RM A55AO3

RESEARCH MEMORANDUM

EFFECTS OF SWEEP AND TAPER R.ATIO ON TEE LONGITUDINAL

CHARACTERISTICS OF AN ASPECT RATIO 3 WINGBODY

COMBINATION AT MACH NUMBERS FROM 0.6 TO 1.4

By Earl D. Knechtel and James L. Summers

Ames Aeronautical Laboratory Moffett Field, Calif.

FOR AERONAUTICS :-i .xk . . WASHINGTON z :.

March 23, 1955 .* _

NACA RM A55AO3

NATIONAL ADVISORY WMMITTEE FOR AERORMJTICS

EFFECTS OF SWEEPANDTAPERHATIO ONTHELONGITUDINAL

CHAHACTERISTICS OF AN ASPECT RATIO 3 WING-BODY

COMBINATION AT MACH NUNBEES FROM 0.6 TO 1.4

By Earl D. Knechtel and James L. Summers

SUMMARY

An experimental investigation was conducted to assess the effects of sweep and taper ratio on the longitudinal characteristics of a wing- body combination at Mach number.6 from 0.6 to 1.4 for a Reynolds number of s.gxlo% The wings were 3 percent thick and of aspect ratio 3. Leading-edge sweep was varied from 19.1° to 53.1° at a constant taper ratio of.0.4, and taper ratio w-as varied from 0 to 0.4 at a constant sweep angle of 53.1°. \

kcreased leading-edge sweep cau‘sed a progressive decrease not only of lift-curve slope, but also of the variation wLth Mach number of lift- curve slope and static longitudinal stability. Throughout the test Mach number range the minimum drag coefficient was less for 53.1° of sweep than for 19.1°. The drag-rise factor was generally larger and showed less variation with Mach number for the model hating 53.1° of sweep. The maxiam lift-drag ratio was higher for the model having 53.1' of sweep at supersonic Mach numbers.

Although the effects of taper ratio were less pronounced than those of sweep, the results of

g regressive increases in taper ratio, at a con-

stant sweep angle of 53.1 , indicated that the model with 0.4 taper ratio had the least variation of lift-curve slope with Mach number for the three taper ratios investigated. Over-all change in static longitudinal stabil- ity was generally least for 0 taper ratio at 0 lift, and least for 0.4 taper ratio at 0.4 lift coefficient. The 0.4 taper ratio also resulted in a lower miniunun drag coefficient and a higher maximum lift-drag ratio at Mach numbers above 0.95 than for the 0 taper ratio. In addition, the drag-rise factor is generally less and varies less with Mach number for the model having 0.4 taper ratio.

2 NACA RMA55A03

The present investigation is part of a continuing program (refs. 1 and 2) directed toward assessing-the effects of plan-form variations on the longitudinal characteristics of wing-body combinations at trsnsonic Mach numbers. The study reported herein was conducted in the Ames 2- by 2-foot transonic wind tunnel to determine.the.l.ongitudinal characteristics of five wing-body combinations having thin wings of aspect ratio 3 which embody systematic variations of sweep angle and taper ratio.

_ - *-

..d. d -. . -

,- The results presented here have the twofold purpose of indicating the

effects on the longitudinal characteristics of-tie models due to (1) vari- -- 'Y ation of sweep at a constant moderate taper ratio, and (2) variation of . .-

taper ratio at a constant large angle of-leading-edge &e&i. Some of the results are compared tith those of available theory.

CD

Gin

CL

cLcL

Cm

C

E

Wn E

.

NOTATION

drag coefficient

minimum drag coefficient

lift CObfficient

- ---

--

Y

.I I . ..T -.-

.a

lift-curve slope

pitching-moment coefficient referred to.guart&chord point of mean aerodynamic chord

L local chord

meanaerodynamic chord

slope of pitching-moment curve

change in pitching-moment-curve slope due to chenge in

maximum lift-drag ratio

body length, including portion removed to accommodate balance

NACA FM A55A03 3

M free-stream Mach number

r local radius of body

r0 maximum radius of body

X body longitudinal coordinate, measured from body nose

a angle of attack, deg

A

h

sweepback angle of leading edge, deg

tip chord tax= ratio, root chord

APPARATUS, TESTS, AND DATA FGZIXJCTION

The investigation was conducted in the Ames 2- by 2-foot transonic HFnd tunnel, which is of the closed-circuit, variable-pressure type having .maximum design operating conditions of 45 pounds per square inch absolute stagnation pressure and 120' F stagnation temperature. Thewindtunnel is fitted with a flexible nozzle followed by a ventilated test section (fig. 1) which permits contLnuous choke-free operation from 0 to 1.4 Mach number.

Five wing-body models of steel were constructed such that sweep angle and taper ratio could be investigated independently for a constant aspect ratio of 3 (fig. 2). For three of the configurations the taper ratio was 0.4 and the leading-edge sweep angles, lg.l", 45O, and 53.1°, while for the other two configurations the sweep angle was 53 .l" and the taper ratios, 0 ma 0.2. For the four wings having 45O and 53.1' of sweep, NACA 0003 airfoils were employed in the streamwise direction; whereas for the wing with lg.l" of.sweep, biconvex sections 3 percent thitzk were uti- lized. The choice of profile for this latter case, for reasons discussed in reference 2, was based upon the known favorable drag cheracteristics of sharp leading edges for wings ha-&q leading edges supersonic over .most of the supersonic Mach number range.

Themodels weremounted inthewindtunnelona sting-supportedinter- nal strain-gage balance as shown in figures 1 and 3. The .models spanned approximately 45 percent of the test section height and blocked approxl- mately 0.5 percent of the cross-sectional area of the test section.

Lift and pitching moment were,measured for all five models at angles of attack from -4' to approximately 13' at Mach numbers from 0.60 to 1.40. Drag was measured over the same ranges of angle of attack and Mach number for only three ,models. These were the configurations having 0.4 taper ratio with 19.1° and 53.1' of sweep, and the configuration hating 0 taper

4 1 NACA RM A55A03

ratio with 53.1° of sweep. . .I

A Reynolds number of 1.5 mig.ion, based on 1 the mean aerodynamic chord of each model, was held constant for the tests. Y All coefficients w-ere based on the wing area fncl&Lng the portion within c:- the body. The pitching-moment coefficient was baaed on the mean aerody- namic chord and referred to the qusrter-chord point. The measured drx

-. ..-

was adjusted to correspond to a condition bf free-stream static pressure acting at the model base.

L .- .I 7-y

Subsonic waLi interference corrections _...? .calculated on the bas.is o-f. -. the theory of reference 3, were found.;to be mall and th-mefore were not amlied to the data;-- No corrections were made for..possible wall inter- ference due to reflected waves at low supersonic speeds. These effects. are dIscussed in theResults snd Discussion section. Corrections for air- stream angularity were not made, since they were found to be less thanthe probable errors in measuring angi&of atta&;Y Drag corrections due to longitudinal pressure gradient were unnecessary throughout the test I&L&I number range, since local Mach number detiations in the vicFnity of the .model were generally no greater than b.003. The data have not been corrected for aeroelastic distortion.

-

- .-

_ - .---

- ; -

Apart from the small systematic errors arising due to neglecting the ._I corrections discussed above, certain random errors of measurement exist -7 ._

..A which determine the precision, or repeatabillty, of the data. hanalyais .- y: was made of the precision of the Mxh number;.angle of attack, and coefff- w .- cients of lift, pitching moment, and drag for the models of the present investigation, and the random uncertainties at three representative Mach numbers and two values of lift coefficient are presented in the following table:

M = 0.8 M = 1.0 M = 1.2 .

CL = 0 CL = 0.4 CL = 0 CL = 0.4 CL = 0 CL = 0.4

M kO.003 kO.003 40.004 f0.004 f.020

c"L *.cm5 *.03o k.020 f.030 *y*g

f0.002 f.030

f.O1O LOO5 ~006 Gcl3 *.006 c, s.004 ~006 *.004 *.005 +iOO3 LOO5 CD *.0O03 f. 0010 rt.0003 k.0006 Loo03 +.0006

4

NACA RM A55A03

RZSKU!SANDDISCUSSION

5

. 3.1 figures 4and 5 are shown, respectively, the variations of Xft

, coefficient with angle of attack and with pitching-moment coefficient at Mach numbers from 0.60 to 1.40 for the five configurations tested. Drag results are presented in figure 6 f_or three of the configurations. The variations with Mach number of lift-curve slope and pitching-moment-curve slope for three values of lift coefficient are shown for the five config- urations in figure 7. The small irregularities which eppear in these curves at low supersonic speeds are believed to be the result of shock waves from the model which reflect from the tunnel walls and impinge upon the afterportion of the model. However, the influence of these reflected waves was confined to the Nach number range from 1.00 to 1.15 and was con- sidered not large enough to affect the conclusions drawn from the data.

Variations of calculated lift-curve slope and pitching-moment-curve slope (whenever possible) with Mach number are shown in figure 7 for the zero-lift condition. These theoretical values include effects of wing- body Interference as computed by the method of reference 4, which, in turn, is based upon theoretical wing-alone JZt characteristics obtained from references 5, 6, and 7, respectively, for subsonic, sonic, and supersonic speeds. The experimental and calculated lift-curve slopes agreed within approtimately 15 percent for all except the model having 53.1° sweep and 0.4 taper ratio. The poorer agreement in this case should not be surpris- ing in view of the limitations of the interference theory with regard to swept trailing edges, ma Ln view of the greater effects of aeroelastic distortion upon the more hfghly swept wing. In the two cases for which theoretical pitchkng~moment-curve slopes could be calculated, quautative agreement between theory and experiment was noted for the model having the least sweep, whereas good a-eement was _obtained for the triangular- wing model. A consistent aiscrepancy between theory and experiment is noted in the &ch numbers at which peaks occur in the lift-curve slopes and pitching+I.oment-curve slopes. Peaks in the calculated curves occur at Mach numbers higher than those of the experimental curves by amounts vsryTng from approximately 0.10 for the relatively unswept model down to approximately 0.03 for the triangular-wing model. This discrepancy probably results from the inability of the present linear theories to account for the fact that a local field of sonic and supersonic flow develops near the 'wing prior to the establishment the free stream.

Effect of Sweep

of these conditions In

m-t and pitcm -moment characteristics.- EI figure 8 a comparison of the ILPt-curve slopes obtained for the three sweep angles reveals no unusual trends. With increased sweep, the lift-curve slope not only became amaller, but also varied less rapidly with Mach number.

6 IWCA RM A55A03

In figure 8 for the three sweep angles investigated a cotupsxfson is shown of the change in static longitudinal stability from that obtatied at the same lift coefficient at 0.6 Mach number. Comparison on this basis t makes evident the large effects of sweep on the-variation of longitudinal . ' stability with Mach number. At high subsonic Mach numbers the .model hav- ing 19.1° of sweep underwent rapid ch&ges in static longitudinal-stabil-

____

ity; whereas the models having greater sweep generally exhibited a smother variation of stability with Mach number. The generally superior stability characteristics at lift coefficients up to 0.4 for 53.1’ of sweep may be somewhat offset, however, by the pitch-up tendencies at larger lift coeffi- cients in the high subsonic speed range. (See fig. p(c).>

Drag characteristics.- In figure 9 the variations ti-th Mach number of minlmum drag coefficient, drag-rise factor, andmaximumlift-drag ratio are compared for sweep pglks of 19.1° and 53;l". Throughout the test .- Mach number range the minimum drag coefficient was less for the wing hav- ,, ing the larger sweep angle, the amount of tUs.&Wference being as great as 40 percent at son5.c speed.

Drag-rise factor wss determined by the slope of curves of drag coeffi- cient plotted against lift coefficient squared, over the linear range of these curves from 0 to 0.4 lift coefficient. The drag-rise factor was

* I-

generally larger and varied less with Mach number for the model having the -yT greater sweep angle. *

Xaximum lift-drag ratios were attained at LWt coefficients from 0.20 t0 0.25. The variation of .=ximum lift-drag ratio with Mach number shown in figure 9 generally reflects the corresponding variations of .minimum drag coefficient and drag-rise factor. Increase in sweep angle from 19.1' to 53.1°~resKLted in slight decreases ti maximum lift-drag ratio at Mach numbers from 0.75 to 0.95 and increases of as much as 15 percent at supersonic Mach numbers.

Effect of Taper Ratio

Lift and pitching-moment characteristics.- A coqarison is shown in -7 -- figure 10 of the slopes of the lift and moment curves for the three taper

ratios. Although the effect of taper was not large, the variation with .- Mach number of lift-curve slope was generally least for the configuration ._z having 0.4 taper ratio. In general, the over-all change in static longi- ~, tudinal stability was least for 0 taper ratio at 0 lift coefficient and least for 0.4 taper ratio at 0.4 lift coefficient.

--.. Drag characteristics.- The variation8 with Mach number of m-lnFrmrm

drag coefficient, drag-rise factor, andmaximumJ.fft-dragratioare shown in figure 11 for the swept wings hating taper ratios of 0 and 0.4. These results indkate that up to 0.95 Mach number the minimum drag coefficients -

NACA RM A55AO3

and maximum lift-drag ratios for the two models coaq?are quite closely. At Mach numbers above 0.95, however, lower tiimum drag coefficient8 and higher .maximum lift-drag ratios sre realized for the 0.4 taper ratio. The drag-rise factor for the model having 0.4 taper ratio is generally less and varies less with Mach number than for the model'with 0 taper ratio.

Comparison with Results of Reference 1

The models for which longitudinal characteristics were presented in reference L had an unswept midchord line, %LCA 64Am3 StramWiSe SeCtiOnS, and taper ratios varying from 1.0 to 0. Accordingly, the leading-edge sweep angle varied from O" to 33.7’, and it may be of interest to compare the effect of the sweep v&iations of reference 1 tith the corresponding effect indicated in the present report. Such data comparisons are made in figures 9 and 2.2 for two configurations of reference 1 and the most nearly comparable models of the present tivestigation for which corre- sponding data were obtained.

In figure 12 the variations with Mach number of lift-curve slope and pitching-moment-curve slope are shown for two models of the present inves- tigation having 19.1° and 45.0° of sweep and two models of reference 1 having 12.60 and 33.7’ of sweep. The effect of leading-edge sweep was quite similar in the two cases, the variations with Bach number of both lift-curve slope and pitching-moment-curve slope generally becoming less abrupt with increasing sweep angle.

331 figure 9 the variations with Mach number of minimum drag coeffi- cient, drag-rise factor, and maximum lift-drag ratio for the same two models of reference 1 are compared with the corresponding results for the models of the present tivestigation having a taper ratio of 0.4 and leading-edge sweep angles of lg.l" and 53.1'. The difference in maxLmum lift-drag ratios at subsonic %ch numbers between the two models having the least sweep is attributable to the use of a sharp leading-edge, and a COnSeqUeniJ loss of leading-edge suction, for the model having 19.1' of sweep. Generally, however, the effect of leading-edge sweep on the drag characteristics, as on the slopes of the lift and pitching-moment curves, was much the same for the models of reference 1 as for the models of the present investigation. These comparisons suggest the possibility that the effects in reference 1 attributed to taper ratio might have been due in part to leading-edge sweep.

.

8 NACA RM A55A03

COIK!LUExCMS . -.

An eqp?imental investigatfon was conducted to determine the effects of sweep and taper ratio on the lift, pitchin@;- moment, and drag chmacter- iStiC of an aspect ratio 3 wing-body combination at Mach numbers from 0.6 to 1.4, at l.5x106 Reynolds number. leading-edge sweep angles of lg.l', 45O,

~ReStit8 obtainedfor ~i12gS having

for wiZlg8 having taper ratios of 0, and 53.1° at 0.4 taper ratio and

to the following conclusions: 0.2, and 0.4 at 53.1' of sweep lead

1. Increased awe-ep re8ulted in progressive decrease8 in lift-curve slope as well. as in-reduced variation with Mach number of'the lift-curve slope and static longitudinal Stability.

- -.-

2. An -ease of sweep Prom 19.1' to 53.1' led to generally decreased minimum drag coefffcient and increased drag-rise factor, and

,-

to increased ,maximum lift-drag ratio at kch numbers above 0.94. -.

3. Increased taper ratio resulted ti.a slightly reduced variation ,. - of lift-curve slope with &ch number, a Somewhat larger Over-& change -. .- in static longitudinal stability with Mach number at 0 lift, and a slightly -I 1. 8maller change in stability with Mach number at 0.4 lift coefficient. -._

4. An increase of taper ratio from 0 to 0.4 resulted in lower .miniumn ' - drag coefficient and higher maximum lift-drag ratio at Mach numbers above 0.95, and generally a.reduction both In. magnitude and in variation of _~ drag-rise factor with Mach number. .-

Ames Aeronautical Laboratory National Advisory Committee for Aeronautic8

Moffett Field, Calif., Jan. 3, 1955 2 ----

RRFEREXCES .- -

1. Summers, James L., Treon, Stuart L., and Graham, Lawrence A,: Effect8 of Taper Ratio on the Longitudm1 Characteristic8 at Mach - .-,- Numbers from 0.6 to 1.4 of a Wing-Body-Tail Combination RaV2ng an Unswept Wing of Aspect Ratio .3. NACA RM'A$+I20,~1954.

._- ;

2. Hall, Charles F.: Lift, Drag, and Pitching &me& of Low-Aspect- Ratio Wings at SUbSOniC and Supersonic Speeds. NACA RM A53A30, 1953. --G

3. Baldwin, Barrett S., Turner, John B., and Knechtel, Earl D.: , Wall Interference in Wind Tunnels with Slotted,and Porous Boundaries .-

at Subsonic Speeds. NACA TN 3176, 1954. I - .

NACA RM A55pLO3 1) 9

. 4. Nielsen, Jack N., Kaa-Uxxi, George E., and Anastasia, Robert F.: A Method for Calculating the i3f-t and Center of Pressure of Wing-

a Body-Tail ComblnatiOns at Subsonic, Transonic, and Supersonic Speeds. NACA RM A53GO8, 1953.

5a DeYoung, John, and Harper, Charles W.: Theoretical Symnetric Span Loading at Subsonic Speeds for Wings Having Arbitrary Plan Form. NACARep.921,1948.

6. Jones, Robert T.: Properties of Low-Aspect-Ratio Pointed wings at Speeds Below and Above the Speed of Sound. NACA Rep. 835, 1946.

7- Lapin, Ellis: Charts for the Coqutation of Lift and Drag of Finite WFngs at Supersonic Speeds. Douglas Aircraft Co. Rep. SM-13480, Oct. 1949.

.

10 NACA RM A55A03 :

A=lS.l’

I = 19.83 ro= 0.79

(Mmsnslons in Inches except OS mted)

. .

t

A=45D’ A= 53.1*

Leoding-edge meep ‘: Tcper ratlo Aspect ratlo I AIrtoil s&Ion (etreomws 4 Am0 Span Mean oer4yncmic chord lip chord Root chord Iat centcrli~

19.1’ I 4 5.0. 53.1’ 0.39 / 0.40 0.40 3.09 , 3.0 3.0 j

3% Llconvu NACA 0003 NACA 0003 3m7sq in. 3a8m In 38Jmwq IO.95 10.80 IO30 3.70 3.82 382 139 2.06 em 5.1 I 5.14 5.14

(a) shweep-uqle vaicdkm.

Fv 2.-Dimemions of the wing-body cm-s. I

,

I

‘ ’ / I

Taper ralla 0 0.2 0.4 Leadlng-edge sweep 53.r 53.1’ 53.1’ Az+ct ralio .3.0 3.0 3.0 Alrfdl st3cn0n (sfreamvlsa) NACAOOO~ NACA 0003 NACA ooO3 Area 3886 sq In. 36.88aq In. 3888 sq In. Span I cl00 IOBO IOBO Mean aemdynamic chwd 4.60 4.13 3.62 llp chord 0 I .20 2.06 ROOt Chwd (IIt Cenk9rflne) 720 6.00 5.14

(Dlmenslons in inches sxcepf 08 noted)

(b) Taper-mtlo variatbns.

FM 2.- f2mMed.

14 HACA

.

--

Figure 3.9 Typical model installation Ln the Ames 2- by 2-faa wind tunnel..

1.0

.8

.6

.4

CL .2

0

72

-.4

% -4 0 4 8 12 16 (aforW.60) a

ID

B

6

.4

c, .2

0

-.2

-.4

--%I -4 0 4. 8 12 16 (aforWl.02) a

(aI h=E$ x=0,4.

F~4.-varbtbnufliftcoeFficientwifh~ofottadc otcmbmtMochrwnber.

ID

.8

.6

.4

CL .2

0 I

-.2

-.4

+8 -4 0 4 8 12 16 (Q fcfh4=.60)

(b) h450’, X-W,

-%I -4 0 4 8 12 I6 (aforMdD2) cl

(c) h=5v, x=0.4.

Fgm 4.- Cimlinued

-% 4 0 4 8 12 16 (aforM=I.021 a

(d) h=53J’, X=02.

Figue 4.-W

. I

8 ,

1.0

.8

.6

.4

CL .2

0

-. 2

-.4

--ES -4 0 4 8 12 16 [aforM=.60) a

Lo

.a

.6

.4

CL .2

0

-2

-.4

I I I I I I I I I I I I I I I I I I I -% 4 0 4 8 I2 16 (aforM=l.02)

I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I

Q

(e) A= 53, A= 0.

F@m 4.-W.

LO

B

.6

.4

GL .2

0

:2

-.4

I

-B .I

Ga

B

B

.4

G .2

0

-.2

-.4

76 .I

to

B

.6

.4

c, .2

0

-.2

-*4

-6 12 08 .04 0 7044708 -.I2 -J6 ;20 -.24 (c, forMa.60)

to

B

.6

.4

c, .2

0

-.2

-.4

-.6 J

(b) 1i.45, h=a4.

Figre srhtirued.

(cl n=53r, x=04.

FiSrcorrtiued

I I

c , .

ID

B

.6

.4

c, .2

0

-.2

-.4

-6 J

to

#a

.6

.4

G .2

0

:2

-,4

-.6 J

ko n=ar, x= 02.

kare5-cantirred

ID

i3

.6

.4

c, .2

0

-.2

-.4

-.6 J2 II8 .04 0 -708 -J2 -J6 -20 -.24

mcA RM A55AO3 25

-20 co

&FOR M=.60)

(0) A= i9.1”, x= 0.4.

-60

-40

I ! !P! P Id Ix- P n 2-l n Y R

I I I I I

-cn -.“” 0 .04 .08 .I2 .16 .20 (co FOR M--.60)

%

(b) h=531°, X=0.4.

Fqure 6.- Wiatbn of drug coeffent with Iii coeffcient at constmt Mach rxmber.

.

Loo

80

.6p

c, .40

20

id7 i i JYI I I I

(c) A=53.1°, X=0. I

F&u-e 6.-C-.

.j.,i,i / ,

I:,.iII ’ ij, ’ ! I

NACA RM A55403

.

CLd 0

for

CL’.4

0

dCm 3q-O

for

CL%4

.2

0

.2

2 L -. 4 -‘I I t t I t / CL=.2

/ I

-. 2

0 ----------_---.__ I II I , I I I

‘26 3 .8 23 Lo Ll L2 I.3 IA M

(0) A =s.r, x=0.4.

Figure 7.-- Voriotlon of IiftaJrve slope and pitching-moment-curve slope with h&h number,,

27

28

Q-0 for

CL= A

.2

0

.- NACA RM A55A03

.08

.04

0’ ” “‘I ” “‘I “1 .6 7 .8 .9 I.0 11 I.2 I.3 1.4

M

dGrn z5i

-.4

.a .9 ID I.1 12 13 1.4

M

(b) A = 45.0, A = 0.4.

Fkyre 7. - Continued.

: z-i :

T :

-- I .--

,-,_ --_l

NACA Fad a55403 29

2

0

*-0 for

CL’.4

.2

-08

t 1 II OS

II II I I I I I I I I 7 .8 .9 LO I.1 I.2 L3 L4

M

M

(cl A = 53.1 ) A = 0.4.

Figure 7.- Corttiied.

30 NACA FiM A55h03

cLol= O for

CL’.4

2

0

dCm.0 dCL for

CL- .4

.2

0

I , ,

I ,+I I !

.08

” .6 .7 .8 .9 LO I.1 1.2 L3 u

M

.--. -.-

.- -

M

(d) h=53J, X=0.2.

Figure 7.-Continued.

NAOA FM A55A03

CL:‘0 for

CL1.4

.2

0

dcm K-O

for

CL -4

.2

0

t / / i i i 1-i i s cr.4 i i T7-l-H

% .08

0 .6 .7 .8 .S 1.0 1.1 1.2 1.3 1.4

M

-.2

0

dcm dCL -.4

.6 I .a .s 1.0 I.1 1.2 1.3 I.4 M

(e) n=53.1, x=0.

Fiqure 7.-Concluded.

t

32 NAOA RMA55A03

for

CL’.4

CL, I I I I I I I I I I

I I I I I I I I I I I .I2 II I I I I I I I I I I I I I

2 I f I I I I I I I I I I III 1

0 III III I I I I I I I I I I .6 .7 .8 .s LO I.1 1.2 I.3 IA

M A -13.1” ------_ 45.00 -- 53.P

0

4%) = for

CL’.4

0

;..

A

.2

0 0

.21 ’ I ’ 1 .6 .7 .8 .S I.0 I.1 I.2 1.3 I.4

M

..v -

.

&we 8.- Effect of sweep on the vortotion wiih Mach number of lift-curve slope and

pitching-moment-curve slope; X = Q4.

EACA FM A55403

CD min *O’

.6 .7 .8 .9 1.0 I.1 I.2 1.3 1.4

M

‘% dc: .2

I I I I I I I I I I I I

0 .6 .7 .8 -9 LO I.1 12 1.3 14

M

I6

0 L D max

8

4

0

‘+ I

.6 .7 .8 .9 1.0 I.1 1.2 1.3 I.4

M n x

ISA0 0.4 -- 53P 0.4 --- 12.6* 0.5 Reference I --33.7O 0 ” ”

Ffgure S.-Effect of sweep on the variation with Mach number of mlnimum drag coeffcient, drag - rise factor, and maximum lift -d&q mtlo.

34 NACA RM A55403

0

M

4- --I--- 2 -- A

4$$-0 for

CL’ .4

4 f

- _ _.. - -- 0

M

FIgtlre KL- Effect of tqw ratio on the variation with R4ad1 number of lift-curve slope

and pitching-moment-curve slope; A = 53.1:

lw.3 RM A55403 35

-3 L

t > L/D max

.02

-01

01 I I I I I I I I I I I .6 .7 .8 .9 1.0 I.1 I.2 1.3 1.4

M

.4

.2

n :6 .7 .8 3 I.0 I.1 I.2 I3 14

M x -0 -- 0.4

16

8

-6 .7 .8 .9 1.0 I.1 1.2 1.3 1.4

M

Fiiure I I.- Effect of taper mtio on the variation with Mach number of minimum drag

coefficient, drag-rise factor, and maximum lift-drag ratio ; A= 53.10.

36 __ NAOA RM A55A03

0 1 I i I I IlI III I I I I I I 0

.6 .? .a .9 I.0 I.1 1.2 I.3 IA

A(F)=0

for c,= .4

.9 I.0 1.1 1.2 w 1.4 M

h A

19.P 0.4 ----- 45.0* 0.4 - -- 12.6” 0.5 Reference I --33.7o 0 . .

Rgure l2.- Comparison of the effect of sweep oe rorled by two geometric methods on the vorlotlon with Moth number of lift-curve slope and pltchlng-moment-curve elope.

.-

_-_.-. ..

NACA-Langley - S-U. 46 - 360

. -- + f;

. -. c. ;

!.