c o p y No. b

RESEARCH MEMORANDUM HIGH-SPEED AERODYNAMIC CHARACTERISTICS O F A

- . -. - . ." - _ _ .

LATERAL-CONTROL MODEL. I - NACA 0012-64

SECTION WITH 20-PERCENT-CHORD PLAIN

AILERON AND Oo AND 45O SWEEPEUCK

By Joseph L. Anderson and Walter J. Krumm

A m e s Aeronautical Laboratory,

CLASSIFICATION CANCELLEd Moffett Field Calif.

NATIONAL ADVISORY COMMITTEE . FOR AERONAUTICS

WASHINGTON September 27,1948

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NATIONAL ADVISORY COMMITTEE: FOR AERONAUTICS

By Joseph L. Anderson and Walter J. Eumn

Wind-tunnel tests were made t o determine the aerodynamic chezacteristics of a -ercent-chord plain aileron on a semispan w i n g having the NACA OOlS64 section. This report includes the results of t e s t s of the wing unswept and swept back 450 and w i t h the aileron deflected from Oo t o loo. Data were obtained for a Mach nuniber range from 0.40 t o 0.85. I

The results with the w i n g w e p t indicated an aileron eve?

balance and reversal of effectiveness at high speeds. With the wing swept back 45O the aileron exhibited only an overbalance and no reversal of effectiveness.

Work has been done cancerning adequate longitudinal control of airplanes at high speeds, but development of adequate la teral control has not kept pace. Ln order t o investigate the problems of lateral control ebnd span loading and t o develop ailerons which are effective at high speeds, a semispan wing model was built d tes ts of the mdelwere made in the h s l&foot h i m p e e d wind tunnel.

The model -8 f i t t ed with a leading-sage aileron and a trailing-edge aileron, This report includes the results of tes ts of the w i n g unswept and swept back 450 with a t r a i l w d g e aileron.

ilNCLASSlFlED

2 - NACA RM No. A8812

The coefficients and symbols used in this report are defined as follows:

where

D

E

L

L'

"

S

v 8

b

aspect ra t io .

drag coefficient

lift coefficient

aileron hinge-mnment coefficient

drag of semispan model, pounds

aileron hinge moment, f-

lift of semispan model, pound8

roUing moment about longitudinal axis parallel t o the air stream in plane of spmetry, foot"pounds

pitching naoment about the 0.25 E of semispan model, foot- PO-

area of semispan d e l , square feet

velocity of the free air stream, feet per second

speed of sound i n the free air stream, feet per second

semLspan of model, feet

b

"r

t

i

+. .

aileron span parallel t o the hinge

3

line, feet

chord of w i n g perpendicular t o 0.25 chord line, feet

root+nesll-square chord of aileron perpendicular to the hinge line, feet

section normal force, pounds per rout

m c pressure ($PI, pounds per squsre foot

angle of attack of d e l , degrees

t r a i l w d g e aileron deflection relative t o airfoil, positive when the trail ing edge is deflected dcnmuaxd, degrees

The model (fig. 1) used for these tests waa a semispan tapered wing with the XACA OOl2-64 aection (table I). For the wing unswept the 0.25 w-hord line was perpendicular t o the air stream. For the wing swept back, the wing was rotated about the 0.50 root-chord point so that the 0.25 wing-~hord l ine w a s 45O t o the air stream. The mdel dimensions for the wing unswept and swept back 45O are given in table 11. The wing was f it.ted w i t h a l-rcent chord, lead-dge, unsealed aileron which w a s not deflected f o r this investigation, 8r.d a eercent-chord, tra- aileron. The t r a i l w g e aileron extended. in span from 0 -561 of the semispan of the unsxept wing t o the t i p and had an unsealed radius nose. This aileron wet8 deflected from Oo to loo f o r this investigation. The aileron hinge momnts were masurd w i t h resistance-type electrical strain gages.

The model w a s mounted in the tunnel with the wing spar extending through the tunnel wall and fastened t o the balance frame. A baffle was installed & the model near the tunnel wall to direct the leakage fluw through the turmel wall axag *om the surface of the model. (See fig. 2. )

Chordwise raws of pressure orifices were located perpendicular to the 0.25 w i n g - c h o r d line at six spgnwise statians: 0.179, 0.417, 0 381, 0.724, 0.867, and 0.935 of the semispan of the unexept wing; so that chordwise and spmwise pressure distributions could be measured.

4

CORR?ICTIONS To aA!L!A

NACA RM No. A8H12

The data were corrected f o r the blockage of the tunnel air stream by the model. The angle of attack asd the drag and rolling- moment coefficients were corrected for the effect of the tunnel u8lls by the mthod of reference 1. As 8 basis for these corrections, the span loading on the w i n g at 0.80 Mach llupiber was used. The corrections w e r e applied t o the data as follows:

. . -. ..

where the corrections are given in the followfng table:

Correctfons

0.51 CL 0.41 CL Ax, degrees

W b g swept back U i n g unswept

ACD .0076 cL2 .0064 cL=

K .g10 =902

The subscript u denotes the uncorrected data. No corrections were made f o r the effect of the tunnel41 boundary layer passing over the model.

RFSUGTS AND DISCUSSION

The test Reynolds numbers f o r the w i n g unswept and swept back 45' are shown in figure 3. For the wing unswept and swept back, and with the aileron deflected, the variation of lift coeffi- cient with angle of attack, the variation of drag coefficient w i t h Mach number, and the variation of pitch-ment coefficient with l i f t coefficient are presented in figures 4, 5 , and 6. The Mach nuTliber of lift and drag divergence for the wing unswept is about 0.80.

The increment of rolling-moment coefficient produced by deflection of the trailwdge aileron for the wing unswept and swept back 45* is presented in figure 7. For the w i n g unswept,

NACA RM KO. A m 0 5

a t most q l e s of attack there is a decrease in aileron effectiveness as the Mach number is increased from 0.60 t o oi85, and a t 0.85 Mach number f o r -bo apd -2O angle of attack a positive deflection of the aileron produces negative increments of rollhg-momnt coefficient. (See fig. 7(a). ) '&e effectiveness is, i n general, reduced for the swept wing.

Figure 8 shows the varia.tion of trailing-edge aileron hinge- mment coefficient with aileron deflection for the wing unswept and swept back 45O. A t 0.70 Mach nuniber for the wing unswept, the aileron becomss overbalanced at negative angles of attack ami the aileron is overbalanced over a greater angl-f-attack range as the Mach nuniber is increased. The aileron, f o r the wing swept back, I s overbalanced a t high Mach numbers but not t o as great an exbent &B for the w i n g -wept. Reference 2 indicates that a trail- edge angle greater than approximately 14O causes a control surface t o overbalance and reverse its effectiveness a t high subsonic Mach numbers. The trailfng-edge angle parallel t o the a i r stream of the wing unswept is 20.63O, and this lazge angle might be the cause of the aileron overbalance and effectivenesa reversal. Sweeping the w i n g hack reduced the effect of the h r g e trail- edge angle.

The spanwise VEtriatloli of section no"4orce coefficient for the w i n g unswept asd swept back i s presented in figures 9 esd 10, respectively. These data were obtained from integration of the chordwise pressure distribution ( n o d t o the 0.25-wing chord) a t each of six spanwise stations. For the w i n g unswept , the loss of l i f t on the w f n g a t about 55 percent of the wing semispas is probably the result of leakage around the Inboard ends of the ailerons. For the w i n g swept back, there is no indication of this loss of lift.

A m s Aeronautical Laboratory, National Advisory C o d t t e e fo r Aeronautics,

Moffett Field, Calif.

1. Sivells, James C., and Deters, Owen J.: Jet-Boundary and Plan-Form Corrections f o r PartialSpan Models w i t h Reflection Plane, End Plate, o r No End Plate in a Closed Circular W i n d Tunnel. NACA TEI No. 1077,1946,

6 L NACA RM lTo. A m 2

2. Axelson, John A.: A Sumnvtry anti Analysis of Wind”unne1 Ihta on the Lift and Hinge-Moment Characteristics of Control Surfaoes up to a Mach Nuniber of 0 . 9 . WACA RM No. A W 2 , 1947 -

NACA RM No. Am12 .k

TABU I. - COORDINATES W PERCENT CHW

FOR mACA 0012-64 SECTION

Stat ion

0 .oo 1.25 2.50 5.00 7.50 10.00 15.00 20.00 30.00 40 .OO 5 0 . ~ 60.00 70.00 80.00 %).oo 95 000 loo. 00

Ordinate

0.00 1.81 2.45 3.27 3.81 4.24 4.87 5-29 5.83 6.00 5.83 5.32 4.48 3.32 1.87 1.03 0.12

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8 - NACA RM No. Am12

TABLE: II. - MODEL DIMENSIONS

Dimension

Semispan area, square'feet

Semispan, f ee t

Aspect r a t i o (based on full span)

Taper ratio

Sweepback of lead- edge, degrees

Sweepback of 0.25 chord line, degreeg

Sweepback of t r a i l i n g edge, degrees

Mean aerodynamic chord, feet

Root chord, f e e t

Projected tip chord, feet

Leading-edge aileron:

Ratio of a i leron chord to wing. chord (perpendicular to quartex- chord lfne)

Span along hinge-line, feet

Root-rnewquare chord of aileron, feet

T r a i l w d g e aileron:

Ratio of ai leron chord to w i n g chord (perpendicular t o quarte- chord line)

Span along hinge line, feet

Roo-knea-quare chord of aileron, feet -

12.25

7.00

8.00

0.50

2.415

0

-7.206

1.841

2.360

1.180

0.15

2.445

0.224

0.20

2 998

0.312

12.35

5.327

4.59

0.484

47.415

45

37 794

2.513

3.224

1 559

0.15

2.445

0.224

0.20

2 998

0.312 ?

MCA RM Ho. A&%X? - I

I I NAGA. 001'2-64 A/ifo/;l section

9

i

.

i

NACA RM No. Am12

Figure 2. - Unswept wing mounted in the h e 16-f o o t high-speed w i n d tunnel.

4

4

t

NACA & No. A m 2

9

8

4

3 3 *4 .5 .6 .I .6 .9

Much number, M

14 NACA RM N Z Am12

LP

.8

0

(a) Wing unswepf

figure 4. - Variafion of ## coefficient with angle of am&. Aileron undeflecfed!

NACA RM No. A8Hl.2 15

/.2

1.0

. .8

.6

G

0

16 - NACA RM No. Am12

.02

0 .3 .4 .5 .6 .I .8 .9

Much number, M (u) Hfm unswept

Figurn 5. - Vuriuth of drug coeflicient with Much number

for consfant ung/es of uttuck Atleron undeflected.

IUCARMNo. A m

.8

4

0

(a) Wing unswept.

Figure 6. - Vuriaiion of piiching-mmenf coefficienf Wh?i /iff coefficient. Aileron undeflected.

18

f l o f 4 c of b

(b) Wing swept bock 4.5"

Figure 6. - Conc/udeo!

4

NACA RM No. A8812 - .02

.Ol

e a U

8 0 4 8

, 02

.O/

0

YO/ 0 ' 4 8 Aileron deflecfion, deg

(a] Wing unswepf "297 Figure ir - tbriafbn of incremenf of rolling-mrnenf

coeflcienf wifh froi/ing - m e u/lemn deffecfhn. .

20 I NACA RM No. A8812

.02

.O/

4 0 0 0 4 8 0

.O/

0

4 8

0 4 8 Ai/eron deflecfion, 6oldep Tj$G5J7

(b) Wing swept buck 4 5 O

.

f

..

WCA RM No. A8H12 - 21

.02 0

-.04

-.06

0 2 4 6 8 l O

.02

0

-.02

-.04

706

-.08

-fO 0 2 4 6 8 l O Aileron deflecfion, Go, deg

(a) Wing unswept figure 8: - Variafion of fruiling - edge aileron hinge -momen f

wefficient with aileron angle.

22 _,. .. ." . - NACA RM No. A m 2

. 01 0

701

-.02

-03

.02

e= .o/

C e .e P

0 2 4 6 8 / 0

.02

.U/

0

-Q/

-02 - n z

0 2 4 6 8 / 0

.wtJ 0 2 4 6 8 / 0 Ai lem deflection, 6*, deg

. . -. - - .

a -4"

* - 2 O 0" 2" 4 O

6 * 8"

" . .

" "- "- 1""""" "- """

L ..

NACA RM No. A m 2

.6

.4

.2

figure 9. - s;Ormwise vuriufion of section normal-

force coeflicimt Wing unswepf. . i

24

.6

.4

.2

0

.2

0

0

NACA RM No. Am12

L

72

-.4

8

r

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0 2* 4" """"""" "

6" -- - - /O*""" .P

0

tP

74

26

4 7 O0 20""- 40 """"""

e

.4

.P

0

NACA RM No. A&€U2

74 0 20 40 623 80 /oo

Pefcenf semispun (ol M, QBO

Figure 10.- Spanwise vanohon of secfion norm/-

hme coeflicient Wing swepf back 45:

NACA RM No. A m

.4

.2

n

20""- 40 """"""

-.7

0 €0 40 60 80 /uo Percent semispun

28

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0

3 20"""- 4%"""""""" 6" -

"

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