and
SDace Adminislr ahon
l. NAsARep°rtNO.TM_4003 ] 2. Government Accession No. 3. Recipient's Catalog No.
4. Title and Subtitle
Low-Speed Wind-Tunnel Results for Symmetrical
NASA LS(1)-0013 Airfoil
7. Author(s)
James C. Ferris, Robert J. McGhee, andRichard W. Barnwcll
9. Performing Organization Name and Address
NASA Langley Research Center
Hampton, VA 23665-5225
12. Sponsoring Agency Name and Address
National Aeronautics and Space Administration
Washington, DC 205.16-0001
5. Report Date
August 1987
6. Performing Organization Code
8. Performing Organization Report No.
L-16279
10. Work Unit No.
505-60-21-01
11. Contract or Grant No.
13. Type of Report and Period Covered
Technical Memorandum
14. Sponsoring Agency Code
15. Supplementary Notes
16. Abstract
A wind-tunnel test has been conducted in the Langley Low-Turbulence Pressure Tunnel to evaluate
the performance of a symmetrical NASA LS(1)-0013 airfoil which is a 13-percent-thick, low-speed
airfoil. The airfoil contour was obtained from the thickness distribution of a 13-percent-thick, high-
performance airfoil developed for general aviation airplanes. The tests were conducted at Machnumbers from 0.10 to 0.37 over a Reynolds number range from about 0.6 x 106 to 12.0 x 106 .
The angle of attack varied from about -8 ° to 20 °. The results indicate that the aerodynamic
characteristics of the present airfoil are similar to, but slightly better than, those of the NACA 0012airfoil.
17. Key Words (Suggested by Authors(s))
NASA LS(1)-0013 airfoil
ExperimentalTheoretical
Low speed
Reynolds number variation
Symmetrical
19. Security Classif.(of this report)
Unclassified
18. Distribution Statement
Unclassified Unlimited
Subject Category 02
20. Security Classif.(of this page)Unclassified 21. No. of Pages122. Price36 A03
NASA FORM 1626 OCT 86 NASA-Langley, 1987
For sale by the National Technical Information Service, Springfield, Virginia 22161-2171
i! I ii
