Full-Scale Semispan Tests of a Business-Jet Wing With a

_*/,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

References

Somers, Dan M.: Design and Experimental Results for aNatural-Laminar-Flow Airfoil for General Aviation Appli-cations. NASA TP-1861, 1981.

2, McGhee, Robert ,].; Viken, Jeffrey K.; Pfimninger,Werner; Beasley, William D.; and Harvey, William D.:Experimental Results for a Flapped Natural-Laminar-FlowAirfoil With High Lift/Drag Ratio. NASA TM-85788,1984.

3. Morgan, Harry L.: High-Lift Flaps for Natural LaminarFlow Airfoils. Laminar Flow Aircraft Certification, LouisJ. Williams, compiler, NASA CP-2413, 1986, pp. 31 65.

4. Harvey, William D.: Boundary-Layer Control for DragReduction. International Pacific Air _4 Space Technolo9yConference Proceedings, P-208, Soe. of Automotive Engi-neers, Inc., 1988, pp. 287 311. (Available as SAE Paper872434.)

5. Harvey, W. D.; McGhee, R. J.; and Harris, C. D.: WindTunnel Testing of Low-Drag Airfoils. Laminar Flow Air-craft Certifcation, Louis J. Williams, compiler, NASACP-2413, 1986, pp. 89 128.

6. Sewall, William G.; McGhee, Robert J.; Viken, JeffreyK.; Waggoner, Edgar G.; Walker, Betty S.; and Millard,

Betty F.: Wind Tunnel Results for a High-Speed, Natu-

ral Laminar-Flow Airfoil Designed for General A_,iation

Aircraft. NASA TM-87602, 1985.

7. Sewall, William G.; McGhee, Robert J.; Hahne, David

E.; and Jordan, Prank L., Jr.: Wind Tunnel Results of

the High-Speed NLF(1)-0213 Airfoil. Research in Natural

Laminar Flow and Laminar-Flow Control, Jerry N. Hefner

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 °.


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 ° .


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 ° .


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


• 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 °.


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 °.


27

15

Cp -s

°[_s. 2

I

!i

0

[(_bo{) 0 0 , C_ _q

T-T0 .2 .q .6 .8 1.0


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


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 °.


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


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 ° .


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


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 ° .


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


%

-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 °.


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.


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,.


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.


/

.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

/,.,

¢


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.


"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,.


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.


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


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.


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

Documents

Full-Scale Semispan Tests of a Business-Jet Wing With a