Investigation of Diving Moments of a Pursuit Airplane in the Ames 16-Foot High Speed Wind Tunnel

MR Oct. 1942.

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

it'll RI' I illE RE PORT

ORIGINALLY ISSUED October 1942. as Memorandum Report INVESTIGATION OF DIVING MOMENTS OF A PURSUIT AIRPLANE IN THE AMES 16-FOOT ifiGH-SPEED WIND TUNNEL By Albert L. Erickson Ames Aeronautical Laboratory Moffett Field, California

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WASHINGTON NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of advance research results to an authorized group requiring them for the war effort. They were pre- viously held under a security status but are now unclassified. Some of these reports were not tech- nically edited. All have been reproduced without change In order to expedite general distribution.

A-65

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS MEMORANDUM REPORT for the Materiel Command, Army Air Forces INVESTIGATION OF DIVING MOMENTS OF A PURSUIT AIRPLANE IN THE AMES 16-FOOT HIGH--SPEED WIND TUNNEL By . Albert L. Erickson SUMMARY A pursuit -type airplane encountered severe diving moments in high--speed dives which make recovery difficult. For the purpose of investigating these diving moments and finding means for their reduction, a 1/6.-scale model of the airplane was tested in the 16—foot high—speed wind, tunnel' at Ames Aeronautical Laboratory. The test results indicate that up to a Mach number of at least 0.7, the limit of the 'tests, the dive-recover r difficulties öan be alleviated and the longitudinal maneuverability improved by the substitution of a long symmetrical fuselage for the standard fuselage.

INTRODUCTION A pursuft airplane developed powerful diving moments in high- speed dives, and these moments have made recovery from high—speed dives very difficult. The difficulties have been discussed in reference 1, and they have been investigated in the full—scale wind.

tunnel and in the i.-fot hih—seed. wind tunnel at Langley Memorial Aeronautical Laboratory (references 2, 'and Ii-).

3, Army At the request of , the Air Forces, a model of the airplane was tested in the '16--foot high—sreed wind tunnel at AAL. The purpose of these tests was to extend the range of the'. previous high--speed tests with a.viewtowad developing 'aéans foi' eliminating the diving difficulties and improving the maneuverability of the airplane at high speeds. A number of fuselage shapes, several changes in the s-pan load distribution, bulges and spoilers 'on the wing and fuselage, and a- modification of the wing center—section profile were tested.

APPPRATUS AND METHOD The l/6—scale.tip—riounted model used in the 8—foot high—speed wind tuneI at LMAL (reference 2) was modified by the addition of wing tips and fitted with trunnion—support fittings in the booms for shows details of the model.

mounting the model on struts,. Reference 5- Stings were attached 18 inches back of the strut trunnions, asshorn in figure 1, for controlling the angle of attack. A ritot survey head was used to explore the air flow in the region of the tail. This head measured the total and static pressures and the pitch and yaw angles of the air stream. In those of the present tests wherein the ed.. aileions and partially extended Fowler' flaps were

effects of droo p

s were simulated.by split flaps having

stAdied j the ailerons and fla p

chords approximately 30 percent of the wing chord at each spanwise . staMard. airplane wing, station. In addition to the tests with the t for the

the model was tested using a wing with revised twiat. Exce p

twist, this revised wing, was identical to the standard wing. The twist was changed only from the boom centêi lines outboard so that the angle of attack relative to the standard wing was increased from

0 at the station where the rounding

0° at the boom center lines to 3 of the wing tips started..

RESULTS The data in this report have been corrected for tunnel—wall effects, and approximately for tare drags and tare moments. The

moment center was 3.23 inches vertically above the trunnion point

with the airpino in the zero angle-of—attabk attitude. TJpf low or downflow with the strut supports in place has not been evaluated..

A slight upflow or downf low would affect only the, absolute values of drag and, for comparative purposes, would have' no effect.

The results crc discussed in the following ordier: 1. Standardconfiguration 2. Effect of fuselage shape Effect of bulges, fillets, and spoilers 3.

Effect of changes to wing center section 4.

Effect of ailerons and flaps 5.

6. Effect of the change in wing twist

Elevator effectiveness 7.

Improvements resulting from use of the long 8.

s:Tmnietrical fuselage Buffeting 9.

Standard Configuration Tests of the standard configuration, complete and in parts, dive-recovery difficulties.

revealed the nature and, cause of the With the complete standard model, the longit.ud.Lnal stability 0.65 increased enormously as the Mach number was increased. above This increase in Stability is illustrated by the decrease in the slores of the curves showing the variation of moment coefficient 2(o) con- with iiftcoefficient in figures 2(a) and 2(b). Figure sists of cross plots - from these curves and shows the variation with Mach number of the thoment coefficient at constant values of the lift coefficient. The lift coefficient at which the moment curves for various Mach numbers intersect is of special significance and.

will henceforth be referred to as the constant-moment lift coeffi- cient. For all lift coefficients greater than the constant-moment value (approximately -0.15 for the standard configuration), the pitching moment decreases; that is, it becomes a diving moment as the speed is increased. For smaller values of lift coefficient, the moment becomes a climbing moment as the speed increases. Thus,.

it is seen that at high speeds the airplane becomes extremely stable. This stdbility is so great that deflection of the elevator, PIS wilibe shown later, produces little change in lift, hence, the difficulty in recovering from dives.

Removing all the accessories (Prestone. oil, and spark-plug cooling scoops, carburetor scoops ., and turbosupercharger installa- tions) from the staMard configuration made no important change in the moment characteristics. Removing the fuselage, however, increased the critical speeft at which the stability started its rapid increase, and also changed. the valie of the constant-moment lift coefficient from a small nerative value (about -0.15) to a positive value of about 0.2 (fig. (c)). Iith the balance occurring at a positive value of the constant-moment lift coefficient, the airplane would. tend. to automatically recover from high-speed dives because a pull-out moment would become effective as the speed.

increased.. Modification of the aIrplane so that the constant moment occurs at a suitable positive value of the lift coefficient, as with the fuselage removed, should provide a moans of alleviating the

II.

dive--recovery difficulties.

With the horizontal tail surfaces removed from the model, relatively smell changes in pitching--moment coefficient occurred as the speed was increased above the critical (fig. I (o)), and the changes that did occur were in the opposite direction to those with the tail in rlace. This result indicated that the moments produced by the tail were undergoing large changes as the speed. increased thereby causing the difficulties. Figure shows downwash angles .5 that were measured at the tail position while the tail was absent.

As the Mach riumher increased. above the critical, the downwash angle $0,

decreased as much as2° or and the variation of downwash with

angle of attack becane only a small fraction of its low-speed value. The decrease in downwaeh was a direct result of the loss in lift as the Mach nLzrnber was increased above the critical rl'ue (fir. I()). The muitud.e of the reduction in downwash corre- sponded approximately to the change in tail angle of attack that would be required to produce the changes in pitching moment shown in figure 3(e).

It was concluded that the dive-recovery difficulties of the airplane are due to the centor section of the wing losing lift as the speed increases abo,,ve the critical. The reduction in lift is accompanied by a reduction in. downwash at the tail and a reduction in the rate of change of downwash angle with airplane angle of attack. The latter change produces a great increase in longitudinal stability at sreeds above the critical speed of the center section.

With the standard fuselage in place, the constant..-moment lift coeffi- cient centers about a negative value of the lift coefficient, and the stability becomes so great that the elevator can produce only small changes in airplane lift coefficient; consequently, recovery from highspeed dives is difficult. With the fuselage removed, a positive value of the constrit-moment lift coefficient occurred, so in this configuration the airplane would tend to automatically recover from high-speed dives.

Effect of Fuselage Shape As it was shown that the standard fuselage caused the moments to break at a lower Mach number and the constant--momont lift coef---- ficient to he negative, several fuselage modifications were tested.

These were the standard fuselage with modified canopy, an underslung fuselae, a long symmetrical fuselage, a long symmetrical fuselage with flat-front cab, and a. longsymmatrical fuselage with the cab from the standard fuselage. The results are compared in figure 6.

These curves are plotted, for three representative lift coefficients: 0.1, 0.2, and 0. 1i. The results show that the fuselage designated "the long symmetrical fuselage" (fig. carried, the moment curve to

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the highest Mach number before breaking, and caused the constant- moment-lift coefficient, to center about a small positive value of the lift coefficient (approximately 0.07). This fuselage also had

a lower drag at Mach numbers above 0.68. The underslung fuselage

gave similar moment characteristics, but it was not considered as 8).

practical a shape (fig.

Adding a flat front to the cab of the long symmetrical fuselage (fig. 9) in order toermit the use of flat bullet—proof glass windshield.made the moment characteristics slightly worse (fig. 6).

Two ca b • changes were tried in an effort to find an arrangement that gave satisfatory moment characteristics, but that would not make it necessary to move the pilot ca nd the controls from their positions in ' The fat—front. cab was moved aft 2 inches the standard airplane.

(corresponding to 12 in. full—scale), and the cab from the standard fuselage was tried on the long symmetrical fuselage. Both of these cab arrangements gave poor moment characteristics (fig.

6) as

compared with the forward cab. The inferior characteristics with these two cabs were probably due to the peak velocities induced by the cabs being near to, and adding to, the peak velocities induced by the wing.

Figures 10 and il givethe complete basic results for the long symmetrical fuselage. Figure 10 gives results for the regular Prestone scoops; figure 11, for modified Prestone scoops. Little difference ui the results is noted,. although there is a slight reduction in dras indicated with the modified scoops.

Effect of Bulges, Fillets, and Spoilers Abrupt bulges were placed on the under side of the fuselage, as shown in figures 12, 13, and 1, to find their effect in causing a shock on the under surface of the fuselage. The first bulge was placed on the under side of the standard fuselage with a revised canopy. This bulge caused the moment curves to rise slightly from Mach numbers of 0. upward, but thoro was no noticeable change in the general effect (fig. 15). This indicates that the upper wing surface had the most powerful influence on thepitching moment.

Other bulges tried on the long symmetrical fuselage wth the cab off had little effect on the Ditching moments. Figure 16 gi7es compar.

-ativeresultswithandwithoutthebulges.

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Several types of fillets were tried on the symmetrical fuselage gave any special benefit (figs. 17, 18, and. 19); but none of these 20) over the constant—radius fillet used on all other tests.

(fig.

The fillet used for most of the tests had a constant radius of one—half inch (3 In. on the airplane) and would be the easiest to build.

Some spoilers and bulges were tried on the under side of the wing. The first sioiler tried was sot 90° to the wing surface, 33--1/3 percent of the wing Drotruded 1/4 inch from the surface at chord from the leading edge. and was extended between the booms.

A second test was made with this same s'oiler extended to the wing ti-ps. For a third teot, the s'oi.1ers were removed and a' smooth bulge one--fourth inch high was located between the booms at the same chord oosition. Those tests were 'oredicated on pressure— distribution data which showed that, at constant angle of attack, as the speed was increased the negative lift on,the lower wing surface increased more rapidly than theositive lift on the upper surface at high sDcedz. The tests were made to determine whether the negative lift increases could be reduced or eliminated. The effects of these various changes are shown on figures 21(a), 21(b), and 21(c). The flow over the lower surface was spoiled to such an extent that from a very low speed, a steady rise in the moment curve took place until the upper surface reached its critical speed, and than at lift coefficients of 0.2 and above, the moment broke in a nagative direction.

Effect of Changes to Wing Center Section As the presence of critical pressures had. been shown to cause the trouble, it appeared that a wing with lower pressure peaks would delay the compressibility break. Accordingly, a glove was built around the original wing between the booms. This glove was set at a lower angle of attack and had a larger chord (fig. 22) than the original wing. This glove had much lower pressure peaks and., with the symmetrical fuselage, raised. the critical speed. and the balancing lift coefficient to a value slightly higher than that for the seine configuration but without the glove (figs. 23(a), 23(b), and 23(c)).

With the standard, fuselage and the glove, on the other hand, the curves broke in the same way and at the same Mach number as for the standard fuselage without the glove.

Effect of Ailerons and Flaps As the loss in lift on the center section caused the large adverse tail moments for a given total lift a change in s pan load distribution that would shift a greater nart of the lift outboard of the booms would relieve the center section of some lift and delay the pressure rise on this section. Accordingly, tests were made with down and extendJngfrom simulated flaps and ailerons deflected 15° the wing tips into the booms. Without the fuselage (fig. 24), an improvement in the characteristics is indicated. The moments increased, at speeds above the critical for lift coefficients of 0.3 or less (fig. 24(e)); whereas without the simulated flaps and ailerons (fig. 3(e)), the moments inceased for lift coefficients only up to 0.1. In addition, the f1as increased the-Mach numbers at which the sudden change in moment occurred. -Several additional simulated flaps

(figs. 23(a), 23(b), and 23(c)Y w i th

runs were.made and ailerons 15 0 down. It can be seen from these comparative curves that the flaps Improved 'the moment characteristics in all oases except when the standard fuselage was used. At a lift coefficient of a positive direction, 0.1, although all other configurations broke in .

the configuration with the standard fuselage broke negatively end, at .the others.

a much earlier Mach number than Effect of the Change in Wing Twist It appeared that an effect similar to that obtained with the split flaps could be obtained in a practically applicable manner by modifying the wing twist. Accordingly, a wing having the twist 3° was modified by increasing the angle of attack at the tips by tested. This change in wing twist improved the characteristics very little, whether used with the long symmetrical fuselage or with the Comolete results of the tests of the standard fuselage (fig.

'5) .

revised wing with the long symmetrical fuselage are iven in The Ineffectiveness of the change in wing twist as figure 26.

compared to the spill flaus an 2arently was largely due to the fact that the twist increased the lift coefficient at each ancie of attack only a small amount compared to the increase in lift ' coeffi- cient produced by the flas.

Elevator Effectiveness Figure 27 indicates, for one configuration, the lift óoeffi- dent at which the airplane would balance at various Mach numbers with several elevator angles. This figure shows that at high speeds

[.1 a given elevator deflection produced relatively small changes in the lift coefficient at which the airplane would balance. Analysis indicates that the loss of elevator effectiveness at high speeds is largely a result of the great increase in stability of the airplane and not to any important extent duo to reduction in the change in tail lift brought about by a given elevator deflection. The results for other configurations were similar, the only important difference being in the value of the constant—moment lift coefficient.

ng From Use of i Im p rovements Result the Long Symmietrical Fuselage The model with the synmtrical fuselage shaved better diving characteristics than th standard. configuration. By taking the roints on the curves of figure 25 where the moment broke, figure 28 lift coefficient attainable axmum was plotted, which shows the without the moment curves breaking ina diving direction. The Mach numbers at which the moment coefficient curves broke agree closely broke c1rVeS With the Mach numbers at which the lift coefficient \ shOw that. at for corresponding conditions. These results (fig.

the limit of -t- 0. 75, zero lift and at Mach numbers up to at least tests, with thesymmetrical fuselage the airplane will not have difficulty in recovering from dives, because the moment is a climbing critical moment when it does. break. If the airplane exceeds- • speed., it will tend to come out of the dive, not stay in it. As a matter of interest, there are also nlotted on this figure curves of the lift coefficient required to maintain level flight at various altitudes.

The long symmetrical fuselage imuroved the longitudinal maneuverability in the critical speei region. The improyement is with the 0.65, •shown in figure 28. For exampleat a Mach number of standard fuselage, the maximum lift coefficient available without encountering the severe diving moments is 0.2. Relaci .r.g the standard fhselare with the long symmetrical fuselage increases the . 5 . At an lift coefficient available for the sce condition to 0 coefficient of 0.2 uroduces oni feet, a lift .

altitude of 25, 000 0 . 65 . Therefore, enough lift for level flight at a Mach number of accelerations that would. require higher lift would put the airplane into the critical diving-noment region. By hanging to the long could be executed. under 2 5 g symmetrical fuselage, acceleration of the conditions of the example without entering the critical region.

Buffeting . o'pservations of the Neithertheresults of forde tests nor model behavior during tests gave any indication as to whether or not tail buffeting occurred. Figures 29 and 30, which give ;rc,SUltS of measurements of the wake at the tosition of the tdll, show that with the standard tail position, thetail will come within the wake of the wing and fuselage at highMach numbers. These results mdi.- cite that raising the horizontaltail surfaces 32 inches above the standard poitiOn should keep them. out of the wake, 0except for Mach numbers above .0.75 at angles of attack above-2 , and hould.

thereby largely eliminate ouffeting Reference 2 makes a similar conclusor Tests were m±de with the tail altered as shown in figure 3.1. The model diensiois indicated cQrrespond to raising the tail inches and moving it back 24 inches on theai.rplane.

Figure 32 shows the aerodynamic cnareteristics in resultg from this change. The only effect as compared to the. standard tail position was an increase in stability.

CONCLUSIONS 1. The difficulty encountered by this pursuit airplane in recovering from high-speeddives is caused by a compressibility shock on the wing center section. This shock causes a loss in lift.

and a reduction in the downwash, which results in a large change in the tail moments; 2. with the standard fuselage, none of the modifications tested eliminated the difficulties.

A long symmetrical fuselage increased, the Mach number at 3.

which the adverse diving moments occurred by at least 005. At Mach numbers ur to at least 075, the limit of the tests, the long symmet- rical fuselage caused the airplane to balance at a sufficiently positive lift coefficient so that recovery from dives could be effected.

Ii-. The longitudinal maneuverability of the airplane at high speeds. can be improved by the use of the long symrnetrical'fuselage.

For example, at 25,000 feet and at a Mach number of 0.65, the airplane can obtain accelerations, as compared with only one g for the 2.5g standard configuration.

Ames Aeronautical Laboratory, National Advisory Committee for Aeronautics, Moffett Field, Calif.

REFERENCES 1. Johnson, C. L.:. A Study of the Divine Characteristics .of the

Lockheed -38

Airplane. Lockheed Aircraft Corporation.

Rep No 26116, 8, 1942

I'Ly 2.

Draley, Eugene C-.: Test's of 1/6_Scale P-38 Model

in the 8—Foot

Tunnel. .NACA RMR, Ma,.

1942.

: High—eea Preston, G. Merritt, and Guryansky, Eugene R.: Drag Analysis of 3.

the Lockheed YP.- 38 Airplane. NACA NR, March

1942.

II . .

Guryansky, Eugene E., and Prestor, G. Merritt: Full—Scale Wind Tunnel Investigation of Buffeting and Diving Tendencies of the YP-38 1941.

Airplane. NACA MR, Mâ±h

Model Design Group: P_32

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0 0 I.2 3.7 A 1.1 2.1 i.3 C Iii 1.5 1.

D 0.1 0.8

Rg

O.b O. K NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS DtMffW5JCW5 ARE i p1c,-I F-gor'l 5UAC5 WING AND FUSEL-A66 Figure 18. - Symnietrical fuselage. Leading-edge fillet with leading edge extended straight forward.

3_. 3j• .• -4.--

3"-4- 4 3

- - - -

- - rII4I At 4

• FILLET

rTA STA.

-jzoNTAL.

vEETICAL.

.0-46 A - 1' _ C I.G' - NATIONAL ADVISORY DIMENSIONS AE - INCHES FROM COMMITTEE FOR AERONAUTICS WING F - USELAGE SURFACES Figure 19. - Symmetrical fuselage with expanding fillet to the rear.

M c Figure 20. - Comparison of effects of various fillets on symmetrical fuselage.

or c'fc "A 960 5Po/Ee5 .33. %

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Q

CE QVE £AJ7%5 CA) iaM.'E

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w/AIc a.4U - .VA1 87-'

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.t) LOg &d'iACE ^ /A/A/,-R W/Ai& c'A., LOW I/JAIE 5tJgd4CE p4A/EI_ £VAI 34AI.ERONS PLAP5 00/'P10 I5 4PJOAPO $ ,,,.js oropP 3Z

CM

64.- iEVSO 5 C W/"Cl CAJ7E I 32.

- 64 3LA TNOA

C

L

-.2

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS = 0.1. - - (a) CL Effect of wing spoilers and wing bumps on pItching moments.

Figure 21. -

.G

A'7

c 33!

OA, LoL-q.--/,e .-u.eP-40 ' - l,AI 5c'AA, 87

.3

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.2

CMOL' 2A.14Oih'

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f - C

• / /5 A/L/'RON 4 iI5 D.c?OPf-'E 84 CONF/6. -

CAI

pf717 /5 - 35 A/LEOAA5 pQ IRI 3Z $L,-r5- -- ----___. -

0. Z

C L -

-2 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS (b) CL = 0.2.

Figure 21. - Continued. Effect of wing spoilers and wing bumps on pitching moments.

/T

X 90 5/0,LER 33 * % CHORD OAf La WE 5iRfl.4cR 87 * OvEF £tJ7/A t-tIAJG 5p.4,4J so,€s q 33 % CHORD OAf LOiS/El?

/A/A/--R W'kJG P,4'JEL - 5 uc-.-lCe

-3

86 - -- 8cJIv 3 C//O p ON LOWE/C 7' l.A'Mi(.

S UR.t/1 CY IA//JAR fT4,JEL Ru"J87 ./C'zJ.v 86

CM

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GL =: 0.4

—.2

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS (c) CL Effect Figure 21. — Concluded. of wing spoilers and wing bunipe on pitching maneate.

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NATIONAL ADVISORY - • COUNITTEE FOR AERONAUTICS Figure 22. - Wing glove mounted over standard wing between booms.

WITH 41LEie01V5 P4R3 DOWA) yA4M - LQG PLAGE AON5 /5

/

60- LONG .5YMNI P EL.4(, /1/ROA'. /5, ZLoV .4//I5/f, 41/A1G 27-NO sEL.. q G,. EON5 Ic 34MO,90 I.SE4G,

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at o

:-:TTT-TT:EEEEfr-

r4fiJCAQ' ,U5eL4 Ie1N 84 . - Sy4A4. 4c •4 54 ';Z7 7 ,41VD,4,e0 * ^4 1- 5E'LA69, a-LOVe 3A1l'I 59- LOAI PueL,q GC GLUVE : ,O ?oc- LONG ?Yd1%4 ) W/fIi WITh' TW".i i9/5.4G L ilL VILJ) [4LL E'(JA15 W/7W W/NG.s 5oaAI5, 4 721/L (€0)] NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 0.1.

(a) CL Comparison curves of configurations tried with and without Figure 23. - and flaps.

ailerons

7- PLAP. .DOw,V

kv

(_.

0.

7\4 -; A.w 4 iv& $"i'o' , 414 g eo'$ /5 Covs r.2A..7 IWJ,c/S 4A(.7t5 -.1• LONG $YMM /E1.4GE J Lbi'E 9/bYOV$ /5 S /GVG SWtM 4/LffiOv /5 t7F /LT X- 2.7 - NO N'qkI.2 4i1S:•/S° 34. . 5 .1.

1 00AI.

1O5 \ 91 LcW SY/V7MFV$EL4GE' W1 &LOVE 64-SiQAJA1 p9 - I ON 5YMM F1SL. 'LOl,4 AAO 5VM41. EL.q(.E WING vvir,I RE V/SEQ. rwisy -. log- '/.o NO 4E4'01V.5 OR.PZAP5 7Q/LS0 e°J i4',VG 545 cL--O NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 0.2.

(b) CL = Figure 23. - Continued. Comparison curves of configurations tried with and without ailerons and flaps.

R/N 4 LONG .VM/4 UL45, A/LCeciv /9 w7-M-r P/U F'1.JT X ,5YM. (/5ELAG IS, - 5- 0AJG 5YMl%1 J3R A/ccc7&i i; ff4m,ur,- /OGE r/ILFT X, Z7 - NO .4cE, ,4/iI F 5 34- .5rco . . -:... .: • LP5 :DQAJ CM It/C ) .4/LE,€O,t/5 01e / .-4i CM 0 - - — / R&M 80- S7NQQ E4GE 94/0 - LONG S 1• •. .• .

d 4- 5 V4F GLO -.. . . . .

• - . 35- LciVG .-YMM fUeLA, G.Lc.)V . ••c4• L c : .

si lo- ,W -F1 /05 L0,.JG Y44,1 PUL-(, W/iW. WIT/I Nw'L' PW/-sr O' o)J

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(

Cz

. NAI ION AL ADVISORY - . .

COMMITTEE FOR AERONAUTICS .

(c) CL = O.k.

Figure 23. - Concluded. Comparison curves of configurations tried with and without ailerons and flaDs.

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C

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I

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0.

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x 75O

.2 4) •r4 c C

.1

I

C3

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS (b) Concluded.. Basic data from Macli number 0.675 to 0.750.

Figure 24. - Continued. Wing, booms, tall, ailerons and flaps drooped 150 from wing tip to booms.

c NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS (d) Variation of CD with Mach number at constant angles of attack.

tail, ailerons and flaps drooped.

Figure 24. - Continued.. Wing, booms, 150 from wing tip to booms.

cJ () / 2!

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS (e) Variation of CM with Mach number at constant lift coefficients.

Figure 24. - Concluded. Wing, booms, tail, ailerons and flaps drooped.

150 from wing tip to booze.

II, im I t 0M .3 .5 M II. LONG ay MMx.a1ckAL FoLAGe JTANPERD WJW 4 5TAEs4 DRO FUSELAGE, 6TANDARD WING TWIST WING J.05 LONG. sYMM.:. FSELiL Iii •8Tft$tDA.D LAG RE VIS E D TWiST WING +.I °

CM

Co

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 25. - Effect of the change in wing twist.

Figure

1) to .

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'-4 0•

0 -4-'

•

cn N- 00 o 0 oa C) IP 0 N- • o Ea

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A •/OO El .7Z5 .•735

CL

iS

f

XX

.2 C L ó -. a - p —.4 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS (c) Variation of with Mach number at constant angles of attack.

CL Figure 26. - Continued. Wing with revised twist, booms, long symmetrical fuselage, constant-radius fillet, and tail. Elevator at 00.

.14 i .12 I

T55r4 1058

ReIN SY M8OL oe

/ f

./0 .08

Co

D2

.741

fJ NATIONAL ADVISORY CONMI TTEE FOR AERONAUTICS (&) Variation of C D with Mach nunber at constant angles of attack.

Figure 26. - Continued.. Wing with revise& twist, booms, long symmetrical fuselage, 00.

constant-radius fillet, and tail. Elevator at

S

C=-.3

C/11

-.1

-.2

NATIONAL ADVISORY COUMITTEE FOR AERONAUTICS (e) Variation of with Mach number at constant lift coefTiclénta.

CM Figure 26. - Concluded. Wing with revised twist, booms, long symmetrical fuselage, constant-radius fillet, and tall. Elevator at 0.

PU •11 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS Figure 27. - Wing with revised twist; booms, long symmetrical fuselage, and tall.

Lift coefficients for balance at various elevator settings.

for balance at various elevator settings.

1 C C REQO/REI OR LEYFZ C 0,R7A11 VA8I( W17-1-1/1r fl416Nr A7 • 41-7-171,€E.5 A 5Rl4K IN Cm CIJ/J'E .5112IVN 84.$ED ON fl'/M .5 .4 .3 CL .2 .1 rej NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS Figure 28. - Maximum lift coefficient available with chanie in Mach number for two configurations.

5Z 5mo,ze C#w.o / 47= \ /f ..?c .50 f) ;' = cx= -2 zo 8 0 to .

--

C 4-

I

/ (

/

U-, A15 -2 or CIC

,g .8 1.0

p

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS (a) Mach number, 0.30 and 0.50.

Figure 29. - Eelative position of the tall and wing wake for the wing and booms alone.

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I

/

/

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,=.7$•_ \\ a ° - / \\ \\ \\ -4 .6' 10 S.. \ \\ it -/0 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS tch number, 0.65 to 0.75.

(b) (b) t .

Figure 29. - 'CcncluOed.. Relative position of the tail and wing wake for the wing and bóons alone.

ZL7' 25CC - (N ( ( .

\ /t4 30 14=. 30 / O 2,9 • 1.0 - Q9 I_ c) czl7L?7ZZZZZZlZ2 9 - ( A'- IN q-- = SCALE Q 9 ('- //) rA rr?277171LZrz== / : .

\\

T

/0--- AUONAI. ADVISORY COMkIT1EE FOR AER3NAJTICS (a) Mach number, 0.30 and 0.65.

30. - Figure Relative positions of the tail and wing wake for the wing, booms, and standard fuselage.

7' IV - - - -- - - \ - 7j_ - \ \ \ __ -

\\

/VOTE5 \ \ 0N - \ \ =(-,ç,)=5c,-iLE

\

0.8 1.0 Q/9 aa, 0

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS (b) Mach number, 0.75.

Figure 30. ;- Concluded. Relative positions of the tail and wing wake for the wing, booms, and standard fuselage.

7AIL. PO5ITN RAISED I _ 5TANDARD TAIL Po5moN

N /

--'

\

/

/

I 1..

/ LINE

rusLAG I

\

/

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS LI Figure 31. - Relative positions of the present standard tail and the raised tail.

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-2

04 0 (

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x In

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Q 473

.70o .73

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0 4011

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cV .

ILI V . V..

4 5 rD c'J fn

•75 0.

70o b .72 x .750 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS (b) Concluded. Basic data from Mach number 0.675 to 0.750.

Figure 32. - Continued. Wing, booms, standard fuselage, and raised tail.

.10 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

ME

•z. •$' .1 7

Al

(d) Variation of CD with Mach number at constant angles of attack.

Figure 32. - Continued. Wln€, booms, standard fuselage, and raised tail.

.1 s.1 .3• .7 I, I, / ii .1

C4

NATIONAL ADVISORY Oki C MITTEE FOR AERONAUTICS (e) Variation of CM with Mach number at constant lift coefficients.

Figure 32. - Concluded. Wing, booms, standard fuselage, and raised tail.