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_*/,NASA...._Technical
:_iPaper_:3133
' September 1991
• I
"- }
Full-Scale SemispanTests of a Business-Jet
Wing With a NaturalLaminar Flow Airfoil
David E. Hahne and
Frank L. Jordan, Jr.
NASATechnical
Paper3133
1991
National Aeronautics andSpace Administration
Office of Management
Scientific and TechnicalInformation Program
Full-Scale SemispanTests of a Business-JetWing With a NaturalLaminar Flow Airfoil
David E. Hahne and
Frank L. Jordan, Jr.
Langley Research Center
Hampton, Virginia
The use of trademarks or names of manufacturers in this
report is for accurate reporting and does not constitute an
official endorsement, either expressed or implied, of such
products or manufacturers by the National Aeronautics and
Space Administration.
Abstract
An investigation was conducted in the Langley
30- by 60-Foot Tunnel on a full-scale semispan model
to evaluate and document the low-speed, high-lift
characteristics of a business-jet class wing that uti-
lized the HSNLF(1)-0213 airfoil section and a single-
slotted flap system. In addition to the high-liftstudies, boundary-layer transition effects were exam-
ined, a segmented leading-edge droop for improvedstall/spin resistance was studied, and two roll-controldevices were evaluated.
The wind-tunnel investigation showed that de-
ployment of a single-slotted, trailing-edge flap waseffective in providing substantial increments in lift
required for takeoff and landing performance. Fixed-
transition studies to investigate premature trippingof the boundary layer indicated no adverse effects
on lift and pitching-moment characteristics h)r either
the cruise or landing configuration. The full-scale re-
sults also suggested the need to further optimize the
leading-edge droop design that was developed in thesubscale tests.
Introduction
While much research on natural laminar flow
(NLF) airfoils has recently focused on drag reduc-
tion for improved cruise performance, few studies
have addressed the use of high-lift systems for takeoff
and landing with this wing class. Although large im-
provements in cruise performance have been shown,
these NLF airfoils will only be used if they carl be
equipped with a viable flap system that is capable
of generating enough lift to meet takeoff and landingrequirements.
Prior to this investigation, some two-dimensionalwind-tunnel tests had been conducted to evaluate
high-lift characteristics of NLF airfoils and to sup-
port associated theoretical studies of flap effective-
ness (refs. 1 and 2). These tests were focused onthe use of simple split flaps. Other studies were
conducted that used theoretical methods to design
more complex flap systems for NLF airfoils (ref. 3).One of the airfoil sections used for the study in
reference 3 was the high-speed HSNLF(1)-0213 air-
foil. This airfoil was developed to extend the nat-
ural laminar flow concepts that were developed for
low-speed airfoils to airfoils intended for higher speed
and Reynolds number applications (refs. 4 to 6). As
stated in references 6 and 7, the HSNLF(1)-0213 air-foil was designed for a cruise section lift coefficient
of 0.26 at a Mach number of 0.7 and a Reynoldsnumber of 9 x 106. Theoretical data on the airfoil
predicted that large increments in lift could be ob-
tained with a slotted flap design (ref. 3). As ex-pected, the amount of additional lift and the angle of
attack for maximum lift depended on the flap gcom-
etry. Two-dimensional theoretical studies indicated
that a single-slotted flap design would offer a good
trade-off between CL,max and flap complexity for useon lightweight business jets (ref. 3). However, be-
cause theoretical techniques cannot reliably predict
maximum lift for three-dimensional wings with flaps,
experimental tests are necessary to accurately evalu-
ate any flap system.
In the present investigation, tests were conducted
in the Langley 30- by 60-Foot Tunnel on a fill-scale
semispan model that incorporated the HSNLF(1)-.0213 airfoil section. The main objective of these tests
was to evaluate and document the low-speed, high-lift characteristics of a business-jet class wing that
used the HSNLF(1)-0213 airfoil section and a single-
slotted flap system that was designed with the aid
of the computer code described in refercnce 8. Thisflap system was the same as the one discussed in ref-
erence 4. Photographs of the model mounted for tests
are shown in figure 1. Figure l(a) shows the model
with the flap retracted and the flow going from right
to left. Figure l(b) shows a close-up of the underside
of the model with the flap deflected 40 ° . In addition
to the high-lift studies, boundary-layer transition ef-
fects were examihed, a segmented lcading-cdge droopfor improved stall/spin resistance wins studied, andtwo roll-control devices were evaluated.
Symbols
Longitudinal forces and moments are presentedin the stability-axis system, and lateral forces and
moments are presented in the body-axis system. Amoment reference center of 0.25_ was used for all
tests.
b wing span, ft
CD drag coefficient, _L_q_ :_
CL lift coefficient, Lift
CL,ma x maximum lift coefficient
C l rolling-moment coefficient,Rolling moment
q_Sb/2
pitching-moment coefficient,Pitching moment
q_,S_
pressure coefficient,' q,'x_
incremental rolling-momentcoefficient
local wing chord with droop off, ft
Cm
P
poc
q_
R
S
x
Y
z
o_
6a
t_f
Abbreviations:
FS
MCARF
NLF
VG
WS
mean aerodynamic chord, ft
local static pressure, lb/ft 2
free-stream static pressure, lb/ft 2
free-stream dynamic pressure, lb/ft 2
Reynolds number based on
semispan reference area, ft 2
chordwise distance from wing
leading edge, positive aft, ft
spanwise distance from wing root, ft
normal distance from wing leading
edge, positive up, ft
angle of attack, deg
aileron deflection, positive trailing
edge down, deg
flap deflection, positive trailing edge
down, deg
spoiler deflection, positive trailing
edge up, deg
fuselage station, in.
Multi-Component Airfoil Analysis
Program
natural laminar flow
vortex generator
wing station, measured plane of
wing, in.
Model Description and Apparatus
The geometry of the semispan model tested is
shown in figure 2, and a summary of the geo-metric characteristics is contained in table I. The
HSNLF(1)-0213 airfoil section used in these tests isshown in figure 3, and section coordinates for this air-
foil are given in table II. The wing incorporated 3°
of twist between wing station 0.0 and the 50-percent
semispan station. An additional 1 ° of twist was in-
corporated between the 50-percent semispan station
and the wingtip for a total of 4° washout. The in-
board portion of the wing was twisted about the 30-
percent chord line, and the outboard portion of the
wing was twisted about the 78-percent chord line. Asmall winglet was located at the wingtip. The model
also incorporated an aileron and spoiler for roll con-
trol. (See fig. 2.) A half body of revolution was
incorporated to simulate the presence of a fuselage
near the wing. This fuselage was representative of a
business jet in both size and shape. A vortex gener-
ator was mounted on the fuselage (fig. 4) just abovethe wing-body juncture to delay flow separation on
the inboard panel of the wing. A multiposition flap
system was incorporated in the model for evaluation.
(See figs. l(b) and 5.) The flap was a 28-percentchord flap that extended from the wing root to a
semispan location of 2y/b = 0.79. Flap deflections
(0 °, 20 °, and 40°), flap gap, and flap overlap were
set by changing three brackets that were located on
the lower surface of the wing. Flap overlap was de-
fined as the distance from the trailing edge of the
wing upper surface (0.92c) to the leading edge of the
flap (negative when the wing overlaps the flap). Flap
gap was defined as the shortest vertical distance be-
tween the wing upper surface (0.92c) and the flapleading edge. The nominal flap overlap and gap were
0 and 2 percent of the local wing chord for the flap
deflected 40 ° and -3 and 4 percent of the local wing
chord for the flap deflected 20 ° .
In an attempt to improve stall/spin resistance, a
segmented leading-edge droop was developed for thewing prior to the full-scale senlispan tests. These ex-
ploratory tests used two subseale models. The mod-
els incorporated the same airfoil section and wing
ptanform as the full-scale semispan model. However,
the subscale model wings that were used for droop
development were not twisted, and they did not in-
corporate a winglet. Static force tests, which wereconducted in the Langley 12-Foot Low-Speed _ln-
nel at a Reynolds number of 3.1 x 105, were used
to develop several candidate droop geometries. The
roll-damping characteristics of these droop designs
were then evaluated in dynamic force tests in the
Langley 30- by 60-Foot Tunnel. This evaluation was
used to select the final droop configuration for the
full-scale tests. These roll-damping tests were con-ducted at a Reynolds number of 9.7 x 105. Refer-
ence 9 contains a description of the techniques used
to develop this droop design. Figure 6 shows the
leading-edge droop location and droop section that
resulted from the subscale tests. The droop design
consisted of two segments: an outboard segment that
extended from the tip inboard to approximately the
70-percent semispan location and a smaller segmentthat was mounted farther inboard between the 40-
and 50-percent semispan locations. The droop sec-
tion was derived from another NLF airfoil (NLF(1)-0215F). This approach was adopted in an attempt to
achieve natural laminar flow on the drooped as well
as the undrooped portions of the wing. Coordinates
for the drooped airfoil section are given in table III.
Unlessotherwisenoted,thedatapresentedhereinarefor configurationsthat donot includethedroop.
Static force and momentmeasurementsweremadein the Langley30-by 60-FootTunnelwiththeexternalscalesystemthat is describedin refer-ence10. Theloadsonboth thewingandthe fuse-lageareincludedin all forceandmomentdata. Inadditionto forceandmomentmeasurements,static-pressuredata, flowvisualization,andhot-fihndatawereobtained.A total of 322pressuretapswerespacedalong the spanof tile wing at eight sta-tions:20-,35-,45-,55-,62-,75-,85-,and95-percentsemispanlocations.Chordwiselocationsof boththeupperandlowersurfaceportsarelistedin tableIV.A total of 45hot-filmsensorsweremountedon thewing (both upperand lowersurfaces)to measureboundary-layerbehaviorat threespanwiselocations.(Seefig. 7.) Surfacetufts wereusedto visualizethesurfaceflowconditions,especiallystall progression.Thesetuftswereremovedwhilepressureandhot-filmdatawereobtained.
Test Conditions and Corrections
Testswere conductedover an angle-of-attackrangefrom-10° to 40°. Aerodynamicforceandmo-mentdatawereobtainedat free-streamvelocitiesof55,66,and77mph,whichcorrespondto Reynoldsnumt)ersbasedon_of about3.05x 106,3.67x 106,and4.26x 106,respectively.Mostof the testswereconductedat a free-streamvelocityof 66mph;un-lessotherwisenoted,thedatapresentedarefor thiscondition. Althoughsomeroll dataweretakentoevahmteaileronand spoilereffectiveness,the testsfocusedon longitudiimlcharacteristicsof the semi-spantoo(tel.
A wind-tunnelcalibrationwas madeprior tomodelinstallationto determinebuoyancyandflowangularitycorrections.Flow-fieldsurveysweremadeto determineflowblockagecorrectionsin themannerdescribedin reference11.Correctionsforjet bound-ary interferenceweremadein accordancewith themethodof reference12and aredescribedin refer-ences13to 15. TheLangley30-by 60-FootTunnel
has a measured turbulence factor of 1.1, which cor-
responds to an average turbulence level of approxi-
mately 0.1 percent of the mean flow velocity (ref. 16).
Results and Discussion
Pressure Distributions
Chordwise pressure distributions for (5f = 0° and
5f = 40 ° are presented in figures 8 and 9. Data wereobtained for angles of attack between -2.2 ° and 16.8 °
at the eight semispan stations that were previously
discussed. Problems with the pressure measurement
system resulted in a limited amount of reliable pres-
sure data. The pressure data presented in figure 8(b)
for cz = 1.5 ° with the flaps retracted (CL _ 0.3)
show that pressure gradients are conducive to lanfi-
nar flow over much of the upper and lower surfaces
of the wing. This conduciveness is indicated by the
decreasing values of the pressure coefficient with the
chord station up to about the 60-percent chord sta-
tion. These results indicate that, even at low speeds,
the amount of laminar flow possible over the wing at
cruise angles of attack should be significant.
A comparison of the pressure data in figures 8
and 9 (_i] = 0° and fI = 40°) indicates that largeincreases in lift result from flap deflection. Even
though the vortex generator wa_s on for the bf = 40 °data, the generator was believed to have no effect on
the pressure data because flow visualization studies
indicated that the vortex generator primarily affected
the flow inboard of the 2y/b = 0.2 pressure port
station. Integration of the pressure distribution datato calculate the lift on the wing and flap indicates
that this increase in total lift results not only from the
lift generated on the flap but also from the enhanced
lift characteristics on the main wing.
Effect of Reynolds Number
The effect of Reynolds number on the longitudi-nal characteristics for the model without the vortex
generator and with the flaps retracted is shown in
figure 10. Changes in Reynolds number had no ef-
fect on lift characteristics except in the maxinmm
lift and post-stall angle-of-attack regions. The in-crease in both the maximmn lift coefficient and stall
angle of attack at the higher Reynolds numbers re-
sulted from the increased resistance of the boundarylayer to separation. With the flaps deflected 40 ° and
the vortex generator installed, Reynolds number ef-
fects on lift and pitching moment were limited to a
small angle-of-attack range just past the stall. (See
fig. 11.) The maximum lift coefficient and stall an-
gle of attack, however, were not affected by Reynoldsnumber variations.
Two-dimensional data from reference 7 for the
basic airfoil with _f = 0 ° showed very little changein minimum drag for Reynolds numbers between4.0 x 106 and 9.0 x 106 at low subsonic Mach numbers.
However, these two-dimensional tests did indicate
that there were some Reynolds number effects onmaximum lift between R = 4.0 x 106 and R =
6.0 x 106. Since the full-scale data indicated virtually
no Reynolds number effects between R = 3.67 x 106
and R = 4.26 x 106 (the maximum Reynolds number
obtainable for these tests), the majority of these tests
3
wereconductedat R = 3.67 × 106 (66 mph). Unless
noted otherwise, data presented herein are for thistest condition.
Effect of Transition
With any NLF airfoil, there is always concern
about the effect on the aerodynamic characteristics
of premature boundary-layer transition that results
from leading-edge contamination (e.g., dirt, insects,
or scratches). In addition to performance impacts,
the potential effect on aircraft stability characteris-tics must also be addressed. To evaluate the effect of
early transition on the wing longitudinal character-
istics, boundary-layer transition was fixed at a chord
location of x/c = 0.05 on both the upper and lower
surfaces of the wing by using the guidelines in ref-
erence 16. Tape with a serrated leading edge, 1/2in.
wide and 1/64 in. high, was used on the upper sur-
face, and a double thickness of the same tape was
used on the lower surface. Transition strips were not
used on either the fuselage or wing flap. The effect offixed transition on the aerodynamic coefficients can
be seen in figure 12 for _] = 0°. The data for this
wing indicate a very slight loss in lift neat" eL,max,which probably resulted from early separation of the
flow downstream of the boundary-layer trip. Pitch
stability was also unaffected by tripping the bound-
ary layer. This was expected because the airfoil wasdesigned so that lift and pitching moment would be
unaffected by premature transition. Figure 12(b)
shows the effect of fixed transition on drag. There
was an increase in CD of 0.003 near CL = 0.3 when
the flow was tripped at x/c = 0.05. Data from two-
dimensional tests (ref. 7) indicated a similar increase
in CD for similar test conditions (Reynolds numberof 3.6 × 106 at M < 0.3). The effect of transition on
tile lift and pitching-moment characteristics for the
flap deflected 40 ° is shown in figure 13. The addi-
tion of a transition strip did not significantly affect
the longitudinal characteristics because very little, if
any, laminar flow existed on the wing at such an ex-
treme flap deflection. This result indicates that land-
ing characteristics do not depend strongly on wingsurface conditions.
Hot-film sensors were used to determine the ex-
tent of laminar flow on the wing with natural transi-
tion. A sample set of oscillograph tracings from the
mid-semispan, hot-film sensors on the lower wing sur-
face at a total CL of 0.22 is shown in figure 14. Thedata were taken over a time period of 1.5 sec and
show that the boundary-layer characteristics change
rapidly over the chord. The tracing from the sen-
sor at the 10-percent chord station shows that the
boundary layer was laminar essentially 100 percent
of the time. At points farther aft along the chord,
the flow was mostly laminar with turbulent bursts.
At 50 percent chord the boundary layer was turbu-
lent 50 percent of the time. Aft of that location, the
flow was predominantly turbulent with some laminar
bursts. The transition point was the chord location
where the boundary layer was turbulent 50 percent
of the time. Therefore, for this example, laminar flow
existed back to 50 percent chord.
Figure 15 shows the extent of laminar flow on the
upper and lower surfaces at three spanwise locations
as a function of total lift coefficient. Figure 15(a)
shows that there was significant laminar flow on theupper surface for lift coefficients that range from-0.40 to about 0.22. A reduction in laminar flow
was seen with increasing C L above 0.22. (The nextdata point was at CL = 0.31, so the initial reduction
in laminar flow could conceivably occur at a lift coef-
ficient that is slightly higher than 0.22.) Figure 15(b)
shows that there was essentially no laminar flow onthe lower surface for lift coefficients below 0.22 but
that there was significant laminar flow on the lower
surface for lift coefficients greater than 0.22. Com-
bining the results for the upper and lower surfaces
indicates that the optimum CL for maximizing the
extent of laminar flow over the wing is about 0.22 orslightly higher.
Because of three-dimensional effects (e.g., the
variation of local Reynolds number along the span),the extent of laminar flow was not constant along the
span. The maximum amount of laminar flow on the
upper surface was 60 percent of the chord (outboard
station), and the minimum amount was 35 percent
of the chord (inboard station). The average amountof laminar flow on the upper surface was less than
50 percent. The lower surface characteristics were
similar, with a laminar flow maximum of 70 per-
cent (outboard station), a minimuln of 30 percent
(inboard station), and an average of about 50 per-cent. The correlation of the hot-film results with the
favorable pressure gradient characteristics discussedpreviously was good. This correlation indicates that
the extent of laminar flow was not being significantlyaffected by small surface irregularities. The amount
of laminar flow in the two-dimensional tests (ref. 5)was 40 to 50 percent on the upper surface and 50 to
60 percent on the lower surface.
Effect of Flap Deflection
The single-slotted flap used in these tests was
designed with the MCARF (Multi-Component Air-
foil Analysis Program) computer code (ref. 8), which
was developed to analyze two-dimensional airfoil-flap
configurations. Because this code cannot accurately
modeltrailing-edgeseparation,it relieson the as-sumptionoflaminarseparationnearthe leadingedgeto predictstall. Thus,the codetendsto overpre-diet CL,ma x when the trailing edge stalls first. Testresults from an evaluation of flap overlap and gap
settings with the flap deflected 40 ° indicated thatthe optimum flap location predicted by the theoreti-
cal code did achieve the best CL,ma x experimentally(0-percent overlap and 2-percent gap). All data pre-sented with the flaps deflected 40 ° were obtained with
the optimum gap and overlap values. The overlapand gap for the 20 ° flap deflection, which were based
on the MCARF predictions, were fixed at -3 and4 percent.
Longitudinal characteristics for the semispanmodel for flap deflections of 0°, 20 °, and 40 ° with the
vortex generator mounted are presented in figure 16.
With 0° flap deflection, a CL,max of 1.45 was achievedat an angle of attack of about 17 ° (fig. 16(a)). Tile
pitching-moment data showed no unstable breaks in
the curves and generally showed linear behavior up
to CL,ma x. The CL data showed that sizable incre-ments in lift were achieved by deflecting the trailing-
edge flap (CL,ma x = 2.44). This result is representa-
tive of other single-slotted flap designs. As expected,
the value of CL,max that was predicted by the codewas higher than the measured value; however, the in-
crement in lift due to flap deflection was accuratelypredicted.
The increase in lift seen in figure 16(a) resulted in
a corresponding increase in nose-down pitching mo-ment that was fairly linear with flap deflection at
low and moderate angles of attack. Flap deflection
had little effect on the overall level of pitch stabil-
ity below 15 ° angle of attack. The pitching-moment
break at the stall with the flaps deflected 40 ° , how-
ever, was slightly unstable in contrast to the flaps-
undeflected configuration, which showed a slight sta-ble break at stall. As expected, increasing the flap
deflection increased the drag for a given value of CL.
(See fig. 16(b).) Unlike the lift and pitching-moment
data, the changes in drag were nonlinear with flapdeflection.
Effect of Vortex Generator
To alleviate premature flow separation on the in-
board section of the wing, a fuselage-mounted vortex
generator was investigated. (See fig. 4.) This pre-
mature flow separation was most clearly seen in tuft
studies with the flaps deflected 40 ° . (See fig. 17.)
Without the vortex generator, a large portion of theinboard panel was stalled. The addition of the vortex
generator greatly reduced the region of stalled flow
and delayed the progression of separated flow from
the inboard to the outboard panel to higher anglesof attack.
Figure 18 shows the effect of the vortex gener-
ator on the longitudinal characteristics of the semi-
span model with the flap undeflected. As can be seen
from the data, the addition of the vortex generator
improved CL,ma x and delayed the stall by 2° angleof attack. Pitching moment and drag (figs. 18(a)
and 18(b)) were unaffected by the vortex generator,
except for negligible changes in the stall and post-
stall angle-of-attack regions. The effect of the vor-
tex generator with the flaps deflected 40 ° was more
pronounced. (See fig. 19.) With the flaps deflected,
maintaining attached flow over the inboard portion
of the wing through the use of the vortex generator
had a large beneficial effect, which increased CL,maxfrom 2.3 to 2.45. As with the flaps undeflected, pitch-
ing moment was essentially uimffected by the vortex
generator.
Effect of Leading-Edge Droop
The concept of incorporating a discontinuous
leading-edge droop on the outboard portion of a wing
to improve stall/spin resistance characteristics has
received much attention in recent years (refs. 9, 17,
and 18). The purpose of the droop is to maintain
attached flow on the outer wing panel to higher an-gles of attack to soften the stall break and to im-
prove roll damping and roll control in the stall and
post-stall regions. These improvements help to bothprevent and control any autorotative tendencies of
the aircraft. In most previous applications, a singleoutboard droop segment has been sufficient to pre-
vent autorotative tendencies. (See ref. 17.) However,recent studies have shown that wings with relatively
high aspect ratios require an additional segment lo-
cated farther inboard to achieve the desired stall/spin
resistance characteristics. As mentioned previously,
the leading-edge droop for the subject configuration
was developed on two subseale models, and the droopconfigurations were evaluated with both static force
data and roll-damping data. Results from these tests
(ref. 17) showed that the selected droop should pro-
vide the desired stall/spin resistance characteristics.
Figure 20 shows the effect of the droop on the flap-
retracted configuration that was measured during the
full-scale semispan model tests. The addition of the
droop did not significantly improve the lift character-istics until well into the post-stall region. The fact
that the desired "flat top" lift curve was not obtained
suggests that the droop design may not provide the
improved stall characteristics shown in the subseale
tests. However, the effectiveness of a leading-edge
droop is not always apparent from an examination
of total lift characteristics. In most cases, the ef-
fect of the droop on static characteristics can best be
seen by separately measuring the forces on the outer
panel of the wing. This approach was used during
the original droop design but was impractical for the
semispan tests. Because the droop was developed on
a wing without twist, without a winglet, and at very
low Reynolds numbers, the design might not have
been properly optimized for the full-scale semispan
model. The effect of the droop on drag was negligi-
ble around the design cruise lift coefficient of 0.26.
Thus, the goal of maintaining natural laminar flow
on the drooped portion of the wing was achieved.
Roll-Control Effectiveness
Roll-control data in the form of rolling-momentincrements are shown in figures 21 to 23. Data for
the wing without the droop and with the flap un(te-
fleeted (fig. 21) show that the aileron was effective in
providing roll control well past CL,max. Additionalroll control was obtained by deflecting the spoiler.
Once the wing stalls in the spoiler region, the spoiler
becomes ineffective. With the flaps deflected, the
additional rolling moment from the spoiler was evenmore pronounced (fig. 22). Because the spoiler was
in front of the flap (fig. 2) the larger lift increment
generated by the flap was lost when the spoiler was
deflected, thus creating these large rolling moments.
Since the aileron was outboard of the flap, the effec-
tiveness of the aileron was only slightly affected by
the deflection of the flap (fig. 22). The effect of theleading-edge droop on the rolling moment is shown
in figure 23 for (_f _ 0° and _a = 20 °. Tim data showa sizable improvement ill roll control above (x = 5°
with the addition of the droop. This result indi-
cates that the droop was affecting the flow on the
outboard portion of the wing, even though the to-
tal lift data stlowed no effect from the droop. This
result indicates a delay in separation of the flow on
the outboard panel and illustrates that changes inroll control can be used to some degree to evaluate
droop effectiveness when a metric outboard panel is
impractical to use.
Summary of Results
A wind-tunnel investigation of the high-lift char-
acteristics of a full-scale senfispan wing with a natu-
ral laminar flow airfoil was conducted in the Langley30- by 60-Foot Tunnel. The results of these tests canbe summarized as follows:
1. Deployment of a single-slotted, trailing-edge flapprovided substantial increments in lift without
significant adverse effects on static pitch stability.
6
These results suggest that an NLF wing designedfor efficient cruise at Mach 0.7 can also be viable
at low-speed takeoff and landing conditions.
2. Fixed-transition studies indicated no adverse ef-
fects on lift and pitching-moment characteristics
as a result of the transition of the boundary layer
near the leading edge for either the cruise or land-ing configuration.
3. Maximum lift coefficient was improved by using a
fuselage-mounted vortex generator to delay sepa-ration of the flow on the inboard portion of the
wing.
4. Although roll-control studies indicated some de-
lay in outboard-panel flow separation, the full-
scale data suggest the need for further optimiza-
tion of the leading-edge droop design to provideimproved stall/spin resistance.
5. The wing aileron provided good roll control up to
stall. Additional rolling moment was generatedby a nfidspan spoiler for flap deflections of 0°
and 40 °. The effectiveness of thc spoiler was
significantly greater with the trailing-edge flapdeflected.
NASA Langley Research CenterHampton, VA 23665-5225August 6, 1991
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and Frances E. Sabo, compilers, NASA CP-2487, Part 3,
1987, pp. 697 726.
8. Stevens, W. A.; Goradia, S. H.; and Braden, J. A.: Math-
ematical Model for Two-Dimensional Multi-Component
Airfoils in Viscous Flow. NASA CR-1843, 1971.
9. Staff of Langley Research Center: Exploratory Study of
the Effects of Wing-Leading-Edge Modifications on the
Stall/Spin Behavior of a Light General Aviation Airplane.
NASA TP-1589, 1979.
10. DeFrance, Smith J.: The N.A.C.A. Full-Scale Wired Tun-
nel. NACA Rep. 459, 1933.
11. Theodorsen, Theodore; and Silverstein, Abe: Experimen-
tal Verification of the Theory of Wind-Tunnel Bottndary
Interference. NACA Rep. 478, 1934.
12. Heyson, Harry H.: Use of Superposition in Digital Com-
puters To Obtain Wind-Tunnel Interference Factors for
Arbitrary Configurations, With Particular Refer_nce to
V/STOL Models. NASA TR R-302, 1969.
13. Heyson, Harry H: FORTRAN Programs for Calculating
Wind-Tunnel Boundary InteTference. NASA TM X-1740,
1969.
14. Heyson, Harry H: Nomographic SolutioT_ of the Momen-
turn Equation for VTOL-STOL Aircraft. NASA TN D-
814, 1961.
15. Heyson, Harry It.: Rapid Estimation of Wind-Tunnel
Correctiorts With Application to Wind-Tunnel a_d Model
Design. NASA TN D-6416, 1971.
16. Rac, William H., Jr.; and Pope, Alan: Low-Speed Wind
Tunnel Testing, Second ed. John Wiley & Sons, Inc.,
c.1984.
17. Johnson, Joseph L., Jr.; Yip, Long P.; and Jordan,
Frank L., Jr.: Preliminary Aerodynamic Design Consid-
erations for Advanced Laminar Flow Aircraft Configu-
•rations. Laminar Flow Aircraft Certification, Louis J.
Williams, compiler, NASA CP-2413, 1986, pp. 185 225.
18. Stough, H. Paul, III; Jordan, Frank L...Jr.; DiCarlo,
Daniel J.; and Glover, Kenneth E.: Leading-Edge Design
for Improved Spin Resistance of Wings Incorporating
Conventional and Advanced Airfoils. Langley Symposium
on Aerodynamics, Volume II, Sharon H. Stack, compiler,
NASA CP-2398, 1985, pp. 141 157.
7
Table I. Geometric Characteristics of Semispan Model
Wing:
Semispan area, ft 2 ................................... 125.00
Semispan, ft ...................................... 22.36
Mean aerodynamic chord, ft ................................ 6.15
Root chord (centerline), ft ................................. 8.28
Tip chord, ft ....................................... 2.91Aspect ratio of semispan .................................. 4.00
Wing incidence (root), deg ................................. 2.00
Dihedral angle of leading edge, deg ............................. 5.50
Leading-edge sweep angle, deg ............................... 4.15
Fuselage station of wing leading edge (centerline), in ..................... 225.18
Airfoil ................................. NASA HSNLF(1)-0213
Flap:
Area, ft 2 ........................................ 24.30
Inboard wing station, in .................................. 32.20
Outboard wing station, in ................................. 211.99
Chord, percent c .................................... 28.00
Aileron:
Area, ft 2 ......................................... 4.39
Inboard wing station, in .................................. 211.99Outboard wing station, in ................................. 264.86
Chord, percent c .................................... 28.00
Hinge line, percent c .................................. 78.00
Spoiler:
Area, ft 2 ......................................... 3.20
Inboard wing station, in .................................. 114.05
Outboard wing station, in ................................. 173.62Chord, percent c .................................... 12.00
Hinge line, percent c .................................. 78.00
Trailing edge, percent c ................................. 92.00
Fuselage:
Length, ft ....................................... 35.00Radius (maximum), ft ................................... 2.50
Winglet:
Area, ft 2 ......................................... 0.65
Span, ft ......................................... 0.88
Root chord, ft ...................................... 1.06
Tip chord, ft ....................................... 0.43
Aspect ratio ....................................... 1.19
Cant angle (outboard), deg ................................ 20.50
Leading-edge sweep angle, deg .............................. 45.00
Airfoil ..................................... NACA 0012-35
TableII. HSNLF(1)-0213Airfoil Coordinates
x/c
0.00000
0.00123
0.00270
0.00499
0.00801
0.01177
0.01627
0.02147
0.03389
0.04104
0.048810.06619
0.07577
0.08992
0.09661
0.10783
0.11955
0.13176
0.14444
0.15759
0.17119
0.19963
0.21445
0.24519
0.277230.29370
0.31042
0.34457
0.36193
0.37942
0.397050.41481
0.43267
0.45O56
0.48633
0.50416
0.52193
0.55725
0.57474
0.60918
0.62604
0.642630.67493
0.70586
0.73539
0.75006
Upper surface
z/c
0.00000
0.00676
0.01001
0.01350
0.01688
0.01995
0.02309
0.02602
0.03167
0.03438
0.03705
0.04222
0.04470
0.04711
0.04945
0.O517O0.05387
0.05596
0.05796
0.05985
0.06166
0.06494
0.06641
0.06901
0.07110
0.07195
0.072660.07365
0.07392
0.07405
0.07402
0.07384
0.073500.07299
0.07147
0.07043
0.06920
0.06605
0.06411
0.05940
0.05664
0.053620.04699
0.03996
0.03300
0.02953
Lower surface
z/c
0.00000
-0.00257
-0.00450
-0.00713
-0.00983
-0.01191
-0.01399
-0.01603
-0.01993
-0.02181
-0.02364
-0.02724
-0.02902
-0.03078
-0.03254
-0.03427
-0.03597
-0.03765-0.03928
-0.04088
-0.04242
-0.04532
-0.04670
-0.04933
-0.05174
-0.05286
-0.05392
-O.O5583
-0.05667-0.05741
-0.05807
-0.05864
-0.05911
-0.05948-0.05990
-0.06000
-0.06003
-0.05969-0.05933
-0.05821
-0.05743
-0.05645
-0.05385-0.05002
-0.04450
-0.04115
TableII. Concluded
x/c
0.77995
0.79503
0.84008
0.86896
0.88077
0.90712
0.94150
0.96027
0.97556
0.98718
1.00000
Upper surface
z/c
0.02256
0.01914
0.009410.00372
0.00154
-0.00298
-0.00812
0.01057
-0.01244
-0.01380
-0.01528
Lower surface
z/c
-0.03565
-0.03365
-0.02890
-0.02630
-0.02535
-0.02315
-0.02040
-0.01895-1.01800
-0.01750
-0.01660
10
TableIII. DroopedAirfoil Coordinates
x/c
-0.02000
-0.01975
-0.01950
-0.01900
-0.01850
-0.01800
-0.01750
-0.01500
-0.01000
0.00000
0.01000
0.02000
0.03000
0.04000
0.05000
0.06000
0.08000
0.09000
0.100000.12500
0.15000
0.15759
0.17119
0.19963
0.21445
0.24519
0.27723
0.293700.31042
0.34457
0.36193
0.37942
0.39705
0.41481
0.43267
0.450560.48633
0.50416
0.52193
0.55725
0.57474
0.60918
0.62604
0.64263
0.674930.70586
Upper surface
-0.00988
-0.00792
-0.00636
-0.00575
-0.00467
-0.00369
-0.00279
0.00106
0.00686
0.01468
0.02075
0.02584
0.03024
0.03407
0.03749
0.04054
0.04576
0.04804
0.05013
0.054800.05876
0.05985
0.06166
0.06494
0.06641
0.06901
0.07110
0.07195
0.072660.07365
0.07392
0.07405
0.07402
0.07384
0.07350
0.07299
0.071470.07043
0.06920
0.06605
0.06411
0.05940
0.05664
0.05362
0.04699
0.03996
Lower Sllrface
z/c
-0.00988
-0.01157
-0.01285
-0.01332
-0.01410
-0.01475
-0.01531
-0.01741
-0.02007
-0.02329
-0.02540
-0.02697
-0.02824
-0.02934
-0.03033
-0.03126
-0.03307
-0.03399
-0.03494
-0.03745
-0.04014
-0.04088
-0.04242-0.04532
-0.04670
-O.O4933
-0.05174
-0.05286
-0.05392-0.05583
-0.05667-0.05741
-0.05807
-0.05864
-0.05911
-0.05948
-0.05990
-0.06000
-0.06003
-0.05969
-0.05933-0.05821
-0.05743
-0.05645
-0.05385
-0.05002
11
TableIII. Concluded
x/c
0.73539
0.750060.77995
0.79503
0.840080.86896
0.88077
0.90712
0.94150
0.96027
0.97556
0.98718
1.00000
Upper surface
z/c
0.03300
0.02953
0.02256
0.01914
0.00941
0.003720.00154
-0.00298
-0.00812
-0.01057
-0.01244
-0.0138O
-0.01528
Lower surface
z/c
-0.04450
-0.04115
-O.O3565
-0.03365
-0.02890
-0.02630
-0.02535
-0.02315
-0.02040
-0.01895-O.018OO
-0.01750
-0.01660
12
Table IV. Chordwise Location of Pressure Ports
I Wing, aileron, and spoiler port locations given in percent of local wing chord; 1flap port locations given in percent of local flap chord
Station 1 Station 2 Station 3 Station 4 Station 5 Station 6 I Station 7 Station 8
Wing upper surface
2.5
5.0
7.5
10.0
15.0
20.0
25.0
30.0
40.0
50.0
60.0
70.0
80.0
90.0
2.5
5.0
7.5
10.0
15.0
20.0
25.0
30.0
40.0
50.0
60.0
70.0
80.0
90.0
2.5
5.0
7.5
10.0
15.0
20.0
25.0
30.0
40.0
50.0
60.0
70.0
79.0
2.5
5.0
7.5
10.0
15.0
20.0
25.0
30.0
40.0
50.0
60.0
70.0
79.0
2.5
5.0
7.5
10,0
15.0
20.0
25.0
30.0
40.0
50.0
60.0
70.0
79.0
2.5
5.0
7.5
10.0
15.0
20.0
25.0
30.0
40.0
50.0
60.0
70.0
80.0
90.0
2.5
5.0
7.5
10.0
15.0
20.0
25.0
30.0
40.0
50.0
60.0
70.0
2.5
5.0
7.5
10.0
15.0
20.0
25.0
30.0
40.0
50.0
60.0
70.0
Wing lower surface
2.5
5.0
7.5
10.0
15.0
20.0
25.0
30.0
40.0
50.0
60.0
70.0
2.5
5.0
7.5
10.0
15.0
20.0
25.0
30.0
40.0
50.0
60.0
70.0
2.5
5.0
7.5
10.0
15.0
20.0
25.0
30.0
40.0
50.0
60.0
70.0
2.5
5.0
7.5
10.0
15.0
20.0
25.0
30.0
40.0
50.0
60.0
70.0
2.5
5.0
7.5
10.0
15.0
20.0
25.0
30.0
40.0
50.0
60.0
70.0
2.5
5.0
7.5
10.0
15.0
20.0
25.0
30.0
40.0
50.0
60.0
70.0
2.5
5.0
7.5
10.0
15.0
20.0
25.0
30.0
40.0
50.0
60.0
70.0
2.5
5.0
7.5
10.0
15.0
20.0
25.0
30.0
40,0
50.0
60.0
70.0
Flap upper surface
2.49
4.98
9.96
19.91
28.88
49.78
72.88
90.70
2.50
5.00
10.00
20.00
28.89
49.98
73.16
91.06
2.50
5.00
10.00
20.00
29.01
50.02
73.22
91.13
2.51
4.99
10.02
20.04
29.07
50.11
73.36
91.29
2.51
5.02
10.03
20.06
29.09
50.16
73.43
91.39
2.52
5.03
10.06
20.11
29.16
50.28
73.61
91.61
13
Table IV. Concluded
Station 1 Station 2 Station 3 Station 4 Station 5 Station 6 Station 7 Station 8
Flap lower surface
2.49
4.98
9.96
19.91
28.88
49.78
64.02
72.88
90.70
2.50
5.00
10.00
20.00
28.89
49.98
64.27
73.16
91.06
2.50
5.00
10.00
20.00
29.01
50.02
64.16
73.22
91.13
2.51
4.99
10.02
20,04
29.07
50.11
64.,'14
73.36
91.29
2.51
5.02
10.03
20.06
29.09
5{). 16
64.50
73.43
91.39
2.52
5.03
10,06
20.11
29.16
50.28
64.66
73.61
91.61
Aileron upper surface
Aileron lower surface
82.91
91.45
93.59
97.86
76.60
88.30
91.23
97.08
82.91 76.60
91.45 88.30
93,59 91.23
97.86 97.08
Spoiler surface
90.{)1 92.84 9{).03
14
ORI_,,,AL PAGE
BLACK AND WHITE PHOTOGRAPH
(a) Cruise configuration.
L-87-300
L-87-1234
(b) Flap deflected 40 °.
Figure 1. Photographs of modelin Langley 30- by 60-Foot Tunnel.
15
Winglet
Aileron _
Sp°iler_ / 6 5.
Single-slotted flap _/ ! /_ 2.50_ radius
18.03
FS 35
Figure 2. Geometry of semispan model. All dimensions in feet unless otherwise noted.
,1
z/c 0 ----_..____.
-,1 , 1 , 1 , I l l l ] J l _ 1L I A I l I0 ,1 ,2 ,3 ,4 ,5 ,6 ,7 ,8 ,9 1,0
X/C
Figure 3. Section shape for HSNLF(1)-0213 airfoil.
16
5.25
18.25
/lm
r"
t
I
!w
1I
9.0
Vortex gener_
1.4 t'-
I 2.5
%
Figure 4. Geometry of vortex generator. All dimensions in inches unless otherwise noted.
.72c .92c
0
Figure 5. Geometry of trailing-edge flap.
17
.1
z/c 0 1
",1
-.1
_Drooped le__9 - edge
_"f Origin _'1a--_ffoil
I I I I I0
I I.1 .2 .3 .4 .5 .6
x/c
WS 188.6 WS 108.8
Iws 671WS 263.0
I
I I I.7 .8 .9
WSO.O
I1.0
Figure 6. Geometry of leading-edge droop. Planform view not to scale.
Sensor locationsUpper surface
0.05c.10c.20c.30c.40c.50c.60c.70c
Lower surface0.10c
.20c
.30c
.40c
.50c
.60c
.70c
Outboard lsensors
80% b/2
oo o
° ° ° lMid-semispan
sensors50% b/2
Inboardsensors
Figure 7. Locations of hot-film sensors.
18
%
1,5 ....
-1,0_-
0--
.S--.2
Station 1,
1
.... £_ .....
.2 .q .6 .8 1.0
x/cy/C./2) =o.2o
1.5.___ _ i!i
!
]
°I-- -_3" '
2 _ ,2 .q .6
x/cStation _, y/(b/2)
....73
.8 1.0
= 0.55
-1 .S ___
-1.0 ---- --]
.5 ....
2 ° '
•_._-- o .2 ._ .6 .8 1.ox/c
Station _,, y/_/2) =0.85
I,S .......
-l ,0 ......
% .s
;(_'"
I
.2
i
r
0 .2 .4 .6 .B
x/c
Station 2, y/(b/2) = 0.35
1.0
-I .5 _ _
-i.ot - --_ ---
% -%®%,-i! o
rL i _ [
.2 0 ,2 .4 .6
x/cStation S, 7/(b/2)
_e
[]
• 8 I .O
= 0.62
C -2 -q -6 .8
x/cStation 8, y/(b/2) = 0.95
%
-I .5
i.0 .............. .........
I
"-.2 0 .2 ,q .6 ,8 I .0
x/cStation 3, y/(b/2/ = o.45
Figure 8.
-1 .S
% _.
---i
D
.2 ._ .6
x/cStation 6, ¥/(b/2)
• 8 I .0
= 0.75
0 Main wing
Z] Flap
0 Aileron
Spoiler
, Lower surface
(a) a=-2.2 ° .
Pressure distributions for clean wing. 6f -- 0°;VG off.
19
-I .S
-t -0
%,5
0
'_5.2
d#'-*D
0 .2 .q
x/c
8
.6 .8 t .0
Station 1, 7/(b/2) = 0.20
-1.5
-I .0
%-,5
0
'5.2
.._,-oo( )00 (
,,,,.a_® ( )(9(9 _
FO .2 .4
ocb[
e¢¢
.6 .8 1.0
x/cStation 4, ,//('o/2) = 0.55
1.S ......
Station _,, y/(b/2) = 0.85
-I .5 ___
-_ .0 -
-.S_0( )(DO 0
_®( ®® ®
O .2 .q
x/c
®
[]
!.6 .8 I .0
Station 2, ',//(b/2) = 0.35
-I .5
-1.0
-.5
0
._s2
_,or,c',( )00 O
._ff.,® ( ®® ®
I0 .2 .q ,6
x/c
_n
.8 I .O
Station 5, f/(b/2) = 0.62
-t .5 F .......... - _ .... • T
I.O ...... '
% _
'-5.2 o .2 ._-- .s .s _.ox/c
Station 8, y/(b/2) = 0.95
%
-I .S __ --
-I .O
.5
0
"52
I [i
4_o()Oo $ o
zao)®( ®® ®
J0 .2 .4
x/cStation 3, y/(b/2) = 0.45
L _
,6 ,8 1 .O
-1.5
-t .0 ----
-.5
"?.2
]
_ rh( )(30 0 1_u-
C_®( ®® ®:9
!0 .2 .4 .6 .8
x/cStation 6, y/(b/2) = 0.75
%
(b) c_=1.5 ° .
Figure 8. Continued.
1.0
0 Main wing
G Flap
0 Aileron
A Spoiler
+ Lower surface
20
%- O0
o____ 2 _.__2 q .6 .8 1.0
x/cStation 1, y/(b/2) =0.20
% _D
-2 - -_)
'00 q _0 _ __
i
-2.2 o .2 . .6 .8 1 .o
x/cStation q, y/(b/2) = 0.55
%
2-.2
_Oi,oo < t
®®®I
0 .2 .q .6 .8 .0
x/cStation 7, ¥/(b/2) = 0.85
%0( )00
( )®tv
I• 2 .q
x/cStation 2, y/(b/2) = 0.35
0
0 _.z__
• 6 .8 .0
%
-2.2 o .2 .4 .6 .a 1.o
x/cStation 5, 7/('o/2) = 0.62
%
-6
_q ..... _
-:ri.4oox/c
Station 8, y/Co/2) = 0,95
%
6
-2 i _0%0( )00
0
-2.2 o .2 .q ,6 ,8 1.0
x/cStation 3, ¥/(b/2) = 0.45
%
-6
)00 ) 0 _ G_.
......-2.2 0 .2 .q .6 .8 1.0
x/¢Station 6, _//(b/2) =0.75
0 Main wing
[] Flap
0 Aileron
A Spoiler
+ Lower surface
(e) a = 9.0 °.
Figure 8. Continued.
21
%
%
%
-_F _ F T T -
-2.2 - • .6 .8 ! .0
x/¢Station I, y/_/2) =0.20
-6
!
-q
0
_._ o .2 ., ._ .8 _.ox/c
Station 2, yl(b/2) =0.35
-6
22 0 2 4 6 .8 0
x/cStation 3, y/(b/2) = 0.45
x/cStation 4, ¥/(b/2) =0.55
%
-6
-q
co -2
o
_2.2
T 1____ _' i
ot:J1
D0
0 0 -)0 O
.3_i)®( )®®
0 .2 .q .6 .8
x/cStation 6, y/(b/2) = 0.75
1.0
%-2
-.2
[_ J
O .2 .q ,6 ,8 1.0
x/cStation "7 y/Co/2) : 0.s5,j
-6
olx/c
Station 8, yflb/2) =0.95
© Main wing
[3 Flap
{> Aileron
& Spoiler
+ Lower surface
(d) e= 12.8 ° .
Figure 8. Continued.
22
%
6[,. ---
-q
-2- I dr'o,
o t9
_J ....
?.2 o .2
Station 1,
t.L4 ,6 -8 1 .0
x/cv/_/2) = o.2o
%
6
-q
2
0
-,2
i0
o !
Q) O_ )oo ) 0 ) aJ
®_)® ,®® ® -_ _e_
o .2 ._ .6 .B _.ox/c
Station q, y/(b/2) =0.55
%
-6 _
°I ',i°22-
x/cStation 7, _//(b/2)
J
• 8 1.0
= 0.85
%
x/cStation 2, yflb/2) = 0.35
%
=.2
Io
-- d - _0,
8 ,2 .q ,6
x/cStation S, _/(b/21
o__o_,-8 I 0
= 0.62
%
-6
x/cStation 8, ¥/(b/2)
• 8 I ,O
= 0.95
%
t °tOotoio
x/cStation 3, ¥/(b/2) = 0.45
Cp
q
2
0
Pio
- -_ " "Oc
®®_
____+ ....
_Oo (_ O (_ t_a fib _u,=,
-L2_ ..... .2 _ .6---._x/c
Station 6, ,//(b/2) = 0.75
(e) _= 14.8 ° .
Figure 8. Continued.
10
0 Main wing
[3 Flap
<} Aileron
A Spoiler
+ Lower surface
23
%
-6
-2
0
2_.2
D
00
0 I
O( I00 l 0
_(_@)®_®® TI _-
0 .2 .4
x/cStation 1, y/(b/2) = O.2O
7-I
.6 .8 I .0
%
-6
-4
-2
0
_2.2
D
0 .2 .q
!
•6 .8 I .0
x/cStation q, y/(b/2} = 0.55
%
-6
_t4
-2
0
-2.2
o%°'>°° )o
®®®_®® ) ®
I•2 .4
x/cStation 7, ¥/(b/2) = 0,85
J.6 .8 I.O
%
-6
!i_2.2
o t%°'>°o¢ o ! o_m!j
I i0 .2 .q .6 .8 1.0
x/cStation 2, f/(b/2) = 0.35
-6
0
_2.2
0 I
000_00 ) 0
) tti]_ _ r_
0 .2 ._ .6 .8
x/cStation S, 1/(b/2) = 0.62
1.0
_ __
0
0 .2 .q .6 .8
x/cStation 8, ¥/(b/2) = 0.95
1.0
-6
%-'i
0
0 .2 .q .6 .8
x/cStation 3. ¥/tb/2) = 0.45
1.0
-6
O
% -] 0%)Oo
0
O ®_i) ®_ (y® ® ) _-_
_12 o .2 .4 .5 .ox/c
Station 6, ¥/(0/21 = 0.75
(f) a = 16.8 °.
Figure 8. Concluded.
1,0
0 Main wing
13 Flap
Aileron
A Spoiler
+ Lower surface
24
%
-6
_u,
-2
0
.2
O, ) 0 0
_ 1
0 .2 ,4
x/cStation 1, _//(b/2)
E:--
>o _o
-6 .8 1,0
= 0.20
%
6 .....
E
/ | (i []
2.p o---U--_ .6 .8-7.0x/c
Station q, y/{b/2) = 0.55
%
6_
2._ o .2 ._ .6 .B _.ox/c
Station 7, ¥/(b/2) = 0.85
%-2 .... E
_o,,oo 4 o _o _o_ i iI I
-2-2 0 .2 .q ,6 .8 1.0
x/cStation 2, ¥/_/2) = 0.35
%
-6
O 2 4 0
x/cStation 5, ¥/(b/2) = 0.(_2
%
t
i
-2 -----+-.--q J
-2.7 U--12 -._ ._ .8x/c
Station 8, ¥/(b/2) =0.95
J
1.0
%
-8
:7I
x/cStation 3, ¥/(b/2) = 0.45
%
-_rI
_q ....
2 ......
_DO0<
0 - (_®(
Station 6,
]
)00 () 0 (_ 0
.q .6 .8 1.0
x/c_//_)/2} = 0.75
0 Main wing
[] Flap
0 Aileron
A Spoiler
+ Lower surface
(a) c_ =-1.3 ° .
Figure 9. Pressure distributions with flap deflected. 6 r = 40°; VG on.a
25
%
_q_ A
J
-2
o- o
_® ®o _ ® _
• 2 -tt .6 .8 1-0
x/cStation l, y/(b/2) =0.20
%
-6F- -
q
-2
-.2
[3
0CLq
rqrq
O)OO 0
0 .2 .q .6 .8 1 .O
x/cStation q, ¥/(b/2) = &55
-6
q
Cp-2
0
2.2
1J
[0 .2
r
......i
• q .6 .8 ; .0
x/cStation 7, y/00/2) = 0.85
%
-6__
ti .... i
0 O+oo o
22 0 2 ti 6
x/cStation 2, y/(b/2) =0.35
• 8 1.0
-6
!
-4 j -_ --i
Cp
-2 FQ_o _OO 0
.2-- !(_i_12 .4 .6 .8 t .0
x/cStation S, y/(b/2) = 0.62
[]
O AC[3
Rn
6
-4
Cp 2
O_
2. 2
T
0 .2
i• q .6 .8 ! .0
x/cStation 8, y/(b/2) = 0.95
%
6_. _
D
%o<,oo i o
°_-2Li _®®_®- -2 0 .2 .4 ,g .8 1.0
x/cStation 3, y/(b/2) = 0.45
E
0 _ A
6 __
-q
Cp -2
--.2
[]
+_OO( _00 0
o ,_o
_®,,_ ¢, i_1_ n
LO .2 .q .6 .8 1.0
x/cStation 6, y/(b/2) = 0.75
0 Main wing
[3 Flap
O Aileron
£x Spoiler
+ Lower surface
(b) o* = 2.4 °.
Figure 9. Continued.
26
-6
-q
Cp -2
0
2.2
D
u
_o c
#)69(
o o_
_®® ( ®
o0 ( 0
D D
•2 .q ,6 .8
x/cStation 1, 7/(b/2) =O.20
1.0
B
_.tt ---
Cp 2
©
O
%,w_)LXD ) 0 )
0
___L2.2 o .2 ._ .6
x/c
rlo
E
O A[]
Ek-
J•8 I 43
Station q, 7/(b/2) = 055
-B
-4
Cp -2
O
%0 _oo o
O ©
_2._-- .2 .4x/c
Station 7, y/(b/2) = 0.85
)
•6 .8 I .O
Cp
-6
-q
-2
0
_2.2
0
qDn._u 0
_®( )®® ®
i
0 .2 .4
x/c
&
[]
o o 8[]o
i• 6 ,8 1.0
Station 2, y/(b/2) = 0,35
-6!i --7
Cp .2 ----
0--
::'.7--
[]> I
,oo _, o 0VIF
• 2 .q .6 ,8 1.3
x/cStation S, _//(b/2) = 0.62
Cp
-6
x/cStation 8, ¥/(b/2) = 0.g5
-6 _
t
Cp -:
2. 2
oo
_®+®®
I.2
D[]
K0
0¢ A0
o_
.½ .6 .8 1.0
x/cStation 3, _//(b/2) = 045
-8
4
Cp -2
o
_.2
*D t)( )00 0
®® ®
%[]
o )co i[3E
0 ,2 .q .6 .8
x/cStation 6, ¥/(b/2) =0.75
1.0
O Main wing
[] Flap
0 Aileron
A Spoiler
+ Lower surface
(e) a = 6.1 °.
Figure 9. Continued.
27
15
Cp -s
°[_s. 2
I
!i
0
[(_bo{) 0 0 , C_ _q
T-T0 .2 .q .6 .8 1.0
x/cStation 1, y/(b/2) =0.20
Cp
15____
-I0
-5o
t_
"%O,)oo ) o o_ _.---_,,<_ _ _ _ _t__
:.2 o .2 .4 .6 .8 _.ox/¢
Station q, y/_o/2) = 055
Cp
-15
-I0
-5
0
_s.2
OqgO )00 m0
0 .2 .q
x/c
JI
•6 .8 ] .0
Station 7, y/(b/2) = 0.85
-15
-10
Cp -s
-.2
%o [2
}00 0 C_,
I0 .2 ._ .6 .8 1.0
x/cStation 2, T/(b/2) = 0.35
Cp
-iS -iS
I0 -tO
Cp-s ._ % -s ...........
%O,oo( o( o_=. olL
/-s.2 o .2 ._ .6 .s 1.0 -<2
x/cStation S, y/(b/2) = 0.62
oq_o' )9_o
0 .2 .q .6 .8 1 .O
x/cStation 8, y/{b/2) = 0.95
-IS
-I0
Cp -s
[
_5.2 0 .2 •q .6 .8
x/cStation 3, y/(b/2) = 0.45
-iS
-10
Cp -s
0
1.o _.2
9) I_®oo{)oo o o _] _,-
0 .2 .q .6 .8
x/cStation 6, y/(b/2) = 0.75
(d) a=10.0 °.
Figure 9. Continued.
1.0
0 Main wing
O Flap
Aileron
A Spoiler
+ Lower surface
28
%
-IS
-I0
-5
0
_s2
D
®
0 .2
(t) _ [ _0 _ _
.4 .6 .8 i .0
x/cStation 1, y/(b/2) = 0.20
Cp
15
tO
S
0
_5.2
D
{_bO( }O00 ( 0[_
0 .2 .q .6 .8
x/cStation _, y/(b/2) = 0.55
.0
%
-5
-ii _D ....
%0 }00 0
0
_5.2 0 ,2 .4
x/cStation 7, y/(b/2) = 0,85
•6 .8 1.0
15
-I0
cp -5
5.2 0 .2 .4 .6 .8 1.0
x/cStation 2, ,//(I}/2) = 0.35
-15 ___ ,
-10 --- ,
%0{)_ _ !
_]
- .2 0 .2 .4 .6 .8 .0
x/cStation 5, y/{b/2) = 0.62
15
-I0
Cp -s
=.2 0 .2 .4 .6 .8 1.0
x/cStation 8, y/(b/2) =0.95
c1,
-IS
-5 D
)OO O
[ _®( )®® e)
S,2 O .2 .q
x/cStation 3, ¥/_/2} =0.45
r [
.6 .8 1.0
%
-10 - '
-S ----
-.2 o .2 ., .6 .8 1.ox/c
Station 6, y/(b/2) = 0.75
O Main wing
o Flap
0 Aileron
A Spoiler
+ Lower surface
(e) a=12.0 ° .
Figure 9. Continued.
29
%
-15I
-I0
-5
t %0 ( 0 0
o , ,.ox/c
Station I, yl(bl2) = 020
-15
-i0
cp -s
0
_S.2
0
%0( )OO O oI_E( ] i_ lq.
0 .2 .q .6 .8
x/¢Station q, ¥/(b/2) = 0.55
1.0
%
-15
-I0
-S --
0
5-.2
i
D
_°°' _oo C o
I I.2 .q .6 -8 1.0
x/cStation 7, y/(b/21 = 0.85
Cp
-1s,___
-10 --
-s ,
° f
5-.2 0
___.
t
po_oq o_,.e_ {.t
•2 .q .6
x/c• 8 1.0
Station 2. y/(b/2) = 0.35
%
-15
-I0
?.2
)
!%O oo(il_)®l '®®
!
t_
0 .2 .q .6 .8
x/cStation S, y/(b/2) = 0.62
1.o
-15
-I0
cp -s
=.2
o_o,:)go _ R
!0 .2 .q .B .8 1.0
x/cStation 8, y/(b/2) = 0.95
-15I
-to
% -_---_)oo o
0 9_),:E)( )®®
_S.2 0 .2 .q
x/c
o _,_
.6 .8 1.0
Station 3, y/(b/2) = 0.45
-15
-10
Cp-5
0
_S.2
U
+D°°( )00 0
)(9%} ®oE IE_n
0 .2 .4 .6 .8
x/cStation 6, y/(b/2) = 0.75
1.0
0 Main wing
[3 Flap
<_ Aileron
A Spoiler
+ Lower surface
(f) _=14.2 ° .
Figure 9. Continued.
30
% Cp
15
5-., 0 .2 .q .6 .8 I .0
x/cStation q, y/(b/2) = 0,55
%
-15
-10
5
,ooq_ 04
0 ,2 .q .6 .8
x/cStation _ ¥/(b/2) = O,8S
---- [
1.0
-15 ___
10--
CP -S-
.2 o .2 .q .6 ,a i.ox/c
Station 2, ,//(b/2) = 0.35
%
1sI
5 - - ' 1..... ]
-2 .q .6 .8 t ,0
x/cStation 5. ¥/(b/2) = 0,62
Cp
S. 2 • 2 .q .6 ,B _ -0
x/cStation 8, y/(b/2) =0.95
%
-1S
-10
-5
1.... I
_oo $ o + o_
• 2 0 .2 .q .6 ,B i .0
x/cStation 3, y/(b/2) = 0.45
%
-5[ --_0
5!?
x/cStation 6, ¥/(b/2) = 0,75
0 Main wing
Flap
0 Aileron
Spoiler
+ Lower surface
(g) c_ = 16.2 °.
Figure 9. Concluded.
31
am
.2
.1
A
--m_y
m
rJ_
Reynolds number
O 3.05 x 106
[] 3.67
4.26
-5 0 5 10 15 2O 25 SO .2
ex, deg
-.4
-I0 .I 0 -.I
C_
-.6
(a) Lift and pitching moment.
Figure 10. Effect of Reynolds number with flap undeflected. 6f = 0°; VG off.
32
CL
L6
L4
L2
L0
.8
.6
.4
.2
0
..¢)
°.4
-._
".8
.01
j('_'
LP,I
III
.0'2 .OS
Reynolds number
O 3.05 x 106
[] 3.67
4.26
i
,'x
.04 .06
v
.06 .09 .08 .og
C D
•10 .11 .12 .IS .14 .11_
(b) Drag polar.
Figure 10. Concluded.
33
-.l
Cm
-.2
",4
-.6 ....
,_/j/-
i1_ -_"
Reynolds number
O 3.05 x 106
[] 3.67
4.26
2.6
2.4
2.2
2.O
1.8
1.6
CL 1.4
L2
1.0
.8 .,-
.6
.4
¢1
]l
//V
r-
J
¢
iX\\
_9
f_
4_//FI A
J
\
FJ
.2-10 -5 0 5 10 11_ 20 26 80 -.1 -.2 -.S
a, deg C_.
-.4 -.1_ -.6
Figure 11. Effect of Reynolds number with flap deflected, t_f = 40°; VG on.
34
Cnl
.2
.1i f
fvv
Transition
O r--re _
[] Fixed at x/c = 005.
1_6
L4
L2 ---
1.0
.8
.6
.2
0
-.2 D _
-.4
1-.6 /
..8;3-10
/
J/-
!
/
J
F "r
)_.,___A:_"/
-5 0 5 I0 15 20 25 _ .2 .1 0
a, deg C,.
(a) Lift and pitching moment.
Figure 12. Effect of fixed transition with flap undeflecled, bf = 0°; VG off.
tl
-.1
35
1.4
1.2
1.0
.8
.6
.4
CL.2
-.2
-.4
-.6
P/
#
//f
,j
\\
Transition
0 Free
[] Fixed at x/c = 0.05
.02 .03 .04 .05 .06
CD
(b) Drag polar.
Figure 12. Concluded.
/
.07 .08 .09 .10 .11
36
-.1
Cm
-.2
-.S
-.4
-.5
°°6
/f7
Transition
0 Free
[] Fixed at ×/c = OOS
2.6
2.4
2.2
2.0
L6
1
L6 "-_'--
C L L4
L2
LO
' /.6 p,
.4
¢/
/
-0
.2-15 -lO -5 0 5 10 15 20 25 SO -.1 -.2 -.S -.4 -.5
a, deg Cm
-.6
Figure 13. Effect of transition with flap deflected. _f = 40°; VG on,
37
Sensorlocation,percentchord
Laminar burst
60 _ 30
50 _ 50
40 _ 70
Turbulent burst
Percent of timeboundary layer
is laminar
.... 10010 ..... : .........
I I I I1.50 .5 1.0
Time, sec
Figure 14. Sample trace of hot-film data. e L = 0.22; 6f = 0°; VG off.
38
100
Extent of
laminar flow, 50percent
chord
0--.5
© Outboard region
[] Mid-semispan region
0 Inboard region
J 26
, I i "_"-"_ i0 .5
C L
11.0
(a) Upper surface.
Extent oflaminar flow,
percentchord
100
5O
0
© Outboard region
[] Mid-semispan region
<_ Inboard region
I
.26
-50 , I , 1 , 1-.5 0 .5 1.0
C L
(b) Lower surface.
Figure 15, Extent of laminar flow oil wing with flap undeflected. 6f = 0°; VG off.
39
-.4
-.6
,OO".... ,'- _ _-t_-O._ .l_ I
[ ......... --,_ ] _ _,_ q r _ _ _-/r
0
[]
<>
8f, deg
0
2O
4O
2.8
Cir.
2,4
2.0
L6
L2
.8/
0 /,(
-.i#
-.8 _"
-_ -]J)
7
v
fn 1
9./XJ
-5
/,.,
¢
(a) Lift and pitching moment.
Figure 16. Effect of flap deflection. VG on.
_y
(
)
)
)
))
0
<
ql/V
7
]
]
1
-.2 -.4
Clll
?
-.6
4O
2.8
2.4
2.O
L6
1.2
.4i
0 t
C L
-.4
-.810
J,I
Jr.
! ] ,
)r
IP
,%Y
i
.06 AO
8r, deg
0 0
[] 20
40
.111
i
.B--_ s!J
i
i
i
iI
I
1!
I
lCD
f-
(b) Drag polar.
Figure 16. Concluded.
"1
.S0 ._I
.40 .45
41
ORIGINAL PAGE
BLACK AND WHITE PHOTOGRAPH
42
Cm
.2
.I
0
#F
Vortex generator
0 o_r
[] on
(]I.
].6
1.4
12
LO
.8
.6
.4 -,
.2--
-.2
-.4 -- -
-.6 J
/..s I_
-10
i
[
//
,¢[
\= ,-,
/ri " "
!
9_
_ /
/mu
(B
-5 0 5 IO 15 2O _ SO .2 .1
a, deg C,.
(a) Lift and pitching moment.
Figure 18. Effect of vort_x generator with flap undeflected. _f = 0°.
Ii
[
\
".1
43
vortex generator
0 off
[] on
16
L4
12
1.0
.8
.6
.2
0
-.2
-.4
-.6
-.8.01
J
#
J
f
g-
..,; .-,-,
w.r-. i
.02 .08 .04 .06 .06 .0"/ .08 .0g .10 .U .12 .13 .14
G D
(b) Drag polar.
Figure 18. Concluded.
44
..1
Cm
-.2
-.S
".6
_m
/JIB
Vortex generator
0 Off
[] On
2.4
2.2
L8
1.6--
C L L4-
L2
LO
.8
.6i
.4
//!
//
¢V
//
\&
L
\b/
/r/
,-2
C_
f,
J,
.2-15 -lo .6 o 5 lo m 2o 25 so -.1 ..2
c_, deg
Figure 19. Effect of vortex generator with flap deflected. _/= 40 °.
-.4 -.5 -.6
45
Cm
.l ._)
o _: : '_'__ _ j '_'___
0
[]
Droop
Off
On
C L
L6
L4
L2
LO
.8
.4
.2
-.2
-.4
9°°8
-10
/
f
//
l
JJ
d/-
d/
i" \_I,,,.,.I
_b,...
-5 0 5 10 15 20 25 30 .1 0
_x, deg C m
[b
-.1
(a) Lift and pitching moment.
Figure 20. Effect of leading-edge droop with flap undeflected. 6/= 0°; VG on.
46
L6
L4
L2
LO
.8
.6
C L .4
.2
-.2
-.4
-.6
-.8.01 .02 .OS .04 .06
Droop
0 Off
[] On
I
v
.06 .07 .(}8 .09 .I0 .11 .12 .IS .14 .15
CI)
(b) Drag polar.
Figure 20. Concluded.
47
.1
0
ACI -. 1
-.2
5a, 5s,deg deg
O 20 0
[] -30 0
-30 60
a,.,
3,..
-.3-10 -5 0 5 10 15 20 25 30
a, deg
Figure 21. Roll-control effectiveness with flap undeflected. 5f = 0°; VG on.
ACI
.1
0
-.1
-.2
-.3
-.4
-.5
-15
ag _s,degO 20 0
[] -30 0
-30 60
?
f
-10 -5 0 5 10 15 20 25 30 35 40co, deg
Figure 22. Roll-control effectiveness with flap undeflected. 6f = 40°; VG on.
Droop
0 Off
[] On
.10 ,_ k
-10 -5 0 5 10 15 20 25 _0
a, deg
Figure 23. Effect of leading-edge droop on roll control with flap undeflected. 6f = 0°; 6a = 20°; VG on.
48
Report Documentation PageNatror_a_Aeronaut,cs and
Space Admir',stralion
1. ReportNAsANO.TP_3133 2. (;overnnlent Accession No. 3. Recipient's Catalog No.
4. Title and Subtitle
Full-Scale Semispan Tests of a Business-Jet Wing With a NaturalLaminar Flow Airfoil
7. Author(s)
David E. Hahne antt Frank L. Jordan, ,Jr.
9. Perfl)rming Organization Name and Address
NASA Langley Research Center
Hampton, VA 23665-5225
12. Sponsoring Agency Name aim Address
National Aeronautics and Space Administration
Washington, DC 20546-0001
5. Report [)ate
September 1991
6. Perfl)rming Organization Code
8. Performing Organization Report No.
L-16905
10. "Work Unit No.
505-61-41-01
11. Contract or Grant No.
13. Type of Report and Period Covered
Technical Paper
14. Sponsoring Agency (?ode
15. Supplementary Notes
16. Abstract
An investigation was conducted in the Langley 30- by 60-Foot Tlmnel on a full-scale semispanmodel to evaluate and document the low-speed, high-lift characteristics of a business-jet (:lass
wing that utilized the HSNLF(1)-0213 airfoil section and a single-sh)tted flap system. In addition
to the high-lift studies, boundary-layer transition effects were examined, a segmented leading-
edge droop for improve(t stall/spin resistance was studied, and two roll-control devices were
evahmted. The wind-tunnel investigation showed that deployment of a single-slotted, trailing-
edge flap was effective in providing substantial increments in lift. required for takeoff and landing
performance. Fixe(l-transition studies to investigate premature tripping of the boundary layerindicated no adverse effects on lift and pitehing-moinent characteristics for either the cruise or
landing configuration. The fifll-scale results also suggested the need to further ot)t.imize the
lea(ling-edge droop design that was (teveloped in the sut)scale t(>ts.
17. Key Words (Suggested by Autilor(s))
Natural laminar flow
High lift
Boundary-layer transitionStall/spin resistance
18. Distribution Statement
Unclassified Unlimited
Subject Category 02
19, Security Classif. (of this report) 20. Security Cla.ssif. (of this page) 21. No. of Pages 22. Price
Unclassified Unclassified 49 A03
NASA FORM 1626 OCT _6 NASA-l,angley. 19!11
For sale by the National 'I?_ehnical Information Service, Springfield, Virginia 22161-2171