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RESEARCH MEMORANDUM

AERODYNAMIC CBAFLACTERISTICS OF A WING WITH QUARTER-CHOFCO I LINF, SWEPT BACK No,ASPECT RATIO 4, TAPER i FLAT10 0.6,AND NACA 65A006 AIRFOIL SECTION

TRANSONIC-BUMP METHOD By Thomas f. .Erg;Jr., and Boyd C. Myers, II Langley Aeronautical Laboratory Langley Air Force Base, Va.

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.NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON September 6, 1949 UNCUSSiFfg-? P a

RATIO 0.6,. e NAGA 6s006 -OIL SF,CTkON

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AB part of a transonic resemch.prog??am, a series of win@Oay combinations qe being investigated in the Langley high-peed 7- by 1o"foot' tunnel over a Mach nmiber range from about 0.60 to 1.18, by use of the tranaonic4Ump test technique. i This paper presents the results-of the investigation of a win-ane and a wing-frmelage configuration amploykg a wing with quartemhard line swept back 600, aspect ratio 4, taper ratio 0.6, and an WA65~006 airfpii section. The results are presentedas lift, drag, pitchine moment, and bendineanent coefficients for'both confipatim. The effect of a wing fence on the wing-fuselage charsbteristics was also investigated. In addition,e.ffective downwaah angles and point dynamic 'pressures for a range of tail heights at a probable tail .length &e presented. for the two configuratiomFnvestlgated. Only a brief ' Wsie is given in order to facilitate the publiehhg of the data.

IIJTRODUCTION

A series of wiwfuselage canbinations ISbeing inveetigated in the Langley high-peed 7- by 10-foot tunnel to study the effects of,- gemtry on longitudinal stability characteristic8 at transonic speeda. In the trm.8dc-bmq technique used, a Mach nmber range of 0.60 to 1.18 is obtained.Previous datapublished in this series are presented in references 1 to 3. 2 - XACA RM ~9~27 This paper yresents the resulk8 of the investigation of the whg- .. alone and WTng-fuselage configurations .employing a wing with the qumter-chord line sweptback 60°, aspect ratio 4, taper ratio 0.6, arid t an NACA 65~006airfoil section parallel to the free stream.

The wing of the semiepan model had 60° of sweepback referred to the quarter-chord line, aspect ratio 4, taper ratio 0.6, asd an MACA 65~006 airfoil section(reference 4) parallel to the free stream. The wing was made of steel and the fuselage of brms. A hrwview drawing of the model is presented in figure I and ordhates of the fwelage of fineness ratlo 10 are given in table I. Details of a wing-fence arrangement tested are shown in figure 2.

The model was mounted On an electrical'straqn-gage balance enclosed in the bump, asd the lift, drag, pitching moment, and bending moment about the model plane of symmetrywere mea8ured with calibrated potentiometers.

Effective dowriwmh angles were detemed far a range of tail heights by measuring the floatfng angles opthe tails at five different positionswith calibrated slide"wire potentfameters.Details of the floating tails are given in figures 2 and 3, while a view of the model mounted on the bmp, showing three of the floating tails, fa given in figure 4. The tails used in %hi8 investigation me the 8- a,a those used in references 1 to 3.

A total-pressure rake W&B used to determine the dynamic+resame ratios for a range of tail heights in a plane whlch contained the 2-ercentmean+rodynamic"chord point of-the free-floating tala. The total-pessure tubes *e spaced 1/8 inch apart near the chord line exte-ded and 1/4 inch ap& asewhere.

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I 3

. c, I

bendlnmnt coefficient about root chord Use (at plane

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9 effective dsnamfc .pressure over span of mdel, po&s per square foot (p~2/2)

S twice wing mea of semiapan model, 0.125 squaze foot - C mean aerodgnamic chord of wing, 0.181 foot; based on

relationship c%y. (us% the theoreticaltip) S ..

C local wing chord - .I Ct mean aerodynamic chord of tail

b twice span of semispan del

spanwise distance fram wing root -. P air density, slugs percubic foot v free-stream velocity, feet per second

M effective Mach number over span of model . Mach M2 local number I Ma average chordwise local Mach nurmber - R Reynolds number of wing based an c

. a angle .of attack, degrees 1 - .E effective downwash angle, de@;rees ratio of point dynamic pressure, along a line c.mta5ning the ! Q qumter-chord points of the mean aerodynamic chords of the free-floating tails, to the local free+tream.dynamic preseure " 4 - NACA RM LgG27 ht tail heightrelative'to wing chord plane extended, perced semispan; positive for tail positions above chord plane extended I a.c. aerodynamic center

Subscripts:

M at con8th-t Mach nmber

CL = 0 at zero lift

The tests were conducted in the Langley high-peed 7- by IO-foot tmel by use of an adaptation of the XWCA uin@low technique far obtaining transonic speeds. Themethod used involve6 the mounting of a model in the high-velocity flow field generated over the curved surface of a. bmrg locatea on the tunnel floor. (see reference 5.)

Typical contours of 1ocal"ach nunibere in the region of the model location orl the bump, obtained from surveys with nomodel in posttion, are shown in figure 5. There ia a Mach number gradient whioh results in a diffQrence af about 0.04 over the'span of the model at the lowest aid higheet .Mach numbers with a maximum difference of about 0.07 present at a Machnumber of about 1.0. The chordwise Machnumber variation is generally lepe than 0.01. Ro.attampt has been made to evaluate the effects of these spanwise and chordwise variations in Mach number.The long-daafied lines shown neax the wlng.root repreeent a locd Mach number 5 percent below the rmxlmm value and indicate the extent of the bmrg boundary layer. The effective test Machnumber was obtained fram contour chart8 similar to thosg presented in-figure 5 from the relationship

The vaxiation of mean test Reynolds number with Mach number is shown in figure 6, The boundaries in the figure indicate the range in Reynolds number caused by variatiorm in test conditions during the course of the investigation.

Force and moment data, effective downmeh anglee, and the ratio of dynamic preeeure at 25 percent of the mean aerodynamic chords of the free-floating tails to flree-tream dynamic pressure were obtained for UCA RM L9G27 - 5 the model confi-guratione,tested through a Mach number range'of 0.60 to 1.18 and.an angle-of-ttack range of -2O to 8.O.

The end-plate tares on drag were obtained through the test Mach n'lrmber range at zero angle of attack by testing the model canfiguratims without end plahs. .For these tests a gap of about 1/16 inch was maintained between the whg root asd the bump .surface, and a sponge- xipep seal waa faatened to the wlng butt beneath the surface of the hw to prevent leeas (fig. 4(b) ) . he drag end-phte tares were aas-d to be invasiant with angle of attack .and the- tar06 obtained at zero angle of attack were applied to dl drag data. Je" cnrrectione have not been evaluated inaemuch as the boudaq conditions to be satisffed are not rieously deffned. However, inR.nmuch aa the effectiye flow field is lmge cmgared with the apas d chard of the model, these correctiom are believed to be Bmall. The possibility of change In aerodynamic characterTstics of the wing due to twlst -resulting fram bending mder aerodpmmic 10- con- sidered. Fram etatic loading tests and reference 6, it was estimsted that. at the highest Ma& number attahed the effect of twist would be .to cause a farvd Eier-c-ntermovement .of about 2 percent. RO c&rectiane have been applied to the data presented.

From meas-nts of 'tail floating angles xlthout a delFzlEtdled, it wss determined that EL tail spacing of 2 inches relative to the wing . chord plane would produce negligible interference effects of reflected shock waves rn the tail floating angle's. Downwaah angles for fhe alone configuration were therefore obtained simultaneously f& the middle, highest, and lowest tail positions in one series of teste and for the

t'wo Intermediate positiorm in succeeding rum. (See ' fig. 3. ) For the wi-elage test8 the effective downwash &ngles'at the chord plane extended were determhed by.munting a free-float* tail on the center lfne of .the 'fuselage. The downwash angles presented are increments from the tal floating angle6 without a model in position. It should be noted that.the floating angles measured are actmUy a measure of the angle of zero pitching m&nt about the tail-pivot axis rather than of the angle of zero lift. It'- been esthated that, fathe tail. arrapgement used, a 2O spaswise domwash gradient over the tail win I result in an error of less than 0,.2O in. the resultant floating angle. . Total pressures obtained *can the tail amey rake have been corrected for bow-wave loss. The static-gressure values. used in cmputing dynamiypre6.sure ratios were 0btalned.b~use of a static probe . without a model In position. 6 NACA RM ~9~27

A table of ‘the figures presenting the results follows: Figure Wimlonefwce data ...... 7 Win&melage fmce data ...... 8 Effect of--~fngfence (wing-fuselage) ...... 9 Effective downwash angles (wing alme) ...... ; ...... 10 Effective dovnwash angles (wing-fuselage)...... ll Downmh gradients ...... 12 DynamZvrassure surveys ... ..-...... 13 summary of wrodynamic characteristics ...... 14

Unlees gated, the &iscussion iB based on the amunary curves presented In figure 14. The slopes have been averaged at CL = 0 over ! a lift-coefficient range of 9.1.

Lfft and Drag Chaxacteristfcs The win@cme .lift”ve slope (fig; 14) was a conetant value of 0 -042 *am a Mach number of 0 -60 to. 0.96. Unpublished low-peed data fram the Langley two-dimeneion+ low-turbulence tunnel far a gemtri- cally similar model also gives a value of‘ 0.042 (Reynolds number, 1.5 X lo6) . Abwe a Machnumber of 0.96, the wiq+Lone lift-curve slope decreaaed, to 0.038 at M. =I 1.04 and remained con&ant to a Mach number of 1.18. The addltim of the fUElela-ge lncreased the lift-curve slope approxlmetelg 10 percent throughout the test Mach number rmge. At a Mach’ nmker of 0.60 (see fig. 14), the Wlng”0ne minhmm ’ drag coefficient was about 0.006 as’ccrmpared with an average di.ag-coefficient value of 0 .W45 obtained at low speed in the langley two-djmemional lowyt;[email protected] tunnel fCr a male1 of th same geaanetric characteristics (Reynolds number, 1.5 X 106 to 6.0 X 10 z) . (Note that the drhg coefficients presented have been corrected for end-late tares while the drag data of references 1 to 3 have not been so ccrrrected.) At zero-lift, the wing-done drwriseMach number is not readily apparent since the rate of increase-in drag coefficient ig quite low. It should be noted that the drag coefficient attained at the highest Mach number waa cmJy 0.012. For the win-selage configuration the drag rim occurred at about M = 1.02 and the rate of drag rise wa~ considera7ily more pronounced than for the wiwone configuration. No correction for the fwelage baee pressure has been applied to the wing- fuselage drag data. 7

i CharacteristicsPitching4cment Near zero,lift coefficient the -one aerodynamic center waa about 34 percent mean aerodpamic chord zrp to a Mach nmbr of 0.97. Ertrapohtian and interpolation of da-ta frcgn reference 7 indicate a theoretical aerodynamiclcenter location of about 30 percent mean aerodgnamtc chvd, although rmpublished Langley twcHUmensianal lo+ - turbulence-tmeldata on a geometrically sfmflm xLng gave an aerodynamic"center position of; cdy 23 percent meazi aerodynamic chord. The addition of the fuselage to the- isolatedwing moved the aeroQpmmlc . center rearward about 2 percent map aer-c chord q to a Mach numiber of 0.9. Above M = 0.9, the reedaercdynamic-center movement caused by the fuselage increased rapidly and was about 30 per" cent mean a&roQnmlc chord at M = 1.15. It .can be seen in figures-7 and 8.tha-t the pitc-ament mesmcate a trend toward instability at the relatiqely low lift coefficient of abbut 0.s throughout the Mach nmiber range.

Effect of Wing Fence

In an attempt to alleviate the -table -t;rend o# the pitchin mament cmes at a relative- low ust coefficient, a .afence Fig. 2) located on the mean aerodynamic chard waa investigated bn the wing- . I fuselage configuration. Nee zero Uft the fence appeqed to decreaee I the slope of the pitcwament curves slightlyat'the lowest and . highest Mach nmibers with a samewhat mare inboard locatfcm of the lateral center .of pressure indicated from the bendwdnt curvee. (See fig.. 9.) At the highest lfft coefficients obtaihed (% Z 0.3 to 0.4) the fence produced a pranounced stabutzing trd tu th;4 pitevnt characteristic8 at all Mach numbers aud the lift-cur-ve slope appeared to be increased somewhat.

- Downwash and Dynamic-Preseure Surveys in Region of Tail Plane

The downwash gradient (a €/aa)Mfar the wing alone vaxied little with tail height through the Machnumber range. (See fig. 12.) near . zero tail heightthe addition of the fuaelage appreciably increased the dourwash gradient (3 as .the tail height apgroached the chord - plane. At the higher lift coefficients, far both the sone and canfiguratione, (a ~/aa)M would Fppear to be generd~ - lower .for tail heights below .the chord plane'and higher for tdl heights abwe the- chord plane through the Mach number range (figs. 10 and U).

.. 8 7 MCA RM LgGr

The resultk of point dynamicqressure e&eys, made in a plane Perpendicular to the chord plane extended ‘at a = Oo, contalning the 25 percent mean-aerodynamic”chord points of the f’reeifloating tails used in the downwash surveys, are presente’d in figure 13. There is very little chmge .in the wake characteristics aa the Mach number is increased to 1.10. The addition of the fuselage had little effect on the aynamic- pressure ratios through the Mach number range.

Langley Aeronautical Laboratojry National Adviscq Carmnittee for Aeronautics Langley Air Force Base, Va.

1. Weil, Joseph, an& GOodeon, Kenneth W.: Aerodynamic Characterletice of a Wing with Quarterqhord Line Swept Back 45O, Aspect Ratio 4, .Taper Ratio 0.6, and NACA 6%006 Airfoil Section. Transmic-Bwnp Method. lUCA RM LgA21, 1949. 2. Sleaman, William C., Jr., and Becht, Robert E.: AerodynamicCharac- I teristfcs Of a Wing with Quarter-Chord Line BweptBack 35O, Aspect Ratio 4, Taper Ratio 0.6, and mACA 65AO06 Airfoil Section. Transonidump Method. WCA RM LgB25, 1949.

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5. Scheiter, Leslie E., and Ziff, Howard L.: Prelimim Investlgation of Spoiler Lateral Control on a; 42O Sweptback Wing at Transonic Speeds. NACA RM Lp19, 1947. 6. Stevens, Victm I.: TheoreticalBasic Span Loabfng Characteristics of Wings wlth Arbitrary Sweep, Aspect Ratio, and Taper Ratio. mm m 1772, 1948. 9

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x/2 42 0 0 0 -005 .0023r .04143 00075 .00298 .O4l67 .0125 .00428 .Oh30 -0250 .00722 .04O24 0%) .01205 .03843 0750 .01613 -03562 .lo00 - 01971 .03128 .1%0 0 02593 ~2526 .2000 03090 .02000 .2500 .03465 .01852 .3000 .03741 .on25 - 3500 ,03933 ,00439 .Moo .ob63 0 L. E. radius = O.OOO52

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Figure 1.- Gaeral arrmgement of mdel wlth bo Bweptback wing, arrpect ratio 4, 'caper ratio 0.6, and NACA em06 ail.foi1.

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Y- End date used with

Figure 2.- Details of wing fence and free-floating tail munted an a model with 60° aweptback wlng, aspect ratio 4, taper ratio 0.6, and RNA 65~005 Grfoil. F 1 0.60 .

Fib 3.- Details of free-floating Wls uses Fn mays behind madel with 60' mptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65~006Moil.

...... ,. EACA RM -27 .

I NACA RM w27

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(a)view shaming mdel &a tested. : Figure 4 .-Model mounted an bmp with three free-floating tails installed. WFng-dane- canfiguration. NACA RM ~~27 -

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(a) View shomg model as tested.

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(b) Cutaway view showing seal fastened to wfng butt. Figure 4 .- Concluded. !

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I M.0.77

IO 12 I4 16 18 "- Nominol boundary-layer thickness

MULOO M47 1.2

1.0

cl . .8 2 d (I] 2 .6

..4 .

L

12

8

4

0

-4L"""'

' ' -.2 0 ' .2 .4 -.2 0 .2 .4 -*2 0 .2 .4 Lift coefficient, CL Lift coefEcient, CL Lift coefficient, CL

...... -

' -- I 20

I

M 0 .1.15 0 0. 1.10 i3 0 1.08 0 1.05 0 .1.03 0 0

0 0 2 0 0 1 0 0'

-1 -2. 0 .2 .4 Lift coefficient, C, u I

Figure 7 .- Cmcluded. M M

0 1.15 1.15 0 1.10 1.10 0 bo 0 .s 0 1.08 1.08 0 1.05 0 U.-0 1.0 5 1.03 1.03 0 1.00 1.00 0.12 .9a .98 0 .95 .85 0 -08 .a .90 0 .ao .80 0.04 0 .70 .70 0 0 .a 0 0 .Bo 0

-4 . "1-~

-.2 0 .2 .4 -.2 0 ' .2 ;4 -.2 0 . .2 . .4 Lift coefficient, CL Lift coefficient, CL Lift coefficient, CL . 22 NACA RM L97

M

0 0 0 0 0 0 0 0 0 0 0

2 0 0 2 Lift coefficient, CL ...... , ...... * , I. .

o Fence off A- Fence .on

-.2 . 0 .2 .4 -.2 0 .2 .',4 -.2 0 .2 .4 Lift cwftlcient, CL Lift coefficient, CL Lift coefficient, CL

...... hD a, P

.. Y -80 -40 '0 40 80 - 80 - 40 0 '40 80 - 80 - 40 0 40 80 Tail height, $ ,percent semispan

Figure 10.- Effective downwash angles in regon of tail pllrae for a model with &lo eweptbdk wing, apct ratio 4, taper ratio 0.6, and NACA 65~006airfoil. wing alone.

...... I

4 4

2 2

0 0

-2 -2

M aQ)

.. -EC,-40 0 40 80 "80 -40 0 40 80 Tail height, ht ,percent semispan I

Eygure lO.-CanclUd€d. 6 6

4 4

2 2

0 0

-2 -2

6 6

4 4

2 2

0 0

-2 -2

-80 -io 0 40 80 -80 -40 0 40 80 -80 -40 0 40 ‘80 Tail height, 4 ,percent semispan

.. . ..

Y

M 6 E a, c 'c1

w-4

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M 6 aQ)- W 4

2

0

-80 -40 0 40 80 -80 -40 0. 40 80 Wing alone "_ "fuselage

E

A

0

B

4

0

-40 -20 0 20 40 20 -20 0 20 40

Mgure 12.- Variation wtth tail helat of doxnvash gradlent for' a model XltQ 600 sveptback hg, aepect ratio 4, taper ratio 0.6, ana NACA 65AOO6 airfoil.

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...... , , . , ......

L

.I M = 0.80 a= yo -M = 0.-70 a= 7" M = 0.86 - v qwake9 .a Wing alone ""_ Wing -fuselage 0 a= 4 a= 4O qwake 1B2ml V. 9 .8 ,

...... ,

M = 1.06

a= 70

Wing alone ""_ ' Wing-fuselage lYa 4O a= 4O 1.2 1.2 qw ake qwake "" "- - " 4 .8 Q. .U T 0 a= oo Qf= 0 1.2 1.2

-qwake x-/. -qwake - " " 9 9 I=as .8 .8 - 80 - 40 0 44 80 - 80 - 40 0 40 80 Tall height,- % ,percent semispan

Rgure 13.- CmluBed.

...... - ...... " . Wing alone "" Ww-fuselage w [u -"B -"B .04 CD L=o ($9 4 0 M G

a

0

......

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