NASATechnicalPaper2011

1982

I IASANational Aeronauticsand Space Adminislration

Scienlific and TechnicalInformation Branch

Effects of Wing-Leading-EdgeModifications on a

Full-Scale, Low-WingGeneral Aviation Airplane

Wind-Tunnel Investigationof High-A ngle-of-A trackA erod_narnic Characteristics

William A. Newsom, Jr.,Dale R. Satran, andJoseph L. Johnson, Jr.Langley Research CenterHampton, Virginia

SUMMARY

A static-force-test investigation has been made on a full-scale, low-wing gen-

eral aviation airplane in the Langley 30- by 60-Foot Tunnel to determine the effects

of wing-leading-edge modifications on the high-angle-of-attack aerodynamic character-

istics. The leading-edge modifications included leading-edge droop and slat config-

urations having full-span, partial-span, or segmented arrangements. Other devices

included wing-chord extensions, fences, and leading-edge stall strips. Some tests

were made to determine control effectiveness and the effects of power.

The investigation showed that good correlation exists between the results of

wind-tunnel data and the results of flight tests, on the basis of autorotational sta-

bility criterion, for a wide range of wing-leading-edge modifications. It was found

that the addition of a drooped leading edge on the outboard wing panel delayed tip

stall to a very high angle of attack and resulted in a relatively small drag penalty

in cruise. Segmented leading-edge droop or slats were found to be equally effective,

but the drag penalties were much higher. Wing-chord extensions, fences, or leading-

edge stall strips were generally ineffective. The outboard leading-edge-droop modi-

fication, which was most promising from the standpoint of stall departure and spin

resistance, had little effect on static longitudinal stability, increased lateral

stability, and generally provided an increase in lateral control at high angles of

attack. Full-span leading-edge modifications tended to degrade airplane stall-

departure and spin-resistance characteristics.

INTRODUCTION

The NASA Langley Research Center is currently conducting a broad research pro-

gram to develop the technology required to provide improved stall departure and spin

resistance of light general aviation airplanes. The program was initiated because

stalling and spinning have been identified as major causes of fatal general aviation

accidents (refs. I and 2). The research encompasses a wide variety of test tech-

niques involving wind-tunnel tests, radio-controlled-model tests, and full-scale

flight tests. Presented in references 3 to 10 are some results obtained in the

research effort thus far. Included in this program are studies to define concepts

which improve the stall characteristics and spin resistance of light general aviation

aircraft as well as studies of the fully developed spin and recovery. Given in ref-

erences 8 to 10 is a summary of the significant results obtained to date relative to

the effects of wing-leading-edge modifications on the stall/spin behavior of a

typical, light general aviation airplane.

The present research effort on wing modifications was inspired to a great extent

by recent research conducted at the University of Michigan and at NASA Ames Research

Center, in addition to earlier work conducted by the National Advisory Committee for

Aeronautics (NACA) to investigate the effects of wing-leading-edge modifications on

the lateral-directional characteristics of wings near stall (refs. 11 to 17). In the

studies of references 16 and 17, the concept of a segmented wing leading edge was

developed to control stall progression and to produce a "flat-top" wing lift curve to

minimize or eliminate loss of damping in roll at the stall. In the tests summarized

in references 8 to 10, an outboard wing-leading-edge modification was developed which

significantly improved lateral-stability characteristics at the stall, spin-

L-15101

resistance, and developed-spin characteristics as determined by a radio-controlled-model and by full-scale flight tests. Becauseof the need for full-scale-Reynolds-numberaerodynamic data for analysis of airplane flight-test results, an investiga-tion has been conducted in the Langley 30- by 60-Foot Tunnel using an airplanesimilar to the flight-test configuration. Someof the results of the wind-tunneltests are summarizedin reference 9. This paper includes that summary,augmentedwith data for additional configurations, pressure distributions, and associatedanalysis.

This investigation was directed at determining the effects of wing-leading-edgemodifications on the high-angle-of-attack aerodynamic characteristics of the subjectairplane configuration. Particular emphasis is placed on those configurations for

which flight-test results were obtained. The leading-edge modifications included

leading-edge-droop configurations and slat configurations having full-span, partial-

span, or segmented arrangements. Other devices tested included wing-chord

extensions, fences, and leading-edge stall strips. Most of the tests to investigate

leading-edge devices were made for the configuration with the horizontal tail

removed, but the effects of the most promising leading-edge device determined in the

flight tests of references 9 and 10 were documented for the complete airplane. Tests

of the complete airplane included rudder and elevator deflections and effects of

power.

In addition to measurements of the forces and moments of the airplane made on

the tunnel balance system, the forces and moments on the left outboard wing panel

were recorded independently by a strain-gauge balance, and the right wing of the air-

plane was provided with several rows of pressure ports to provide pressure measure-

ments for many of the tests. Flow surveys and flow-visualization studies utilizing a

tuft grid, smoke, and "mini-tufts" were also employed during the investigation. The

investigation was conducted at angles of attack ranging from -9 ° to 41 ° and at side-

slip angles ranging from -15 ° to 15 ° for a Reynolds number of about 2.5 × 106 , based

on the mean aerodynamic chord.

SYMBOLS

All longitudinal forces and moments are referred to the wind-axis system and all

lateral-directional forces and moments are referred to the body-axis system. Moment

data are presented with respect to a center-of-gravity position of 25 percent of the

wing mean aerodynamic chord at fuselage water line 34.87 (in.). Dimensional quanti-

ties are presented in U.S. Customary Units.

b wing span, ft

c local wing chord, ft

mean aerodynamic chord, ft

f0 <+)c n section normal-force coefficient, (Cp, 1 - Cp,u) d

Cp section wing-pressure coefficient, (p - p )/q_

C D airplane drag coefficient, Drag/q S

CD,wp

%

CL,wp

CI

%

C n

C R

I

CT

Cy

D

n

P

P_

q_

S

Swp

V

X

Y

(I

6e

6f

6r

N2D

1

outboard-wing-panel drag coefficient,

airplane lift coefficient, Lift/q S

outboard-wing-panel lift coefficient,

Drag(wing panel)/q Swp

Lift(wing panel)/q Swp

rolling-moment coefficient, positive right wing down, ' Rolling moment/q Sb

pitching-moment coefficient, positive nose up, Pitching moment/q Sc

yawing-moment coefficient, positive nose right, Yawing moment/q Sb

c C 2resultant-force coefficient, _ L,wp

+D,wp

effective propeller-thrust coefficient,

Drag(prop off) - Drag(plop running)

q S

side-force coefficient, Side force/q S

propeller diameter, ft

propeller speed, rps

local static pressure, Ib/ft 2

free-stream static pressure, Ib/ft 2

free-stream dynamic pressure, ib/ft 2

wing area, ft 2

outboard-wing-panel area, ft 2

velocity, ft/sec

chordwise distance from leading edge, ft

spanwise distance from plane of symmetry, ft

angle of attack, deg or rad

angle of sideslip, deg

elevator deflection, positive trailing edge down, deg

flap deflection, positive trailing edge down, deg

rudder deflection, positive for left yaw, deg

incremental drag coefficient relative to basic airplane configuration

incremental rolling-moment coefficient

ACn incremental yawing-momentcoefficient

incremental side-force coefficient

Stability derivatives:

_C _C1 nC - C -

Subscripts:

I lower

u upper

_CyC -Y8 bR

DESCRIPTION OF AIRPLANE

Basic Configuration

A three-view sketch of the low-wing general aviation airplane used in these

tests is presented in figure I, and a photograph of the airplane mounted in the

Langley 30- by 60-Foot Tunnel is shown in figure 2. This configuration differed

externally from the flight-test airplane of reference 8 only in that the flight-test

airplane was fitted with a tail-mounted spin-recovery parachute, streamlined wheel

fairings, and wing-tip-mounted _/8 sensor booms. Presented in tables I and II(a)

are the geometry characteristics of the airplane tested in the wind tunnel and the

coordinates of the basic-wing airfoil section, respectively.

The location of the balance used to measure loads on the left outboard wing

panel during the wind-tunnel tests is also shown in figure I. For this installation,

the wing panel was separated from the airplane along the line shown and reattached

with all loads carried through an internal straln-gauge balance. The opening at the

separation line was sealed with a thin rubber membrane. The right wing of the air-

plane was provided with static-pressure ports to provide pressure measurements for

most of the tests. Figure 3 is a drawing of the right wing panel showing the span-

wise locations of the 6 rows of pressure ports. Each row consisted of 15 chordwise

ports on the upper surface and 8 on the lower surface. When the leading-edge-droop

modifications were added, they were provided with ports in their upper and lower

surfaces to replace the wing ports covered.

Wing-Leading-Edge Modifications

Leading-edge droop.- Most of the tests using a drooped leading edge were made

with the leading-edge airfoil configuration developed in reference 8 as a device

which would improve lateral stability at the stall. The modification to the basic

wing consisted of a glove installed over the forward part of the airfoil which

provided a 3-percent-chord extension and a droop which increased the leading-edge

camber and radius as shown in figure 4. Coordinates of the new airfoil section

created by this basic-droop piece are presented in table II(b). Several different

configurations using various spanwise segments of the basic leading-edge droop were

tested in flight and in the wind tunnel as shownin figure 5. These modificationswere created by changing the spanwise location of the abrupt discontinuity at theinboard end of the droop piece (figs. 5(a), (b), (d), and (e)), creating a gapin what would otherwise be a full-span droop (figs. 5(f) and (g)), or fairing theinboard edge of the discontinuity (fig. 5(c)).

An exaggerated leading-edge-droop configuration was created for the wind-tunneltests by using the basic leading-edge droop as a starting point. Shownin figure 6is the configuration resulting from mounting an additional drooped portion onto theoutboard panels of the wing with the original drooped-leading-edge glove still inplace. Coordinates of the airfoil section resulting from this leading-edge configu-

ration are presented in table II(c).

A third leading-edge-droop configuration, shown in figure 7, was made by

superimposing the leading edge of a NASA LS(I)-0417 airfoil section onto the lead-

ing edge of the basic wing. This airfoil section was selected primarily because of

its large, rounded leading edge. The chord line of the LS(I)-0417 was tilted down-

ward 1.5 ° to accomplish the upper-surface alignment. The resultant airfoil section

was faired back into the lower surface of the basic wing starting about 0.10c behind

the leading edge. Coordinates for the LS(I)-0417 drooped section are given intable II(d).

Upper-surface modification.- A modification to the upper surface of the basic

wing was designed in reference 18 to improve maximum lift of the NACA 64-series

airfoil used for the basic wing of the test airplane. This increase in thickness of

the airfoil, as shown in figure 8, extended over the forward 42 percent of the wing

chord. Coordinates of the new airfoil created by this modification are presented in

table II(e).

Leadin_-ed@e slats.- Two leading-edge-slat configurations were used in the

tests. Sketches of the slat arrangements are presented as figures 9 and 10, and it

can be seen that the slats, which differed only in chord width, were tested in both

partial- and full-span configurations.

Leadin_-ed@e stall strips.- Two sets of leading-edge stall strips were used in

the tests. The stall strips were made to mount on the basic wing and on the original

drooped leading edge at the same spanwise location as the gap of figure 5(g). As

shown in the sketches of figure 11, the stall strips were triangular in cross section

and were made in three chord widths.

Win@ fences.- Sketches of wing-fence arrangements used in the tests are shown in

figures 12(a) and (b). The fences were located at 57-percent semispan and were made

in 2 chord lengths. The long fence was full chord and the short fence, which was

tested only in skewed-in and skewed-out configurations, had a chord length of 23 per-cent of the wing chord.

Chord extension.- The leading-edge extension to the chord of the wing is shown

in figure 12(c). A simple glove made to fit over the wing leading edge was used to

increase the wing chord by 8 percent. The leading edge of the glove was made using

the same coordinates as those of the basic wing.

Fillet droop.- In order to carry the lines of the basic leading-edge-droop

airfoil into the side of the fuselage, a tapered fairing was made which mounted on

the wing fillet. This fillet fairing is shown in the sketch of figure 12(d).

TESTS

The investigation was conducted in the Langley 30- by 60-Foot Tunnel. Except

for the loads on the left outboard wing panel being measured by the internal bal-

ance mentioned previously, all forces and moments were measured on the tunnel scale

system. The static pressures on the right wing were recorded by using a set of

scanivalve units. Forces and moments presented are the average of 10 sets of data

recorded at 1-sec intervals for each test condition.

During the tests of the complete airplane configurations, measurements were

included with both ailerons deflected ±25 ° and the elevator deflected 12 ° to -23 ° .

Most of the tests were made with power off, but for some tests of the complete

airplane, power was set to produce an advance ratio (V/nD) of 0.5 (C_ = 0.11).Reynolds number for the tests was about 2.5 × 106 based on the mean aerodynamic

chord (_) for a free-stream dynamic pressure q_ of about 11 ib/ft 2. The test angle

of attack, which was set using an accelerometer mounted in the model, ranged from

-9 ° to 41 ° and the sideslip angle ranged from -15 ° to 15 ° .

In addition to the measurement of forces and moments on the airplane during the

investigation, flow surveys and flow-visualization studies were made. These studies

included the use of a tuft grid, smoke, and "mini-tufts" which were illuminated by

ultraviolet light.

The longitudinal data from the tests have been corrected for blockage, airstream

misalignment, buoyancy effects, mounting strut tares (including propeller slipstream

effects), and wind-tunnel jet-boundary effects on the wing and the tail. Effects of

the propeller slipstream at the tail are also accounted for in the tail-on jet-

boundary corrections. Lift and drag corrections have been made for the integrated

average airstream misalignment, and lateral-directional data are referenced to side-

slip angles which include a correction for the integrated average lateral-airstream

angle. An indication of the correction involved and a plot of the actual flow dis-

tribution in the tunnel is presented in appendix A of reference 19.

PREVIOUS FLIGHT RE_SULTS

In order to provide the reader with additional background information to aid in

the interpretation of this paper, a brief discussion of the flight tests (refs. 8

and 10) conducted with several of the wing-leading-edge modifications is presented.

Shown in figure 13 are sketches of the eight principal wing configurations previously

studied in flight. Shown under each configuration are summary comments describing

the spin results obtained in flight tests.

The airplane with the basic wing had two spin modes: one moderately flat and

the other flat. The moderately flat spin mode was characterized by an angle of

attack of about 50 ° , and recovery from this mode occurred I I/2 turns after applying

normal recovery controls. The flat spin mode was characterized by an angle of attack

of about 70 ° . Airplane controls were found to be ineffective for recovery, and the

use of a spin-recovery parachute was required. During the flight program, the air-

plane exhibited a strong tendency to enter the moderately flat spin mode, but it was

reluctant to enter the flat spin mode with normal prospin controls.

The full-span-droop configuration (modification A) was found to readily enter

a flat spin, regardless of the prospin controls employed. The flat spin mode was

characterized by an angle of attack of 60° to 70 ° , which was comparable with the flat

spin mode of the basic configuration. The airplane controls were ineffective for

providing acceptable spin recovery, and the spin-recovery parachute was required.

The airplane with the outboard wing-leading-edge droop (modification B) exhib-

ited a steep, slow spiral-type motion following prospin control inputs. Immediate

recovery was achieved (I/8 turn) by simply relaxing either prospin rudder or

elevator.

Addition of the fairing to the outboard leading-edge droop (modification C)

caused the spin characteristics to be severely degraded. The spin entry appeared

identical to that for the basic, outboard leading-edge-droop configuration, but after

I I/2 turns the rotation rate increased rapidly, the angle of attack increased, and

the airplane entered a flat spin mode. The spin mode was characterized by an angle

of attack of about 74 ° . Recovery controls were ineffective and the spin-recovery

parachute was deployed for recovery.

Spin characteristics relative to those obtained for the basic airplane were

degraded by shortening the outboard leading-edge droop to modification D. This con-

figuration also entered the flat spin easily. When the outboard leading-edge droop

was lengthened to modification E, no change was noted from results obtained with mod-

ification B (that is, the very steep, easily recoverable spin was obtained).

Finally, it was found that both of the segmented leading-edge-droop modifica-

tions (F and G) resulted in flat spins.

The foregoing flight-test results indicate extremely large effects of wing-

leading-edge modifications on spinning, and these results imply large variations of

aerodynamic autorotational tendencies for the various wing configurations. These

trends were quite evident in examination of the wind-tunnel data, as will be dis-

cussed. Also, the detailed flow phenomena and pressures responsible for the trends

were identified.

WIND-TUNNEL RESULTS

The results of this wind-tunnel investigation are presented in the figures

listed in table III.

Lift Characteristics

Basic wing.- Presented in figure 14 is the variation of lift coefficient with

angle of attack measured on the outboard wing panel during one test run of the air-

plane with the tail removed. The data show a normal trend and repeatability at the

lower angles of attack. Above an angle of attack of about 12 ° , however, the data

show a scatter band of lift values corresponding to random fluctuations in the wing-

balance readings. The lift fluctuations were periodic and did not appear to be

related to hysteresis effects; but rather, they appeared to be caused by random flow

separation and flow reattachment on the outboard wing panel.

The results of tuft studies for the basic configuration (presented in fig. 15)

illustrate the random-flow-separation problem on the outboard wing panel and provide

basic flow information for correlation with the lift data of figure 14. Two photo-

graphs of the flow patterns are presented for different time intervals corresponding

to the high- and low-lift readings at _ = 20 ° • The flow pattern for the high-lift

condition shows separated flow inboard on the wing but attached flow on the outboard

wing panel. For the low-lift condition, the flow on the outboard wing panel is shown

to be separated. The results of oil-flow studies (fig. 16) obtained on a I/3-scale

model in the University of Maryland!s Glenn L. Martin Wind Tunnel illustrate even

more clearly the fluctuations in surface flow conditions for the basic configuration

near the stall angle of attack. It is interesting to note the similarity in flow

patterns for the model and the aircraft. The two photographs of the model at

= 14 ° are presented to illustrate the random flow changes in flow patterns which

occurred at different time intervals - a result similar to that for the full-scale

aircraft at _ = 20 ° . In one photograph, the oil-flow studies show attached flow

on the outboard panel of the right wing and stalled flow on the left wing panel;

whereas, the other photograph, taken at a different time interval, shows the oppo-

site trends with flow attachment on the left outboard wing panel and separated flow

on the right wing panel. For angles of attack greater than about 25 ° , the flow-

visualization tests indicated that the entire wing of the basic configuration was

stalled. In view of the conditions described, it is necessary that some of the dataherein be used with caution.

Modified win_.- As mentioned previously, photographs were made of the flow

across the wing surface for each configuration tested. An example of the results

obtained is presented in figure 17. Shown are the stall patterns at _ = 30 ° and 35 °

for the wing with the addition of a drooped leading edge on the outboard portion of

the wing (modification B). It can be seen that this leading-edge-droop configuration

tended to have attached wing-tip flow to very high angles of attack. Closer examina-

tion of the flow associated with the outboard-droop configuration, using a tuft grid

and a smoke generator, indicated that the effectiveness of this configuration in

maintaining attached flow at the wing tips was the result of a vortex flow generated

at the inboard edge of the droop. The vortex flow apparently acted as an aerodynamic

fence to stop the spanwise progression of the separated flow region toward the wing

tips such that the tips continued to generate lift to high angles of attack. The

outboard wing panel then appeared to have aerodynamic characteristics generally

similar to those of a low-aspect-ratio wing.

Comparison of Wing Modifications

Basic leading-edge droop.- The significance of maintaining attached flow on the

wing tips to high angles of attack is illustrated in figure 18 by plots of the lift

and drag coefficients measured on the wind-tunnel scale system and plots of the

resultant-force coefficient measured on the wing-tip balance. The wing-tip-balance

data are included because the wing-tip aerodynamics on unswept wings are believed to

be closely related to the damping or autorotational tendencies exhibited by the

wing. As pointed out in reference 8, previous research has indicated that autorota-

tion is encountered when the variation of the resultant-force coefficient of the wing

angle of attack becomes negative; that is, when 5CR/_ < 0. For the subject config-

urations, the variation in the slope of the resultant-force coefficient of the wing

tip with _ is expected to provide information for a good prediction of autorota-

tional tendencies. The data of figure 18 show that the tip of the basic wing stalled

abruptly at an angle of attack of 20 ° and the lift decreased rapidly at higher angles

of attack. The addition of an outboard leading-edge droop (modification B) is

shown to eliminate the abrupt stall of the wing tip and to maintain or increase lift

up to _ = 40 ° . The change in slope of the resultant-force-coefficient curve from

negative to positive values at the higher angles of attack is believed to be

important in indicating the elimination of autorotation and the improvement of spin

8

resistance. It is interesting to note that the addition of a fairing on the inboardend of the outboard droop (modification C) to eliminate leading-edge discontinuity,or the addition of a full-span leading-edge droop (modification A), reintroducedabrupt tip stall and caused the slope of the resultant-force-coefficient versuscurve to becomevery negative at high angles of attack. This is probably the resultof eliminating the vortex formerly generated by the discontinuity. In flight testsreported in reference 8, modification B was very spin resistant; whereas, the basic-wing configuration showeda flat spin mode. Modifications C and A also exhibited aflat spin modein the airplane flight tests. Correlation of the values of 5CR/_

for the four configurations of figure 18 with the airplane flight-test results from

reference 8 can be made on the basis of figure 19. The values of 5CR/_ plotted

against angle of attack in figure 19 predict autorotation for all configurations

except B. It is interesting to note that all leading-edge modifications extended the

angle of attack at which _CR/_ became zero, but apparently the elevator power was

great enough to drive the airplane to angles of attack where only modification B

could provide attached flow on the wing tips.

Static aerodynamic data for correlation with flight-test data on the effect of

spanwise variation of the length of the leading-edge droop are shown in figure 20.

The test configurations included full-span droop (modification A) and partial-span

droop with inboard discontinuity at semispan stations of 72 percent (modification D),

57 percent (modification B), and 38 percent (modification E). The data of figure 20

show trends similar to those previously discussed in figure 18, but bring out two

additional points. First, shortening the outboard-droop length by moving the

inboard end of the droop from 57 to 72 percent of the semispan eliminates almost all

the effectiveness of the outboard-droop arrangement for providing stall departure and

spin resistance. In fact, the data show droop-modification D to have aerodynamic

characteristics very similar to those of the basic wing. The second significant

point regarding figure 20 is that droop-modification E provided aerodynamic data gen-

erally similar to modification A, except modification E tended to delay to a higher

angle of attack the rapid destabilizing change in C R near _ = 30 ° which was noted

for the full-span droop.

A summary of the data of figure 20 is presented in figure 21 in terms of _CR/_

plotted against wing lateral stations in percent of wing semispan. Presented in fig-

ure 21 are the values of 5CR/5_ for angles of attack from 20 ° to 40 ° along with

results of flight tests (ref. 10) which define boundaries of inboard discontinuity of

leading-edge droop which effectively prevent entry of the airplane into the flat spin

mode. The data of figure 21 indicate fairly good qualitative agreement between the

flight data and wind-tunnel static data, based on the criterion that autorotation is

encountered when _CR/_ < 0. Modification B is seen to provide stabilizing ten-

dencies over the angle-of-attack range, but shortening the length of the droop is

seen to produce negative values of 5CR/5_ at semispan stations corresponding very

closely to that identified in flight tests (67 percent of b/2) for loss of effec-

tiveness of the drooped leading edge. In the flight tests (ref. 10), an inboard wing

station was identified for loss of effectiveness of the leading-edge droop in provid-

ing resistance to the flat spin mode (35 percent of b/2). The wind-tunnel data show

that large negative values of 5CR/5_ can be encountered at that point for angles of

attack above 30 ° . Apparently, angles of attack of 30 ° or above can be induced at the

wing tips by rotation under spin conditions.

Wing-fillet droop.- With regard to modifications A, E, and B, an interesting

point brought out in the tests was that adding a leading-edge-droop modification to

the wing/fuselage fillet (see fig. 12(d)) altered the aerodynamic characteristics of

these configurations considerably. The data for the fillet droop modification on

modification A is shownin figure 22(a). It can be seen that the addition of thedroop to the fillet eliminated the initial break in the total wing-lift curve. With-out the wing-fillet droop, the initial lift-curve break in modification A dataoccurred near 5 = 10 ° and a stall break occurred near 5 = 25 ° . Values of C R in

figure 22(a) indicate that the addition of droop to the fillet decreased the angle of

attack at which _CR/55 for the wing tip changed from positive to negative values,

which suggests that the fillet droop would introduce autorotative tendencies at lower

angles of attack than the basic full-span-droop configuration.

The effect of adding the wing/fuselage-fillet droop to modification E or to mod-

ification E with a spanwise, inboard droop extension with length equal to 0.095b/2

is shown in figure 22(b). The total lift data for all configurations show an ini-

tial break near 5 = 10 ° with a stall break near 5 = 30 ° . Values of CR in

figure 22(b) show that addition of the drooped fillet to modification E provided a

slight increase in angle of attack at which 5CR/_5 changed from positive to nega-

tive values, indicating a stabilizing effect on autorotation tendencies. However,

modification E with the inboard droop extension and the fillet droop appeared to be

less stable on the basis of the variation of C R with 5.

The addition of the fillet droop to the modification-B wing arrangement

(fig. 22(c)) was also found to increase the autorotational tendencies of the airplane

on the basis of the variation of C R with 5.

Segmented leading-ed@e droop.- Presented in figure 23 is a comparison of data

measured on the airplane with two configurations of segmented leading-edge droop.

The segmented configurations (figs. 5(f) and (g)) are geometrically similar to those

tested in reference 8. The lift data of figure 23 show trends similar to those

reported in reference 8 in that initial stall occurs around 5 = 10 ° and a secondary

stall occurs at a higher angle of attack, resulting in a double-peak lift curve.

Values of C R show that the wing-tip stall is delayed by the segmented leading edge

from 5 = 20 ° to 5 = 30 ° . Values of 5CR/55 for the segmented configurations are

compared with those of the basic wing and of modification B in figure 24. The

segmented leading edges were not as effective as modification B in providing positive

values of _CR/_5 to high angles of attack. The segmented leading edge with the

larger cutout (modification G) shows only small negative values of 5CR/_5 at5 = 40°; whereas, the smaller segmented cutout (modification F) produced large

unstable values at 5 = 40 ° •

LS(I)-0417 leading-edge droop.- As described previously, a modified leading-edge

droop was created on the wing by superimposing the lines of an LS(I)-0417 airfoil.

Tests of this configuration (see fig. 7) provided the results shown plotted and com-

pared with the basic wing in figure 25. The figure shows an initial break near

5 = 10 ° for all configurations. The outboard LS(I)-0417 droop provided some

protection in preventing autorotational tendencies by delaying the angle of attack at

which stall occurred from _ = 20 ° to 5 = 30 ° . The full-span LS(I)-0417 droop

provided some increase in maximum lift coefficient after the initial stall break.

The initial stall break apparently was associated with wing-root flow-separation

problems, and the use of a droop on the fillet in combination with the full-span

LS(I)-0417 droop probably would have provided an increase in maximum lift coefficient

at the initial stall break. The resultant-force coefficient, however, indicates that

the full-span LS(I)-0417 droop provided no improvement in tip stall characteristics

compared with those of the basic wing.

Exaggerated leading-edge droop.- To determine the aerodynamic effectiveness of

increasing the radius of the leading-edge droop, tests were conducted using modifica-

I0

tions A and B with an extra droop section mounted outboard, as shown in figure 6.

The results presented in figure 26 show generally similar trends of lift, drag, and

wing-tip resultant force as those shown for the basic droop configurations (fig. 18),

which indicates that increasing the leading-edge radius provided little or no benefit

on the stall departure and spin resistance relative to those of the basic droop con-

figurations. Detailed studies of the penalties introduced on performance character-

istics of the airplane by modifying the leading-edge radius will be discussed in a

subsequent section.

Leadin@-ed_e slats.- The results of the tests of various slat configurations are

compared in figure 27. As shown by the data in figure 27(a), outboard-slat arrange-

ments provided trends in lift and drag generally similar to those of modification B

(basic leading-edge droop), but the outboard slat provided more lift on the outboard

wing panel. The full-span slats gave large increases in maximum lift coefficient and

delayed tip stall to very high angles of attack as seen in the data of figure 27(b).

The outboard-wing-panel data of figure 27 are summarized in figure 28 in terms of

_CR/SU plotted against u. Figure 28 is a very clear illustration of the similarityin autorotational stability of the outboard-droop and outboard-slat configurations.

The full-span slat arrangements are shown to provide autorotational stability except

near an angle of attack of 35 ° for the small-chord slat.

Chord extension.- The aerodynamic data obtained in tests of the extended-chord

configuration are shown in figure 29 compared with the basic wing data. It can be

seen that the extended chord on the outboard wing section had little effect on the

aerodynamic characteristics of the airplane. The results of figure 29 suggests that

the effectiveness of the outboard-leading-edge-droop configuration is apparently

associated to a great extent with the droop shape as well as with the abrupt discon-

tinuity of the inboard end of the drooped section.

Win@ fences.- Figure 30 is a summary of the results of tests made with the vari-

ous fence arrangements on the wing. The figure shows that the long-chord fence

provided some delay in the angle of attack at which the lift-curve slope changes

from positive to negative values; but, based on values of C R for the wing tip, the

fences were not very effective in improving the stall characteristics or the autoro-

tational tendencies.

Leading-edge stall strips.- Presented in figure 31 are the data from the tests

of the stall strips on the leading edge of the basic wing and on modification A of

the basic leading-edge-droop arrangement. Results for only the smallest chord stall

strips are shown, but the results were the same for the other stall-strip chord

widths. In general, the stall strips provided little if any aerodynamic benefits in

terms of improved stall characteristics or improved autorotational tendencies.

Wing upper-surface modification.- Data for the tests involving the configuration

having a modified upper wing surface show in figure 32 that there was some improve-

ment in maximum lift coefficient at the stall break for the full-span arrangement,

but very little improvement in extending the angle of attack at which autorotational

tendencies begin.

Chordwise Pressure Coefficients

Presented in figures 33 to 36 are representative plots of the chordwise pressure

coefficients obtained in the test. Data are included for: the basic wing (fig. 33);

modification B (fig. 34), which showed improved aerodynamic characteristics; modifi-

11

cation C (fig. 35), wherein the improved aerodynamic characteristics of modifica-

tion B were lost; and modification A (fig. 36), which had increased lift but poorautorotational characteristics.

Basic win_.- The data of figure 33(a) show that the basic wing had peak nega-

tive pressures near the wing leading edge and relatively high loading along the

inboard and forward portion of the wing. Increasing the angle of attack to 21.6 °

(fig. 33(b)) reduced the loading inboard because of stalling and increased the peak

leading-edge loading near the wing tip. At angles of attack of 31.9 ° and 41.5 ° ,

the chordwise stations show mostly zero pressure gradients along the wing chord with

values of Cp about 0.6, indicating flow separation over the entire wing.

Modification B.- The pressure data for the best outboard-droop configuration

near _ = 11 ° (fig. 34(a)) show chordwise variations very similar to those of the

basic wing (fig. 33(a)). However, at _ = 21.6 ° (fig. 34(b)) the outboard-droop

configuration shows decreased values of wing-leading-edge peak pressures at the

spanwise stations along the droop portion of the wing (stations 0.63b/2 to

0.92b/2). As shown in figures 34(c) and (d), the peak pressures along the drooped

leading edge maintained relatively high values even to an angle of attack of 41.5 ° .

Modification C.- Comparison of the chordwise pressure data for modification C

(fig. 35) with the data for modification B (fig. 34), with its abrupt discontinuity

at the inboard edge, shows that similar chordwise pressure variations and peak values

of pressure coefficients were obtained for both configurations except near _ = 40 ° .

Comparison of the data of figures 34(d) and 35(d) indicates that the creation of mod-

ification C by adding the inboard fairing to modification B reduced the peak values

of the leading-edge pressure coefficients along the drooped portion of the leading

edge and resulted in separated flow behind the wing leading edge as indicated by the

low, constant values of pressure coefficient (Cp = 0.60) aft of the wing leadingedge.

Modification A.- Comparison of the pressure data for the full-span-droop con-

figuration (fig. 36) with those for modification B (fig. 34) shows, as expected,

increased loading inboard along the wing for the full-span-droop configuration.

Outboard, the pressure data are very similar for the two configurations except for

angles of attack near 40 ° . The pressure data for modification A near _ = 40 °

show reduced values of peak pressure coefficients near the leading edge and separated

flow behind the leading edge. These data very closely resemble the data for

modification C.

Spanwise Load Distribution

Values of section normal-force coefficient obtained from figures 33 to 36 were

integrated and plotted as a function of semispan location in figure 37. In addition,

similar plots are included for the segmented leading-edge-droop configurations

investigated (fig. 38) and for the slat configurations in figures 39 and 40.

The data of figure 37 show generally similar span-load distributions for all the

configurations at an angle of attack of 11.2 ° • Modification A, as expected, showed

much higher inboard loading than the other configurations. Increasing the angle of

attack to 21.6 ° resulted in modification A having the highest loading, and all con-

figurations show a general trend of increased loading at the outboard stations. At

= 31.7 ° , modifications B and C continued to show heavy loading near the wing tips;

12

whereas, modification A showedreductions in loading near the tips. At _ = 41.5° ,the data show trends in span loading that are generally in good agreementwith thestatic-force-test results of figure 18 in that modification B provided high loadingat the tips, whereas the other configurations showreduced loading over the wingspan, especially at the wing tips.

Span-load distribution data for the segmentedleading-edge droop (modifica-tions F and G), presented in figure 38, showchanges in span loading with increasesin angle of attack that are generally similar to those noted in figure 37 for theoutboard droop arrangements. The segmentedleading edge with the largest cutout (G)provided higher wing-tip loading than the configuration with the smaller cutout (F).These results are generally as expected, based on the static-force-test data offigure 23.

Span-load data for the small-chord slat (fig. 39) and for the large-chord slat(fig. 40) are in good agreement with the static-force-test data of figure 27. Thedata of figures 39 and 40 show that the slats increased the span loading outboard onthe wing as the angle of attack was increased and maintained the high outboard load-ing up to the highest test angle of attack.

Drag Characteristics

In order to provide drag coefficient data for use in determining the performance

penalties of the leading-edge devices under investigation, the drag coefficient data

presented earlier were replotted to an expanded scale and incremental drag values

were determined for the configurations which appeared most promising for improved

stall departure and spin resistance. The curves of figure 41 show incremental drag

values plotted against CL for various configurations. Incremental drag values in

the cruise range (C L = 0.4) of 0.002 are shown for modification B, whereas modifica-

tion C has _C D = 0.0054 and modification A has AC D = 0.007. Calculated perfor-

mance figures indicate that AC D of 0.002 would reduce the airplane cruise speed

about 2 mph and AC D of 0.007 would penalize the airplane cruise speed about 6 mph.

The data of figure 41 show that the addition of the droop on the fillet would reduce

the drag penalty of modification A in cruise to a value of _C D = 0.004. Modifica-

tion B is shown to produce no penalty on the climb performance of the airplane

(C L = O.75).

Incremental-drag-coefficient data for the segmented leading-edge-droop configu-

rations are presented in figure 42. They show that the segmented leading edge with

the _naller cutout had the lower drag penalty but that both segmented configurations

generally produced a higher drag penalty than modification B.

The incremental-drag-coefficient data for the slat configurations are presented

in figure 43. The data show, as expected, that all slat arrangements produced very

large drag penalties on the airplane. The large-chord full-span-slat configuration

showed values of AC D at C L = 0.4 that almost doubled the drag of the basic air-

plane in cruise.

Complete Airplane

The aerodynamic characteristics of the complete airplane are presented in

figures 44 to 56.

13

Longitudinal characteristics.- The effects of power and flap deflection on the

longitudinal characteristics are shown in figure 44. Power had about the same incre-

mental effect on maximum lift as deflecting the flaps 30 ° , but neither power nor

flap deflection had any significant effect on the pitching moment. The effects of

elevator deflection with power off and on are shown in figure 45. The elevator was

effective for producing incremental pitching moment over the whole angle-of-attack

range and elevator deflection range, and as would be expected, the presence of

increased slipstream velocity with power on gave increased elevator effectiveness.

In the data of figure 46 which show the effect of the horizontal tail on the longi-

tudinal characteristics, it can be seen that the airplane has good static longitudi-

nal stability with the tail on and is unstable with the tail off.

Presented in figures 47 to 49 are the longitudinal aerodynamic characteristics

of the complete airplane with wing-leading-edge droop pieces in place. The leading-

edge modifications had little effect on static longitudinal stability of the air-

plane, indicating that the downwash characteristics in the vicinity of the horizontal

tail were unaffected by the wing-leading-edge modifications.

Lateral characteristics.- Wide variations in effective dihedral and directional

stability occurred starting at about _ = 15 °. These variations are apparently

associated with random asymmetric wing stall. The effect of this asymmetric stall,

which was first mentioned in the discussion of figure 14, on the measured rolling

moment is illustrated in figure 50. The data plotted in this report are averages of

10 measurements made at each test condition. In addition, data are presented,

between _ = 12 ° and _ = 36 ° , for the maximum and minimum values of rolling moment

measured on the tunnel scale system to emphasize the unusually large variations of

the readings in that angle-of-attack range. These variations in rolling moment and

yawing moment could adversely affect the accuracy of and betweenCI 8 Cn 8

= 15 ° and _ = 30 ° because these derivatives were calculated using the average of

the tunnel scale readings. Therefore, the values of lift, rolling moment, and yaw-

ing moment at any given time during a test could, as shown in figures 14 and 50, be

varying drastically as flow is detached and reattached on the wing panels.

The static lateral-stability characteristics of the complete basic airplane are

shown in figure 51. With the vertical tail on, the airplane was directionally stable

with power on up to about _ = 25 ° . The level of directional stability was higher

with power on than with power off. It is interesting to note, however, that in the

range of angles of attack between 0° and about 12 ° the airplane had more positive

effective dihedral (-C ) with the vertical tail off than with the tail on. Abovet8

15 ° angle of attack, effective dihedral and directional stability were subject to

wide variations.

The static lateral-directional characteristics of the complete airplane with

wing-leading-edge droop pieces in place are presented in figures 52 to 54. A com-

parison of the basic configuration and modification B (figs. 51 and 52) shows that

the addition of the outboard leading-edge droop reduced the directional stability and

provided a large increase in the effective dihedral at angles of attack greater than

15 ° . The increase in dihedral effect was, as expected, based on longitudinal data

which showed that the outboard droop modification provided attached flow on the wing

outboard span up to angles of attack near 40 ° . Between _ = 0 ° and _ = 20 o,

lateral-directional stability data for modification A presented in figure 53 show a

slight increase in directional stability and an increase in effective dihedral rela-

14

tive to the basic configuration. When the fillet droop was added to modification A,

the data (fig. 54) show that, compared to the data of figure 53, the directional

stability was reduced over the whole high-angle-of-attack range, but there was an

increase in -C at angles of attack between 15 ° and 40 ° . Comparison of the data

of figures 52 and 54, however, show that the increase in -C_8 provided by modifi-

cation A with drooped fillet was not as large as that provided by modification B in

the high-angle-of-attack range.

Rudder effectiveness of the airplane is shown in figure 55 and the aileron

effectiveness is shown in figure 56. In general, the rudder provided substantial

increments of yawing moment and the aileron provided substantial increments of roll-

ing moment over most of the angle-of-attack range. Presented in figure 56 are values

of _Cy, _Cn, and AC I provided by maximum deflection of the ailerons for rightroll. The data show that modification B provided relatively small increases in

aileron effectiveness even though attached flow was maintained to angles of attack

near 40 ° .

Results of the pressure surveys indicated that most of the benefits of the out-

board droop in delaying flow separation of the wing tips were near the wing leading

edge. The ailerons were apparently exposed to regions of separated flow despite the

benefits of the leading-edge droop in maintaining attached flow forward of the

ailerons. The data of figure 56 do show, however, that modification A apparently

provided improved flow conditions over the ailerons between _ = 10 ° and _ = 30 °

as indicated by the increased aileron effectiveness in that region. At angles of

attack above 30 ° , modification-A data indicated a sharp decrease in aileron effec-

tiveness, apparently because of flow separation on the outboard wing panel.

SUMMARY OF RESULTS

The results of an investigation to determine the effects of wing-leading-edge

modifications on the aerodynamic characteristics of a full-scale low-wing general

aviation airplane may be summarized as follows:

I. Good correlation was obtained between the results of wind-tunnel static data

and the results of airplane flight tests, on the basis of the autorotational stabil-

ity criterion, for a wide range of wing-leading-edge modifications.

2. The addition of a drooped leading edge on the outboard wing panel delayed tip

stall to a very high angle of attack and resulted in a relatively small drag penalty

in cruise.

3. The effectiveness of the outboard-droop arrangement in delaying tip stall is

attributed to a vortex flow field at its inboard discontinuity which prevented sepa-

rated flow from progressing outboard on the wing. The outboard wing panel, with the

addition of the drooped leading edge, appeared to have aerodynamic characteristics

generally similar to those of a low-aspect-ratio wing with significant delay in wing-

tip stall.

4. The use of segmented leading-edge droop, slats, or exaggerated leading-edge

droop on the outboard wing panel was effective for delaying tip stall, but was

accompanied by increased drag penalties.

15

5. Leading-edge droop on the wing/fuselage fillet minimized flow separationproblems at the wing/fuselage juncture. The fillet droop eliminated the initiallift-curve break and reduced drag for most of the full-span leading-edgemodifications.

6. The outboard leading-edge-droop modification, which was most promising fromthe standpoint of stall departure and spin resistance, had little effect on staticlongitudinal stability, increased lateral stability, and generally provided someincrease in lateral control at high angles of attack.

7. Full-span leading-edge droop, wing/fillet droop, full-span slats, and seg-mented leading-edge droop (with small gap) degraded airplane stall-departure andspin-resistance characteristics based on the autorotational stability criterion.

8. The wing upper-surface modification, leading-edge stall strips, wlng-chordextension, wing fences, and LS(I)-0417 leading-edge droop provided little or noimprovement in airplane autorotational characteristics.

Langley Research Center

National Aeronautics and Space Administration

Hampton, VA 23665

May 11, 1982

16

REFERENCES

I. Silver, Brent W.: Statistical Analysis of General Aviation Stall Spin Accidents.[Preprint] 760480, Soc. Automot. Eng., Apr. 1976.

2. Ellis, David R.: A Study of Lightplane Stall Avoidance and Suppression.FAA-RD-77-25,Feb. 1977.

3. Bowman,JamesS., Jr.; Stough, Harry P.; Burk, Sanger M., Jr.; Patton, andJamesM., Jr.: Correlation of Model and Airplane Spin Characteristics fora Low-WingGeneral Aviation Research Airplane. AIAA Paper 78-1477, Aug. 1978.

4. Burk, Sanger M., Jr.; Bowman,JamesS., Jr.; and White, William L.: Spin-TunnelInvestigation of the Spinning Characteristics of Typical Single-Engine GeneralAviation Airplane Designs. I - Low-WingModel A: Effects of Tail Configura-tions. NASATP-1009, 1977.

5. Burk, Sanger M., Jr.; Bowman,JamesS., Jr.; and White, William L.: Spin-TunnelInvestigation of the Spinning Characteristics of Typical Single-Englne General

Aviation Airplane Designs. II - Low-Wing Model A: Tail Parachute Diameter and

Canopy Distance for Emergency Spin Recovery. NASA TP-1076, 1977.

6. Bihrle, William, Jr.; Barnhart, Billy; and Pantason, Paul: Static Aerodynamic

Characteristics of a Typical Single-Engine Low-Wing General Aviation Design for

an Angle-of-Attack Range of -8 ° to 90 ° • NASA CR-2971, 1978.

7. Bihrle, William, Jr.; Hultberg, Randy S.; and Mulcay, William: Rotary Balance

Data for a Typical Single-Engine Low-Wing General Aviation Design for an Angle-

of-Attack Range of 30 ° to 90 ° . NASA CR-2972, 1978.

8. Staff of Langley Research Center: Exploratory Study of the Effects of Wing-

Leading-Edge Modifications on the Stall/Spin Behavior of a Light General Avia-

tion Airplane. NASA TP-1589, 1979.

9. Johnson, Joseph L., Jr.; Newsom, William A., Jr.; and Satran, Dale R.: Full-

Scale Wind-Tunnel Investigation of the Effects of Wing Leading-Edge Modifica-

tions on the High Angle-of-Attack Aerodynamic Characteristics of a Low-Wing

General Aviation Airplane. AIAA-80-1844, Aug. 1980.

10. DiCarlo, Daniel J.; Stough, Harry P., III; and Patton, James M., Jr.: Effects of

Discontinuous Drooped Wing Leading-Edge Modification on the Spinning Character-

istics of a Low-Wing General Aviation Airplane. AIAA-80-1843, Aug. 1980.

11. Weick, Fred E.; and Wenzinger, Carl J.: Effect of Length of Handley Page Tip

Slots on the Lateral-Stability Factor, Damping in Roll. NACA TN No. 423, 1932.

12. Weick, Fred E.; and Wenzinger, Carl J.: Wind-Tunnel Research Comparing Lateral

Control Devices, Particularly at High Angles of Attack. VII - Handley Page

Tip and Full-Span Slots With Ailerons and Spoilers. NACA TN No. 443, 1933.

13. Weick, Fred E.; and Wenzlnger, Carl J.: Wind-Tunnel Research Comparing Lateral

Control Devices, Particularly at High Angles of Attack. I - Ordinary Ailerons

on Rectangular Wings. NACA Rep. No. 419, 1932.

17

14. Weick, Fred Ernest; Sevelson, Maurice S.; McClure, James G.; and Flanagan,

Marion D.: Investigation of Lateral Control Near the Stall. Flight Investi-

gation With a Light High-Wing Monoplane Tested With Various Amounts of Wash-

out and Various Lengths of Leading-Edge Slot. NACA TN 2948, 1953.

15. Weick, Fred E.; and Abramson, H. Norman: Investigation of Lateral Control Near

the Stall. Flight Tests With High-Wing and Low-Wing Monoplanes of Various

Configurations. NACA TN 3676, 1956.

16. Kroeger, R. A.; and Feistel, T. W.:

Through Wing Aerodynamic Design.

Apr. 1976.

Reduction of Stall-Spin Entry Tendencies

[Paper] 760481, Soc. Automot. Eng.,

17. Feistel, T. W.; Anderson, S. B.; and Kroeger, R. A.: A Method for Localizing

Wing Flow Separation at Stall To Alleviate Spin Entry Tendencies. AIAA

Paper 78-1476, Aug. 1978.

18. Szelazek, C. A.; and Hicks, Raymond M.: Upper-Surface Modifications for

Improvement of Selected NACA 6-Series Airfoils. NASA TM-78603, 1979. C_max

19. Hassell, James L., Jr.; Newsom, William A., Jr.; and Yip, Long P.: Full-Scale

Wind-Tunnel Investigation of the Advanced Technology Light Twin-Engine Airplane

(ATLIT). NASA TP-1591, 1980.

18

TABLEI.- GEOMETRICCHARACTERISTICSOFTESTAIRPLANE

Wing (basic) :Span, ft .................................................................... 24.46Area, ft 2 ................................................................... 98.11Design wing loading, Ibf/ft 2 ................................................ 15.89Root chord, ft ............................................................... 4.00Tip chord, ft ................................................................ 4.00Meanaerodynamic chord _, ft ................................................ 4.00Aspect ratio ................................................................. 6.10Dihedral, deg ................................................................ 5.0Incidence :

3.5At root, deg ...............................................................At tip, deg ................................................................. 3.5

Airfoil section ............................................. Modified NACA642-415Aileron (each) :

Area, ft 2 .................................................................. 2.60Span, ft ................................. :. ................................ 3.82Chord, ft .................................................................. 0.68

Flap (each_:Area ft _ 2 72

, o ee o e eoo e o e o o o e oee eoo e o o e e eelee • eoe ee • I • • • • oe • oe J ee e'e • ee e ee e e e me e •

Span, ft ................................................................... 3.760.68Chord, ft ..................................................................

Horizontal tail:

Span, ft ...................................................................... 7.69-3.0

Incidence, deg ...............................................................

Root chord, ft ............................................................... 3.60I .67

Tip chord, ft ................................................................

Mean aerodynamic chord, ft ................................................... 2.75

section NACA 651-012eooeeeeeeoeoeeemeeeeeoeleeeeeeeoeooooeoee'eooeeeoeeseeAirfoil

Tail length (distance 0.25_ to 0.25 mean

aerodynamic chord of tail), ft ............................................ 11.62

Elevator:

Area (total), ft 2 .......................................................... 7.22

1.13Root chord, ft .............................................................0.70Tip chord, ft ..............................................................

Span, ft ................................................................... 7.69

Area (forward of hinge line at tip), ft 2 ................................... 0.92

Vertical tail:4.09

Span, ft .....................................................................3.60Root chord, ft ....................... .................... ....................1.67

Tip chord, ft ................................................................

Airfoil section ...................................................... NACA 651-012

Rudder:

Area (total), ft 2 .......................................................... 3.61

1.13Root chord, ft ..................................................... ,........

Tip chord, ft .............................................................. 0.70

Span, ft ................................................................... 4.09

Area (forward of hinge line at tip), ft 2 ................................... 0.46

Propeller diameter, ft ......................................................... 5.92

Propeller pitch, in ............................................................ 46

19

TABLEII.- COORDINATES OF AIRFOIL SECTIONS USED IN TESTS

tations and ordinates given in_percent of airfoil chord

(a) Coordinates of modified NACA 642-415airfoil (basic wing)

Upper surface Lower surface

Station Ordinates Station Ordinates

0

.299

.526

.996

2.207

4.673

7.162

9.662

14.681

19.714

24.756

29.803

34.853

39.904

44.954

50.000

55.O4O

60.072

65.096

70.111

75.115

80,109

85.092

90.066

95.032

100.000

0

1.291

1.579

2.038

2.883

4.121

5.075

5.864

7.122

8.066

8.771

9.260

9.541

9.614

9.414

9.016

8.456

7.762

6.954

6.055

5.084

4.062

3.020

1.982

.976

0

0

.701

.974

1.504

2.793

5.327

7.838

10.338

15.319

20.286

25.224

30.197

35.147

40.096

45.046

50.000

54.960

60.000

65.000

70.000

75.000

80.000

85.000

90.000

95.000

100.000

0

-1.091

- 1. 299

-1.610

-2.139

-2.857

-3.379

-3.796

-4.430

-4.882

-5.191

-5.372

-5.421

-5,330

-5.034

-4.604

-4.076

-3,698

-3.281

-2.865

-2.343

-1.875

-1.458

-.990

-.573

0

20

TABLE II.- Continued

(b) Coordinates of leading-edge-droop airfoil

Upper surface Lower surface

Station Ordinates Station Ordinates

-3.833

-2.885

-1.633

-.817

-.190

.350

.875

1.254

1.604

1.983

2.319

2.883

4.121

5.075

5.865

7.123

8.065

8.771

9.258

9.542

9.615

9.415

9.017

8.456

7.763

6.954

6.056

5.085

4.063

3.021

1.983

.977

.000

-2.769

-2.658

-2.217

-1.773

-1.329

-.885

-.444

.000

.444

.885

I .329

2.206

4.673

7.163

9.662

14.681

19.715

24.756

29.802

34.852

39.904

44.952

50.000

55.040

60.071

65.096

70.110

75.115

80.108

85.O92

90.065

95.031

100.000

-2.769

-2.658

-2.217

-1.773

-1.329

-.885

-.700

.000

.444

.885

1.329

2.206

4.673

7.163

9.662

14.681

19.715

24.756

29.802

38.585

40.096

45.046

50.000

54.960

60.000

65.000

70.000

75.000

80.000

85.000

90.000

95.000

100.000

-3.833

-4.631

-5.540

-5.983

-6.160

-6.210

6.225

-6.210

-6.201

-6.191

-6.182

-6.164

-6.111

-6.059

-6.006

-5.90O

-5.793

-5.687

-5.580

-5.394

-5.330

-5.034

-4.604

-4.076

-3.698

-3.281

-2.865

-2.343

-1.875

-1.458

-.990

-.573

.000

21

TABLEII.- Continued

(c) Coordinates of exaggerated leading-edge-droop airfoil

Upper surface Lower surface

Station Ordinates Station Ordinates

-5.000-4.167-3.333-2.500-I .667-.833

.000

.4171.0422.2074.6737.1629.662

14.68119.71424.75629.80334.85339.90444.95450.00055.O4O60.07265.09670.11175.11580.109

85.092

90.066

95.032

100.000

-6.667

-4.521

-2.938

-1.667

-.625

.313

1.042

1.417

1.896

2.883

4.121

5.075

5.864

7.122

8.066

8.771

9.260

9.541

9.614

9.414

9.016

8.456

7.762

6.954

6.055

5.084

4.062

3.020

I .982

.976

0

-5.000

-4.167

-3.333

-2.500

-1.667

-.833

.000

.417

1.042

2.207

4.673

7.162

9.662

14.681

19.714

24.756

29.803

34.853

40.096

45.046

50.000

54.960

60.000

65.000

70.000

75.000

80.000

85.O00

9O.0OO

95.000

100.000

-10.979

-11.688

-11.833

-11.667

-11.537

-11.406

-11.276

-11.210

-11.113

-10.930

-10.544

-10.155

-9.764

-8.678

-8 •191

-7.401

-6.612

-5.811

-5.330

-5.034

-4.604

-4.076

-3.698

-3.281

-2.865

-2.343

-1.875

-1.458

-.990

-.573

0

22

TABLEII.- Continued

(d) Coordinates of LS(1)-0417 leading-edge-droop airfoil

Upper surface Lower surface

Station Ordinates Station Ordinates

-1.375-.833-.4170

.6251.2502.2074.6737.1629.662

14.681

19,714

24.756

29.803

34.853

39,904

44.954

50.000

55.040

60.072

65.096

70.111

75.115

80.109

85,092

90,006

95.032

100.000

-2.000 -1.375

0 -.833

.604 -.417

1.145 0

1.688 .625

2.188 1.250

2.883 1.875

4.125 2.500

5.075 3.750

5.864 4.375

7.122 6.250

8.066 9.229

8.771 12.333

9.260 30.530

9.541 35.147

9.614 40.096

9.414 45.046

9.016 50.000

8.456 54.960

7.762 60.000

6.954 65.000

6.055 70.000

5.084 75.000

4.062 80.000

3.020 85.000

1.982 90.000

.976 95.000

0 100.000

-2.000

-3.188

-3.750

-4.063

-4.396

-4.646

-4.833

-5.042

-5.438

-5.604

-6.021

-6.417

-6.063

-5.417

-5.375

-5.330

-5.034

-4.604

-4.076

-3.698

-3.281

-2.865

-2.343

-1.875

-1.458

-.990

-.573

0

23

TABLEII.- Concluded

(e) Coordinates of upper-surface-modification airfoil

Upper surface Lower surface

Station Ordinates Station Ordinates

0

.104

.500

1.000

2.000

3.000

5.000

7.000

10.000

15.000

20.000

25.000

30.OOO

35.000

40.000

45.000

50.000

55.040

60.072

65.096

70.111

75.115

80.109

85.092

90.066

95.032

100.000

0 0

1.250 .701

2.270 .974

3.150 1.504

4.290 2.793

5.130 5.327

6.310 7.838

7.150 10.338

8.020 15.319

8.900 20.286

9.333 25.244

9.520 30.197

9.600 35.147

9.600 40.096

9.600 45.046

9.420 50.000

9.016 54.960

8.456 60.000

7.762 65.000

6.954 70.000

6.055 75.000

5.084 80.000

4.062 85.000

3.020 90.000

1.982 95.000

•976 100.000

0

0

-1.091

-1.299

-1.610

-2.139

-2.857

-3.379

-3.796

-4.430

-4.882

-5. 191

-5.372

-5.421

-5.330

-5.034

-4.604

-4.076

-3.698

-3.281

-2.865

-2.343

-1.875

-1.458

-.990

-.573

0

24

TABLEIII.- PRESENTATIONOFWIND-TUNNELDATAFigure

Lift characteristics:Basic-wing lift variation ...............................................Basic-wing tuft and oil-flow photographs ................................Modified-wing tuft photographs ..........................................

1415 and 16

17

Comparisonof wing modifications:Basic leading-edge droop ................................................ 18 to 21Wing-fillet droop ....................................................... 22Segmentedleading-edge droop ............................................ 23 and 24LS(I)-0417 leading-edge droop ........................................... 25Exaggerated leading-edge droop .......................................... 26Leading-edge slats ...................................................... 27 and 28Chord extension ......................................................... 29

3OWing fences .............................................................Leading-edge stall strips ............................................... 31Wing upper-surface modification ......................................... 32

Chordwise pressure coefficients:

Basic wing .............................................................. 3334Modification B ..........................................................

35Modification C ..........................................................

Modification A .......................................................... 36

Spanwise load distribution:

Basic wing; B, C, and A modifications ...................................

Segmented leading-edge modifications ....................................

Slat configurations .....................................................

37

38

39 and 40

Drag characteristics:

Basic-droop leading-edge modifications .................................. 41

Segmented leading-edge modifications .................................... 42

Slat configurations ............. • ....................................... 43

Complete airplane:

Longitudinal characteristics ............................................. _... 44 to 49

Lateral characteristics .............................................. ,... 50 to 56

25

I I

o

/-Outboard panel parting line

_ing balance

_2____J_°_--5.17-_

Diameter 5.92-__

Figure I.- Three views of test airplane. Dimensions are in feet.

26

N -..J Figure 2.- Test airplane in Langley 30- by 60-Foot Tunnel.

I

£

/

\i!

I I i• 38 b/2 1 I

I 441 b/2 II"I I .63 b[2/

I II I tL p I I

I i I

J I ii t i

• 78 b/2 I II .85 b/2

I i .92 b/2

I I [ wI

Figure 3.- Spanwise locations of rows of static-pressure ports on

right wing of airplane.

28

!c_J

-r-

4-

v°_o

O.Pl

Ig-,-.4

Ig-,-4

29

_Leading-edge . . .

droop

trOll

-_ F--o.o_Section A-A(enlarged)

"0.95 b/2

(a) Full-span droop

(modification A).

/

O.95b12

F--0.57 b12--

(b) Original outboard droop

(modification B).

f

JO.95b/2

/ I

I 0.57 bl2--A _,_0.41

I--"

I(c) Original outboard droop

plus inboard fairing

(modification C).

0.95bt2

0.72 b12--

I(d) Short outboard droop

(modification D).

Figure 5.- Leading-edge-droop modifications studied in flight and wind-tunnel tests.

30

-'1I--o.o3_

Section A-A

(enlarged)

0.g5 1_2 /_

Inboard extension | I

(e) Long outboard droop

(modification E).

JO.g5hi2 /_,

_0.57 b/2

_ 0.46b/2 0

---,-A ,_.

(f) Narrow-gap segmented droop

(modification F).

/_0. g5 b/2 /-'-'

O.35 r2.

O.

k-_A

(g) Wide-gap segmented dro

(modification G).

Figure 5.- Concluded.

31

00__ _ Dr00p0n,_0pSection A-A

(enlarged)

f==-,-A

O.95hi2

0.57 b12

J i---_A

(a) Original outboard droop ("B")

+ Exaggerated droop.

---_A

O.95 b/2

0.57 b/2

J I-_,-A

I

(b) Full-span droop ("A") + Exaggerated droop.

Figure 6.- Geometry of exaggerated leading-edge droop.

32

Section A-A

(enlarged)

--_A

O.9.5b/2

F 0.57 b/2

l-_A

(a) Outboard droop.

--A

-_A

0.95 b/2

(b) Full-span droop.

Figure 7.- Geometry of LS(I)-0417 droop configuration used in tests.

33

F Upper-s u rface modification

-0.426-_

Section B-B

( en la rged )

O.95 b/2

--_B

J ]

O.57 b/2

i

(a) Outboard modification.

1

Figure

P-_B

d I-_B

O.9.5b/2

(b) Full-span modification.

8.- Geometry of upper-wing-surface modifications used in tests.

34

4

_rAU_A

0 O.01_.o_-,,,,_-,<.--_-_---

,,oy___o.o_Section A-A

(enlarged)

O.95 b/2

1"i

I I

O.57 b/2 t

/

(a) Outboard slat.

4

_r A_A

I I

0.95 b/2

(b) Full-span slat.

Figure 9.- Geometry of 0.08_ slat used in tests.

35

0.1_ _ 0.01_:

Section A-A

(enlarged)

0.95 b/2

•'------ O.57 b/2

I A

(a) Outboard slat.

I

O.95b/2

t 1

(b) Full-span slat.

Figure 10.- Geometry of 0.155 slat used in tests.

36

_--0 57 b/2

B

B

Section B-B

(enlarged)x

(a)O.005C(b)O.021_(c) 0 052_

(a) Basic wing.

Figure 11.- Geometry of leading-edge stall strips used in tests.

37

0.95 b/2 r

0.57 hi2

_A

--,-A

0.35 b12

X

Section A-A(enlarged)

X

(a)0.005C(b)0.021C(c) 0.052_,

(b) Drooped leading edge.

Figure 11.- Concluded.

38

_ - O.57 b/2--

-A

Skewed

"- B\ If Skewedin

B

=-- O.57

I

b12

"C

i 1 I

U,,-,,,

I

.-D

SectionA-A(enlarged)

(a)

0.06_J

o_ _oo_Short fence -__----_.._

0.17_-_Section B-B(enlarged)

(b)

-t I-o. Section C-C(enlarged)

(c)

c fillet

it droop

SectionD-D(enlarged)

(d)

Figure 12.- Geometry of wing fences, chord extension, and fillet droop

used in tests.

39

1

--

I'"

I J gl_ .

!_,

L ----

t"

": I

u _ l

_..e-_ 4,_I

i i._ _ _"_'- I_'" _'-"

I)

(..I ,

**

,I.-. _ rh

_-_ gl

.r,.

L_

.---i

_,.,.

or4-I,

-r

( .,,i

I .o_

I,i..i

°_

0._.IA

0

!

O_

I

4O

CL, wp

41

42

a = 20 0 , low lift

L-82-141 a = 20 0 , high lift

Figure 15.- Tuft surveys showing fluctuations in stall pattern on full-scale model with basic wing.

1- --~. ~-----.-

L-82-142

Figure 16.- Oil-flow surveys of 1/3-scale model with basic wing showing fluctuations in wing stall pattern.

43

44

a = 30°

L-82-143

Figure 17.- Tuft surveys made on full-scale model with wing modification B.

----- --- --- - - - - -- -~---- ----~-~ ---- -- ----'

=--.

4-

_, +

=

o= ._-_u

0 o u.

oD04

.(',.I

t_--_ d_J, _ cf

_ _P"" _r,. /

J

• . . • 0 .

J

o

o

o'3"

o

i

:>

•_ ,._.,.4

c _!

_go ,_

•_, _

!

_o

i1)

g-,.4

45

.6

0

Confi gurati on

Basic wing

---------- Original outboard droop ("B")

..... Original outboard droop + inboard fairing ("C")

------==- Full-span droop (A")

°

\

-.4

-.6 --

-.8

0

Stable

/

\ Unstable

\

I

/

J I _ I I I I I _ l10 20 30 40 50

c_, deg

Figure 19.- Effect of leading-edge-droop modifications on autorotational

characteristics of airplane.

46

cL

CD

1.8

].6

1.4

Configuration

0 Basic wing

[] Short outboard droop ("D")

Original outboard droop ("B")

/_ Long outboard droop ("E")

_, Full-span droop ("A")

22

1.2

].0

8

.6

.4--

L

d v

f

CR

2.0

1.8

1.6

1.4

1.2

l.O

.8

6

.4

L .2'

-.2

-.4 0

-lO 0 10 20 30 40 50 -lO 0 lO 20

_, deg o, deg

30 40 50

Figure 20.- Effect of variation in length of leading-edge droop on longitudinal

aerodynamic characteristics of airplane. Tails off.

47

bCR

_o

-2 -

-4 -

-6

-8 -

_.=

_R

- ,=)

>

=_o,

r

A

E

::}o

m

._:2_

o

w,-,,

O

"I:O

.CCzb

JII

4

Force data

a, deg

0 20[] 25

30A 35E_ 40

c

.,-, Stable

Unstable

Most outboard location of discontinuity forleading-edge-droop effectiveness in flight tests

Most inboard location of discontinuity forleading-edge-droop effectiveness in flight tests

_"'_I I_ I

I I I t I

0 20 40 60 80 100

Spanwise distance, percent semispan

Figure 21.- Variation of autorotational stability provided by leading-edge droop

as function of spanwise location of droop inboard discontinuity.

48

CL

cD

16

14

1.0

6

-]0

4

£

i4

=

Configuration

0 Basic wing

[] Full-span droop ("A")

Full-span droop ("A") + fillet droop

[ll

II

I

." D

Ji

: _0" Il

2.2

2.0

1.8

1.6

1.4

t.2

CR

l.O

,8

.6

.4

I

,2'

0J

0 lO 20 30 40 50 -lO 0

_, deg

30 40 50

Figure

(a) Modification A.

22.- Effect of wing/fillet droop fairing on characteristics of

leading-edge configurations.

three

49

Configuration0 Basicwing[] Longoutboarddroop("E")<_Longoutboarddroop("E")+ fillet droopA Longoutboarddroop("E")+ fillet droop+

inboarddroopextensionof 0,095b/22.2

!//

30 40 50

(b) Modification E.

Figure 22.- Continued.

50

CL

C D

1.4

1.2

l.O

.8

.6

.4

.2

i

0

l-.2_

-.4

-I0

Configuration

0 Basic wing

[] Original outboard droop ("B")

<_ Original outboard droop ("B") +

/ f

r d

0 I0 20 30 40 50

a, deg

1

CR

fillet droop

1 , i

__ #lj, ....

IF

-lO 0 I0 ZO 30 40 50

_, deg

(c) Modification B.

Figure 22.- Concluded.

51

Configuration

0 Basic wing

[] Large segment ("G")

Small segment ("F")

m, deg

CR l.0

,8

0

40 50 -l0 0 l0 20 30 40

_, deg

50

Figure 23.- Effect of segmented leading-edge droop on longitudinal aerodynamic

characteristics of airplane. Ta£1s off.

52

BCR

.6

.4

.2

-.2

-.4

-.6

Configuration

Basic wing

._-_ Outboard droop ("B")

..... Large segment ("G")

--=---=--- Small segment ("F")

//

Stable

Unstable

I

0 I0 20 30 40 50

_, deg

Figure 24.- Effect of segmented leading-edge droop on autorotational stability

characteristics of airplane.

53

1.6

1.4

1.2

l.O

.8

CL

.6

CD.4

.2

0

-.2

-.4

-I0

Configuration

0 Basic wing

[] Outboard droop

Full-span droop

I

f

°. ,

0

_ _ 1.2 ,

.4

0

I0 20 30 40 50 -lO 0 I0 20 30 40 50

o, deg 6, deg

Figure 25.- Ef6ect of LS(I)-0417 shape droop on longitudinal aerodynamic

characteristics of airplane.

54

C L

CD

1.8

1.6

1.4

1.2

l.O

_o

,j

Configuration

0 Basic wing

[] Original outboard droop ("B") + exaggerated droop

<_ Full-span droop ("A") + exaggerated droop

/_ Full-span droop ("A") + exaggerated droop + fillet droop

! _ 2.2

z

_z : 2.0

. ./

-- "-- =e--_,e_ •

,_Ir I _ " 1.6

mm

JJ

aFI

,3

CR

1,4

1.2

l.O

,8 ""

-.4

-l 0 0 10 20 30 40 50 20 30

_, deg e, deg

\

-lO 0 lO

q

40 50

Figure 26.- Effect of exaggerated droop on leading edge of outboard wing panel.

55

CL

CD

14

1.2

]°0 ....

.8 --

.6

.4

.2

f/ ,

-.2_

-.4

-I0 0 lO 20 30 40 50

Slat configuration

0 Basic wing

[] 0.o8_

.15_

CR

1.8

t t/#

1.2

1.O

.8

.6

.4

.2

oi-I0 0 lO 20 30 40 50

_, deg _, deg

(a) Outboard slats.

Figure 27.- Effect of leading-edge slats on longitudinal aerodynamic

characteristics of airplane. Tails off.

56

CL

cD

2.2

2.0

1.8

1.6

1.4

1.2

1.0

.8

.6

.4

-.2

-.4-lO

/,

{

r

N,

1

0 10 20

e',deg

Slat configuration

0 Basic wing

[] 0.15_

<_ .15_ + fillet droop

/_ .08_ + fillet droop

2.6

_._

.,

,)-q9

j

30 40 50

CR

r

2.4,

2.2 ....

2.0 ---

1.8 --=_

1.6 ----

1.4 k__

1.2

l,O --

.8 -- --

.6

.4

r

q

.2 _ =

_W0-lO

!I

I

............ L_

ira_

---4 ......... l_ m

l

/

I

I

V/

I ..........

J- _ t .....t

10 20 30 40 5O

_, deg

(b) Full-span slats.

Figure 27.- Concluded.

57

_CR

.4

.2

0

Slat configuration

Basic wing

Original outboard droop ("B")

O. 08_

.15_

Stable

-.2Unstable

-.4 J I J I I I0 l 0 20 30 40

_, deg

(a) Outboard slats.

l I5O

Figure 28.- Effect of leading-edge slats on autorotational stability characteristics

of airplane as compared to leading-edge-droop configuration.

58

_CR

.6--

,#

J

Slat configuration

Basic wing

\ -----_ Original outboard droop ("B")\

% ..... 0.15_

\ _-_ .08_ + fillet droop

%

\ //Stable

Unstable

I0 20 30

_, deg

i _ I i I l I40 50

(b) Full-span slats.

Figure 28.- Concluded.

59

14 [

1.2 1-i

1.0--

.8,

CL

.4 --

CD

o I

i

lJ

-,2 _#,

-.4 "

-10

Configuration

O Basic wing

[] Chord extension

r i

b

[P,-I!

WI - :J

/H

CR

1.6

1.4 .... i-- ----- __Z1.2i-

f.0 ...... _ \s...

.6

.4

0-10

iJ

0 10 20 30 40 50 0 l0 20 30 40 50

_, deg _, deg

Figure 29.- Effect of extended outboard wing chord on longitudinal aerodynamic

characteristics of airplane.

60

1.4

1.2

Configuration

0 Basic wing

[] Long fence

Short fence (skewed in)

/_ Short fence (skewed out)

1.8

1.6 ---

10 ZO 20

_, deg a, deg

30 40 50

Figure 30.- Effect of wing fences on longitudinal aerodynamic characteristics

of airplane.

61

1.8

i

1.6

1.4

1,2

1.0

CL.8

CD |

W

",_0 " --''

N_

-10 0 I0 40 50

(a) Basic wing.

Figure 31.- Effect of 0.005_ stall strips on longitudinal aerodynamic

characteristics of airplane.

62

CL

CD

1.0

I0 20

_, deg

Confi gura ti on

0 Basic wing

[] Full-span droop ("A")

Full-span droop ("A") + stall strip

2.2 ......... ]___

CR

1.8 -- ---

1.6 __L--_ - --- _---- --

1.0 .....

.8

° ii.4

.2

0

-I0 0 I0 20 30 40 50

, deg

(b) Full-span droop ("A") •

Figure 31.- Concluded.

63

CL

CD

1.6

1.4

1.2

1.0

.8

.6

.4

.2

0

-.2

°.4

-I0

i __ - :

/_ : : i

: : |

• I

0 10

Confi guration

0 Basic wing

[] Outboard modification

<_ Full-span modification

/_ Full-span modification+ fillet droop ....

T + _ +:: :

1,6 +- ................

1.4 ........ _

CR .8

::+

,4 _LJ _

0

E

i

iJ

20 30 40 50 -I0 0 10 20 30 40 50

_, deg _, deg

Figure 32.- Effect of airfoil upper-surface modifLcation on longitudinal

aerodynamic characteristics of airplane.

64

cp

Cp

cP

-6

-4

-2

0 (_(_

2 --

4-.2 0 .2

•_bo_> o ) o'0

L _q

.4 .6 .8x/c

Station l, y/(b/2) = O.g2

1.0

-6

-4 L

-2

0 )

2

4-.2 0 .2

' wt_

I].4 .6 .8x/c

Station 2, yl(b/2) = 0.85

1.0

-6

-4

:3

-2 ----(_) 0

0

) 0

-.2 0 .2 .4 6

x/c

Station 3, y/(b/2) = 0.78

.8 0

Cp

Cp

Cp

-6

-4

-2

0

2

4

-6

-4

-2

0

2--

4

-8

-6

-4

-2

0

2

0

-.2 0 .2 .4 .6 .8 1.0

x/c

Station 4. y/(b/2) = 0.63

o c o

-.2 0 .2 .4 .6 .8 l.Ox/c

Station 5, y/(b/2) = 0.44

3

v

-,2 0 .2

o4 0! o

.4 .6 ,8 1.0

x/c

Station 6, y/(b/2) = 0.38

0 Basicwing

+ Denoteslowersurface

(a) _ = 11.2 ° •

Figure 33.- Chordwise pressure coefficients for basic wing.

65

-8

-6

(

-4

Cp 13o..

o )0 ) o OQ0 ...... _--_ _-'-"--_

2-.2 0 .2 .4 .6

x/c

Station 1. y/(b/2) = 0.92

CP

Cp

-6 )

.8 1.0

-4

-2

0

2

0

c__o c;)

oc o( o _2_

4-.2

-6

-4

-2 --

0-

2

4

.2 .4 .6x/c

Station 2. yl(b/2) = 0.85

.8 1.0

)0

0)

e® )

.2 0 .2 .4 .6x/c

Station 3. y/(b/2) = 0.78

o oc o(,

-8

-6

-4

Cp

-2

0

2

-6

-4

-2

Cp0

-6

-4

-2

Cp0

2

4

0

_0

_c

0 0 ( 0_c v

-.2 0 .2 .4 .6 .8 1.0

x/c

Station4,yl(b12)= O.63

-.2

(Be

)o o¢;i

.2 .4 .6 .8

xlc

Station5.yl(b/2)= 0.44

D

o(

1.0

.8 ]0 -2 o .2 .4 .6 .8x/c

Station 6, y/(b/2) = 0.38

1.0

0 Basicwing

+ Denoteslowersurface

(b) a = 21.6 ° •

Figure 33.- Continued.

66

Cp

-6

-4

-2

0

2

Cp

Cp

-4

.2 0

_Oc o 0 o--_

®®

.2 .4 .6

xlc

Station 1. yl(b/2) = 0.92

.8 1.0

-2

0

2

4

6

-4

-2

_x)( 0 o ) 0

®®

OO

-.2 0 .2 .4 .6

x/c

Station 2. y/(b/2} = 0.85

,8 I.O

0 ----

2

4

6-,2

)_D .3P°(.' O O ( O

®®

0 .2 .4 .6

x/c

Station 3. y/(b/2) = 0.78

.8 l.O

-6

-4

Cp

-2

0 --i----.

2

o

@(9 (

0 o

-.2 0 .2 .4 .6

x/c

Station 4, y/(bl2} = 0.63

.8 1.0

-4

Cp

-2

0--

2

_Z)O() 0 ) 0

(9(9 ")

6-.2 0 .2 .4 .6

x/c

Station 5, y/(b/2) = 0.44

.8 1.0

-4

Cp

-2

_z)o o( o0 (

®® (

2

4--

.2 .4 .6 .8x/c

Station 6, y/(b/2)= 0.38

1.0

(c) _ =-31.9 °.

Figure 33.- Continued.

0 Basicwing

+ Denoteslower surface

67

-4

-2

0

Cp

2

4

6

Cp

Cp

-4

-2

0

2

4

6

-4

o ( 0

.2 0 ,2 .4 .6x/c

Station 1, y/(b/2)= 0.92

.8 l.O

i

I_EXD_ 0 ( 0 0 ) O0

®®

.2 0

-2

0

2

4

6-.2 0

.2 .4 .6 .8x/c

Station2, y/(b/2) = 0.85

1.0

)0 )00

1

.2 .4 .6

xlc

Station3,y/(b/2)= 0.78

.8 l.O

Cp

Cp

Cp

-4

-2

0

2

(_ 0 0 ) 0

-.2 0 .2 .4 .6x/c

Station 4, y/(bl2) = 0.63

D

L_DO< O ) O ) O

®® (

.2 0 .2 .4 .6x/c

Station 5, y/(bf2)= 0.44

-4

-2

0

2

4

.8

6-.2 0

.8

(_DO 0 ) O ) 0 O0

G® (

1.0

.2 .4 .6

xlc

Station 6, y/(b/2) = 0.38

1.0

.8 l.O

0 Basicwing

+ Denoteslowersurface

(d) _ = 41.5 o .

Figure 33.- Concluded.

68

Cp

Cp

CP

-4

-2

0

2

4

6-,2

(lIE)O(

m_l _ H..

0 o) o)

0 .2 .4 .6x/c

Station ], y/(bt2)= 0._2

.8 1.0

-4

-2!

0!--Fo ! o_e¢-

-.2 0 .2 .4 .6

x/c

Station 2, y/(bl2) = 0.85

Cp

-4

-2

0

2

_'0_)_ 0 ( 0

tl _"

-.2 0 .2 .4" .6

x/c

Station3, yl(b/2)= 0.78

Cp

-6

-4

-20

2

4-,2

o0

o,_

0 .2 .4 .6 .8xlc

Station 4, y/(b/2) = 0.63

-6

-4

"_ ___)Oc 0 ( 00_@-- _ )_-_ ._

(9

2

4.8 1.0 -.2 0 .2 .4 .6 .8

xlc

Station 5, y/(b/2) = 0.44

-8

-6

C

P

-4 _r-

-2 %r_

o )olo

0 _

2 0 12 .4 .6xlc

Station 6, y/(b/2) = 0.38

1.0

1.0

.8 1.0 -.2 .8 1.0

0 Basic wing

Original outboarddroop("B")

+ Denotes lowersurface

(a) _ = 11.2 ° •

Figure 34.- Chordwise pressure coefficients for wing with

modification B.

69

cp

Cp

Cp

-4

-2

0

2

4

6.2

-6

-4

0 ........

2

4

-6

-4

-it --

0 --_

2

4L.2

0 ( 0 ) 0 () c)A

0 .it .4 .6

xlc

StationI,yl(bl2)= 0.92

.8 1.0

_0¢) 0

......

.2 .4 .6 .8

xlc

Station2,y/(b/2)= 0.85

1.0

0) 0 )0

F

.it .4 .6

xlc

Station3,y/(bl2)= 0.78

.8 1.0

-6

cp

-4

-2

0

2

4.2

t,,

"o00o(

0 0 )0[

.it .4 .6

x/c

Station 4, yt(b/2) = 0.63

.8 1.0

-6

-4 ..........

D-2

0 ..... _ ....

®

It

4-.2 0 .2 .4 .6

xlc

Station5.yl(bl2)= 0.44

.8 1.0

cp

-4

-2 --

0

It

4-.2 .2 .4 .6 .8

x/c

Station 6, yl(b/2) = 0.38

v

1.0

(b) o( = 21.6 °.

Figure 34.- Continued.

0 Basic wing

13,Originaloutboarddroop("B")

+ Denotes lower surface

T

70

CP

Cp

C

p

-8

-6

-4

2

0

h

F,

1" -_J(

0

-8

o )0 ) o 3(_P

.2 .4 .6

x/c

Station1,yl(b/2)= 0.g2

-6

-4

-2 -- _ --C

0 (

0_. _,-

22 0 2

-8

-6

4

-2 ..-00

0

h_ h,.2-.2 0 .2

.8 1.0

O( o()oo

.4 .6

x/c

Station 2, y/(b/2 ) = 0.85

.8 1.0

o )0 ) 0 oo

__ )----.-_ _-

.4 .6

xlc

Station3,yl(br2)= 0.78

.8 1.0

-6

Cp

-4--

-2

O

2

(iZ_)O()0 0 ( 0

4

-4

0 .2 .4 .6

xlc

Station4.yllbi2)= 0.63

.8 1.0

Cp

-2

IDO<0

2

o o.)

4 ........

6O .2 .4 .6

x/c

Station 5. y/(bi2) = 0.44

.8 1.0

-4

Cp

-2

0

2

O O ()

®® )

o_22

6-._ o .2 .4 .6

xtc

Station 6. y/(b/2 ) = O.38

.8 1.0

(c) a = 31.9 ° •

Figure 34.- Continued.

0 Basic wing

tx Original outboard droop ("B")

+ Denoteslowersurface

71

cP

cp

cP

-6

-4 ----

-2 --

0--

2

00 (

4-,2

-8

-6

-4 ----

-2

0

2-,2

0 .2 .4 .6

x/c

Station 1. y/(b/2) = 0.92

(1_)0¢

I, O( 0 )0

)

0 2 4 _6x/c

Station2.y/(b/2)= 0.85

-8

-6

-4

-2

0 _)

0

2-,2

t,

0 o (

(

.8

.8

o (

0 .2 .4 .6

x/c

Station3,yl(bl2)= 0.78

O0

1.0

) O0

1.0

O0

• v

.8 1.0

-6

Cp

-4

h

-2 I'

0---_

2

4

.2 0

-4

(_DC_ 0 0 ) 0

" )

.2 .4 .6

x/c

Station4.yl(b/2)= 0.63

Cp

-2 ......

_o o(o(0--

(®

2

o

6-.2 0 .2 .4 .6

x/c

Station 5, y/(bl2)= 0.44

-4

cP

-2--

0

2--

4

6.2

(_$DO 0 o )o_i

)®® (

10 .2 .4 6

xlc

Station 6, yl(b/2)= 0.38

(d) (z = 41.5 ° .

Figure 34.- Concluded.

OOee

.8 1.0

.s ii0

oo

_t-c_e

8 I.O

O Basicwing

F. Original outboard droop("B")

+ Denotes lowersurface

72

cp

cp

Cp

-4

-2

0---_

2

4

6.2 0

-4

-2

0

2

4

6-.2

-4

-2

0

2

4

6-.2

-6

)O)o_0

.2 .4 .6 ,8

xlc

Stationi,y/(bl2)= 0.92

-4

-2 .I12

cp

0 I,[

2

41.0 -.2 O

-6

-"--)o,0o

.2 .4 .6

x/c

Station4,y/(bl2)= 0.63

0 _

.8 1.0

0 Basicwing

I',, Original outboard droop + fairing ("C")

+ Denotes lower surface

E<_)O 0 ) 0) 0

-4

-2

Cp

0

0

0

2 ............

0 .2 .4 .6 .8 1.0

x/c

Station2,yl(b/2)= 0.85

4-.2

-8

-6!

-4

Cp

-2 --

0 .2 .4 .6 .8 1.0

xlc

Station5.yl(bl2)= 0.44

D

v

0) o

0 c)

2.2 0 .2 .4 .6

xlc

Station6,yl(b/2)= 0.38

0 .2 .4 .6 .8 I.O .8 I.O

x/c

Station 3. y/(b/2)= 0.78

(a) (:z = 11.2 ° .

Figure 35.- Chordwise pressure coefficient for wing with

modification C.

73

-6

-4

Cp ' 0 0 ) 0

o _-

2

-.2 0 .2 .4 .6

xlc

Station i, y/(b/2J = 0.92

.8 l.O

-6

Cp

-4

-2 h,._IX-_Oq 0 (

o

0t_ _-

2

40 .2 .4 .6 .8

x/c

Station 2. y/(b/2 )= 0.85

1.0

-6

Cp

-4

-2 _0(

0o( o_

4-.2 0 .2 .4 .B .8 l.O

xIc

Station 3. yl(b/2) = 038

Cp

Cp

Cp

-6

-2 ----_(_h

O ----

2

) 0o ( O(

4-,2 0 .2 .4 .6 .8 1.0

x/c

Station 4, yl(bl2) = 0.63

-6

-4

-2

0

2 i

4!-.2 0

-4

-2

0

2

4

6-.2

e=Do)0 )0 )0). )

.2 .4 .6

xlc

Station5.yl(bi2)= 0.44

.8 1.0

l

®® (

o o

.2 .4 .6

x[c

Station 6, yl(b/2) = 0.38

.8 l.O

0 Basicwing : : ::

Original outboard droop + fairing ("C")

+ Denotes lower surface

=- = z

±

(b) a = 21.6 ° •

Figure 35.- Continued,

74

CP

cp

Cp

-6

-4

II-2

0

2

4

-.2 0

-8

L_ ,

'%0( o o ) o

r--(

.2 .4 .6 .B

x/c

Station 1, y/(b/2) = 0.92

1.0

-6

-4

-2

0

2',2

-B

0() 0 o( o( Oo

--( _-

.2 .4 .6

x/c

Station2,yl(b#2)= 0.85

.8 1.0

-6

-4

-2

2.2 0 .2 .4

x/c

Station 3, y/(b/2) = 0.78

O 0)0)_

.3--_ _-_

.6 .8 1.0

cp

Cp

-6

-4

-2 --

0

2

4

.2

-4

h

(_0< 0 o _ 0

0 .2 .4 .6x/c

Station 4, y/(b/2) = 0.63

.8 1.0

-2

0C_IDCX O O ( O A__

( 4

6

-4

0 .2 .d .6 .8 1.0

xlc

Station 5. y/(b/2) = 0.44

Cp

-2

0

2--

_DO e 0 0 ) 0

EX9

4[ ..........

6-.2 0 .2 .4 ,6 .8

x/c

Station 6, y/( b/2) = O.38

(c) (I = 31.7 °.

Figure 35.- Continued.

10

0 Basic wing

I'_ Original outboarddroop + fairing ("C")

+ Denotes lower surface

75

Cp

Cp

Cp

-6

-4

-2

4.2

-6

-4

-2 __i_t

2

4

-6

-4

-2 ----

0

2 .......

4

(3_X_ 0 0 ) 0 ) O_;p

0 .2 4 .6

x/c

Station 1, y/(b/2) = 0.92

.8 l.O

"(_Dc< 0 ) 0 ) 0 P O0

0 .2 .4 .6 .8 1.0

x/c

Station 2. y/(bl2) = 0.85

G_BDO()0 ) 0 0 ( O0 Cp

.2 .4 .6

x/c

Station 3, y/(b/2) = 0.78

.8

-6

Cp

Cp

-4

-2

0 ----'

2 ------

4.2

_(_(_l 0 0 (

(

-4

-2 --

0--

2

4

6",2

0 .2 .4 .6

xlc

Station 4. y/(b/2) = 0.63

-4

-2

0-

2

0 .2 .4 .6

x/c

Station5,y/(b/2)= 0.44

0

61.0 -.2

C o(

®® (

o

) o

0 .2 .4 .6

xIc

Station 6. y/(b/2) = 0.38

(d) _ = 41.5 ° .

Figure 35.- Concluded.

.B 1.0

.8 l.O

.8

0 Basicwing

L_ Original outboard droop

+ fairing ("C")

+ Denotes lower surface

76

C

P

C

P

C

P

-4

-2

2

4

6-°2

-4

-2

0 -- --

2

4

6-.2

-4

h0

4

6-.2 0

0 0

.2 .4 .6 .B

x/c

Station I, y/(b/2) = 0._

CP

1.0

-6

-2 --_

07---

2

4.2

-6

t'h

"jOoc o0

o_

.2 .4 .6

x/c

Station 4, y/(b/2) = 0.63

.8 1.0

0 Basicwing

IX Full-span droop ("A")

+ Denotes lower surface

.o CXF.

)o ) o

.2 .4 .6

x/c

Station 2. y/(b/2) = 0.85

.8 1.0

C

P

-4

-2 ----_

0

2

4

-6

0 io

_ () O_

.2 .4 .6

xlc

Station5.yl(bl2)= 0.44

.8

3@@

1.0

_-vL) 0 0 (

.2 .4 .6

x/c

Station 3, y/(b/2} = 0.78

_o__ )-(B_-C

P

-4 ---

-2

E

0

2

4-.2

o o0 ) o

.8 1.0 0 .2 .4 .6 .8 1.0

x/c

Station 6. yl(b/2) = O.38

(a) _ = 11.4 ° •

Figure 36.- Chordwise pressure coefficients for wing with

modification A.

77

CP

C

P

Cp

-6

-4

-2

0

-6

-4 ---

-2

0

-6

-4

-2

0

Ix_ _o I 0 C 0

2

4-.2 0 .2 .4 .6

xlc

Station],y/(b/2)=0.92

0

8 1.0

o°O,I°to

4-°2 .2 .4 .6 .8

xlc

Station2,y/(b/2}= 0.85

1.0

r,

olAc_

4-+2 0 ,2 .4 .6 .8 1.0

x/c

Station3.yl(bl2)= 0.78

Cp

Cp

Cp

-8

-6

-4

-2

0

2

t_

)o o

__.__L __1_ .....,2 0 .2 .4 .6

x/c

Station4.y/(b/2)=0.63

J

--1

0 Basicwing

.8 1.0 _ Full-sNn droop ("A")

+ Denotes lower surface

-4----

_o(>, o () o <) o () ooi

2 ........ i ..........i

__ I4-.2 0:: .2 .4 .6 .8 1.0

xlc

Station5,yl(bl2)= 0.44

-6

-4

0 5x> o ocO_ __

2 .......................

4:.2 0 .2 .4 .6 .8 ].0

x/c

Station 6, yl(bl2} = O.38

(b) _ = 21.5 ° .

Figure 36.- Continued.

78

C

P

CP

Cp

-6

-4

-2

0

2

--0._ --

%0< o ) o

9

4

-.2 0 .2 .4 .6x/c

Station 1. yl(b/2) = 0.92

-8

-6--

-2

0_() 0 0 ( 0

02-.2 0 .2 .4 .6

xlc

Station2.y/(b/2)= 0.85

-6

-4

-2

0

2

(_)0( 0 0 0

4-._ 0 .Z .4 .6

xlc

Station3,yl(b/2)= 0.78

oO

,8 1.0

00('_,m,

.8 1.0

) oo

.8 l.O

-8

Cp

-6--

-4

-2

0

2-.2

h

o_ o o o

0 .2 .4 .6xlc

Station4.yl(b/2)= 0.63

.8 l.O

-6

-4

-2

Cp

O

2

_)o(

t_

__ 1

o ) o

-,2 0 12 .4 .6x/c

Station5,y/(bl'Z)= 0.44

-6

-4

-2

Cp

0

2--

c9o o4

4-.2 o .2 .4 .6

x/cStation 6, y/(b/2) = 0.38

(c) a = 31.7 ° •

Figure 36.- Continued.

O_

.8

8-

1.0

1.0

0 Basicwing

iX Full-span droop("A")

+ Denotes lower surface

79

cp

Cp

Cp

-6

-4

-2

0

2

4

-6

o ) o

.2 0 .2 .4 .6

x/c

Station ], y/(bf2) = 0.92

.8 1.0

-4

-2

0

2--

=.

4-.2

-6

0 .2 .4 .6 .8 I.Ox/c

Station 2, y/(b/'2)= 0.85

-4 --_

-2

0

2--

4.2

__ht

(_DCx o ) O P O (DO,.%--_,¢._,

.2 .4 .6

xlc

Station 3. y/(bfZ)= 0.78

.B l.O

C

p

Cp

cp

-6

-4

-2

0

2--

4..2 0

-6

0_ 0 0 (

(

.2 .4 .6

x/c

Station 4, y/(b/2) = 0.63

-4

o oo

.8 1.0

(d)

Figure

h-2

%0

%0 0 0 0 ) 0

b.

4-.2 0 .2 .4 .6

xlc

Station5,yl(b/2)= 0.44

-6

-4 --- E

r,-2 ----

0

2

4-.2 0 .2 .4 .6

xIc

Station6,yl(bl2)= 0.58

.8 1.0

O0O_

.8 1.0

a = 41.8 ° •

36.- Concluded.

o Basicwing

E, Full-span droop ("A")

÷ Denoteslower surface

80

cn

1.6

1.0

.8

.6

.4

! 4

.2 .4 .6 .8

Semi-spanwise station

Alpha 11.20

-I1.0

caConfigu ration

Basic wi ng

[] Original outboard droop("B")

O Original outboard droop +inboard fairing ("C")

A Full-span droop ("A")

1.8

1.2

1.0

.8

.6

.40

t J

b

.2 .4 .6 .8 1.0

Semi-spanwise station

Alpha 31.70

cn

1.6

1.4

1.2

1.0

.8

.6

.40

J

.2 .4 .6 ,8

Semi-spanwise station

Alpha 21.60

1.4

cn 1.2 ....

1.0

.8

_r )('

' _;

rJ

1/l__

s. '_-

.6

1.0 0 .2 .4 .6 .8

Semi-spanwise station

A Ipha 41,50

1.0

Figure 37.- Spanwise pressure distributions _or airplane with several

leading-edge-droop configurations.

81

cn

cn

1.6

1.4

1.2

1.0

.8

.6

_wI}./

•4d

1.6

1.4

1.2

1.0

.8

.6

.2 .4 .6 .8Semi-spanwisestation

Alpha 11.30

1.0

Configuration

0 Basic wing

[] Narrow-gap segmenteddroop ("F")

<_ Wide-gap segmenteddroop ("G")

1.8

1.6

1.4

1.2

cn

1.0

.8

/

.6

.40

Z)

.2 .4 .6 .8

Semi-spanwlse station

Alpha 31.60

[I/

%

II

_L

1.8

1.6

1.4

cn 1.2

1.0

.8

1.0

.40 .2 .4 .6 .8 1.0 .6 0 .2

Semi-spanwise station

Alpha 21.5

.4 .6 .8

Semi-spanwise station

Alpha 41.3°

Figure 38.- Spanwise pressure distributions for airplane with segmented

leading-edge droop.

-%

1.0

=

=

82

1.4

1.2

1.0 ........

ca

.6

.40

1.6

1.4

1.2

1.0

cn

.8

.6

.4

11.2 .4 .6 .8 1.0

Semi-spanwJsestation

Alpha 11.20

Configuration cn

© Basic wing

[] Outboard slat

Full-span slat+ fillet droop

0 .2 .4 .6 .8

Semi-spanwisestation

Alpha 31.70

1.0

.........

1.8 / EL

16.J ______

rcn 1.4 ................ / ---

1.0 C_8"/_ _9" -'_'_

.8

.60.2 .4 .6 .8 1.0 .2 .4 .6 .8

Semi-spanwise station Semi-spanwise station

Aipha 21.5° Alpha 41.40

Figure 39.- Spanwise pressure distributions for airplane with 0.08_

leading-edge slat.

1.0

83

cn

1.4

1.2

1.0

.8

.6

.4

0 Basic wing

[] Outboard slat

<_ Full-span slat+ fillet droop

.2 .4 .6 .8

Semi-spanwise station

Alpha 11.20

1.0

cn

2.0

1.8 _.

16 /1.4 -- -"_ "_

-S1.2 .....

1.0 ......

•8----t.- _.6

.40

I.2 .4 .6 .8

Semi-spanwise station

Alpha 31.70

2.4

i.0

cn

1.6

1.4

1.2

1.0

.8

.6

.4

0

A

..... -f _-_ cn

2.2

2.0

1.8

1.6

1.4

1.2

1.0

.8

.60

#

/

1 j

Jr-- w

9

.2 .4 .6 .8 1.0 .2 .4 .6 .8

Seml-spanwise station Semi-spanwise station

AIpha 21.50 Alpha 41.40

1.0

Figure 40.- Spanwise pressure distributions for airplane with 0.155

leading'edge slat.

84

.02 -

Confi gurati on

Original outboard droop ("B")

Original outboard droop + fairing ("C")

Full-span droop ("A")

Full-span droop ("A") with fillet droop

ACD

.01 - <

-.02 I I I I I I

0 .2 .4 .6 .8 1.0 1.2

CL

Figure 41.- Incremental values of drag coefficient for airplane with several

wing-leading-edge configurations.

85

Configuration

Wide-gap segmented droop ("G")

=-=-- _- Narrow-gap segmented droop ("F")

..... Original outboard droop ("B")

ACD

.02 --

.01 --

-,02 I I I I I J0 .2 .4 .6 .8 1.0 1.2

CL

Figure 42.- Values of incremental drag coefficient for airplane with segmented

leading-edge droop.

86

,/.-

,-,j ,,-,j0 c_,

r-

g_

I,

I I

i/

,"", _ '-"0

m

N

!

.._

r-I

I

I 0I:}'_.,-4

-,.-I,'_

_gr-t -N

O O

ID -,-4•,--I NO

oI

o

O _O

_1 .,-.I

_g

_ O

g-,.4

8V

'-I. o

o00<3

.. I

#

_;_ ...... ]J £/q LIE

E

_J

b

m

m

m

m

L

<

-< +

.... o

_+ - ! oo

---- -----i

¢ B

_-_--- ;1....... 1_

i

Ji

I

,(

/

---- u'_

o

¢,_

i

_ °

o

{:::

0

"0

"_,_

°_0 o

_'_

4..I

-,'4

g4 _0 _

I

g.,-t

i

{

i

88

C m

CL

CD

6e, deq

© 12

[] 0

0 -5

z_ -[0

[_ -15

-23

-]0 0 lO 20 _ 40 50 .6 .5 .4 .3 .2 ,! O -,] -,2 ",3 -.4 -.5 -_6 -.7 -,$

a, decj Cm

.2 ,4 .6 .8 1.0 1.2

CD

(a) C' = 0.T

Figure 45.- Effect of elevator deflection on longitudinal aerodynamic

characteristics of complete basic airplane.

89

Cm

CD

be, de_

0 12

[] G

0 -5

Z_ -10

-[5 :

1,2 m _-

,2 [L__

L

L_k

Lit .,d_i_-_

I it l_, , _ 1 __

-:;;z _,( l

-lO 0 lO 20 30 40 50 .7 ,G .5 ,4 .3 ._ l 0 -.l -,2 -.3 -.4 -.5 -.6 *,7 -,8 -,9 -.2 0 .2 .4 .6 .8 10 1.2

a. de_ Crn CO

(b) CT = 0.11.

Figure 45.- Concluded.

9O

C m

CL

C D

.2

.1

0--

%), --

-,2

-,3

-,7

_k_

C_ 5f, deg

©0 0

o O.ll 0

0 0.11 30

1.6

1.4

1.2

1.0

.8

.6

.4 P

.2

o

-10 0

-

t

IIIII

lO 20 30 40 .50 -.l -.2

a. de,9 Cm

.2 .1 0 -.3 -.4 -.5 -.6 -.7 -.2 0 .2 .4 .6 .8 ].0 1.2

CD

(a) Tail on.

Figure 46.- Effect of horizontal tail on longitudinal aerodynamic characteristics

of complete basic airplane.

91

Cm

- [] 0.11

-.3_ 0 0.11

1.6

1_,_.1.4

,.2 t

.4 - -

- 4

-10 0 10 20

o, deg

!

30 40 50 0 -.I

6 r (leg

O

0

30

-.2 0 .2 .4 .6 .8 1.0

CD

(b) Tail off.

Figure 46.- Concluded.

92

C m

C L

C D

.2

.I

0

-.I

-.3

-.4

-.6

-.7

CtF 6f. d_

O0 0

Q 0.II 0

0 0.II )0

1.61"S_] )_.,.'=-_ .a,-__},.

.8 _ i

.6 I_,_ _±j_

, ;g.2

°r: ;Ill-.4:1I I I ....... 72_-____-.6 ]IIIII

-lO 0 1O 20 _) 40 50 .2 .l 0 -.1 2.2 -.) -.4 -,5 -.6 -3 -.8

o, de9 Cm

i!_ _

-.2 O .2 .4 .6 .8 1.0 1.2 1,4

C D

Figure 47.- Longitudinal characteristics of complete airplane with modification B

wing. (See fig. 5(b).)

93

2 II I

0

1.[

-2

-.5 m_ iL

-.8 llll _

C_ 6f. de9

00 0

00.11 0

00.11 3O

1.0

CL

.8

C D

.6

/

z

0 10 20 30 40 50 .2 ,1 0 -.1 -.2 -.3 -.4 -5 -,6 -,7 -,8 -.2 0 .2 .4 .6 ,8 1.0 1.2 1.4

a. d_ Cm Co

Figure 48.- Longitudinal aerodynamic characteristics of complete airplane with

modification A wing. (See fig. 5(a).)

94

c_ 6f. deg

OO o

[3 o.11 o

<> O.ll 3o

c L

c D

2.D_

7 " •

L.8_-

1.6 -

l.d

1.2

l.O

i.1

,o

-.4

-.6

-lO 0

l I

I[

LEI_I0 20 30 40 50

o, deg

,2 .I 0 -.I -.2 -.3 -.4 -.5 -.6 -.7 -.8 -.2 0 .2 .4 .6 ,8

CoCm

..... t

t

L--L

LL___1.o 1.2 1.4

Figure 49.- Longitudinal aerodynamic characteristics of complete airplane with

modification A wing and fillet droop. (See figs. 5(a) and 12(d).)

95

•02

0

C{ -.02

-.04

LJlk_

_L

!LF

k__

LI

L

L

!-

LL

I; +_.

t_f

!-

- + ,

[i11

1 I

L_t •

i

iI

i .

11

'I

!

J

10

4

FF t

/

!

-. 06

-10 0 20 30

a,deg

Tunnel scales

O Averageof readings[] Minimum reading

Maximum reading

40 50

Figure 50.- Rolling moment on complete airplane showing fluctuations at

high angles of attack.

96

.O2

Cy6 0

-.02

v

.004

-.002

Ct13

• OO6

• 004

• 002

L

F[\I1'I_ _>II1\I1¢,_

°' _ / "_,J-. 004

-.006

-lO

%11

0 lO 20 30 40 50

c_, deg

(a) Tail on.

Figure 51.- E6fect of vertical tail on static-lateral-stability

characteristics of complete basic airplane.

97

CYB

Cn_

C

IB

.02

I

o,__ _ m__ _ ____ __-.02

-.002

.004

• 002

-.004

.......... L ......

C_- 6f, deg

0 o o

[] 0.11 0

<_ O.ll 30

:-__ ___:-_ <...

! i

• 0O4

.002

-.002

-.OO4

-.006

-.008

-10 10

/I_111,_I1_]>II_K,Iil C_III \_\ ,_

L IV '_

°120 30 40

z

50

c_, deg

(b) Tail off.

Figure 51.- Concluded.

98

.O2

Cy_

CnB

-.02

• OO4

.002

0

-. 002

-. 004

6f, deg

0 o

[] o<> 30

IT

0

0.II

0.II

C

IB

-. 002

-. 004

-. 006

-. 008

0 0 I0 20 30 40 50

_, deg

Figure 52.- Lateral-directional characteristics of complete airplane

with modification B wing. (See fig. 5(b).)

99

CY 8

CnB

.02

•008

.006

•0O4

•002

-. 002

-. 004

/\_

/ .o,

I

-,.-4'I,

J

6f, deg C_

0 o o

[] o o.11

_30 O.ll

.... Z

C

tB

•006

.004

.002

oo_ .

-. 004

-.006

-. 008

-I0

15.

/1,\

!fk

/s_il// '.I,I _-V_ _>_I ]-_-

0 lO 20 30 40 50

ct, deg

Figure 53.- Lateral-directional characteristics of complete airplane

with modification A wing. (See fig. 5(a).)

100

Cy_

.O2

- .02

6f, deg

0 o[] o<_ 3o

c÷0

0.II

0.II

• OO4

Cn_

•002

0

-. 002

-. 004

-. 002

-. 006

v-;W

1 _ =,,--_k a

-. 008-I0 0 I0 20 30 40 50

e, deg

Figure 54.- Lateral-directional characteristics of complete airplane

with modification A wing and fillet droop. (See figs. 5(a) and

12(d).)

101

ACy

o-. i

6r, deg

0 -z5[] Z5

0 o

ACn

•O4

.o2 ...e _e _e

o ,...._£I_._,_,.

-.02

-. 04

i

ACl

.02

-.02

-. 04

-lO

_/ ..... _Z_ .

0 lO 20 50

, !

oi I30 40

m, deg

(a) C' = 0.T

Figure 55.- Effect of radder deflection on lateral-directional

aerodynamic characteristics of complete basic airplane.

102

ACy

.I

O'

-.I

-.2

6r, deg

0 -25

[] 25

o

• O4

.O2

._._)--s_

\

AC -.02 .......

-- -- r

-. 04

-. 06 --

-. 08

ACI

•O4

.02

-.02-I0 0

\W =10 20 30 40

m, deg

(b) C' = 0.11.T

Figure 55.- Concluded.

5O

103

Wing configuration

O Basic

[] Original outboard droop ("B")

Original outboard droop + inboard fairing ("C")

/_ Full-span droop ("A")

ACy

.I

-.l

_Cn

.O2

-.02

ACI

• O6

.O4

.O2

0

-I0 0 I0 20 30 40 50

_, deg

Figure 56.- Effect of aileron deflection on complete airplane with various

wing configurations. Aileron deflection = 25 ° for right roll.

104

I. Report No. 2. Government Accession No.

NASA TP-20 1 1

4. Title and Subtitle

EFFECTS OF WING-LEADING-EDGE MODIFICATIONS ON A F[TLL-SCALE,

LOW-WING GENERAL AVIATION AIRPLANE - WIND-TUNNEL INVESTIGA-

TION OF HIGH-ANGLE-OF-ATTACK AERODYNAMIC CHARACTERISTICS

7 Author(s)

William A. Newsom, Jr., Dale R. Satran,

and Joseph L. Johnson, Jr.

9. Performing Organization Name and Address

NASA Langley Research Center

Hampton, VA 23665

12. Sponsoring Agency Nameand Address

National Aeronautics and Space Administration

Washington, DC 20546

3. Recipient's Catalog No,

5. Report Date

June 1982

6. Pedorming Or_nizationCode

505-41-13-02

8. Performing Or_n_zation Re_rt No,

L-15101

10. Work Unit No.

11. Contract or Grant No.

13. Type of Report and Period Covered

Technical Paper

14. Sponsoring Agency Code

15, S_pplementary Notes

16. Abstract

The tests were made in the Langley 30- by 60-Foot Tunnel. Wing-leading-edge modi-

fications included leading-edge droop and slat configurations having full-span,

partial-span, or segmented arrangements. Other devices included wing-chord

extensions, fences, and leading-edge stall strips. Good correlation was apparent

between the results of wind-tunnel data and the results of flight tests, on the

basis of autorotational stability criterion, for a wide range of wing-leading-edge

modifications.

17. Key Words(Suggested by Author(s))

Spin resistance

Stall resistance

Wing-leading-edge modifications

General aviation

19. Security _assif.(ofthisreport) 20. SecurityClassif.(ofthis _ge)

Unclassified Unclassified

18. Distribution Statement

Unclassified - Unlimited

Subject Category 01

21. No. of Pages 22. Dice

105 A06

.-_os For sale by the National Technical Information Service, Springfield, Virginia 22161 NASA-Lang]ey. ]982