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ADVISORY

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WAlrrmm Ium)lvrORIGINALLY ISSUED \,

June1943 asAdvanceRestrictedReport.3R?3

WIND-TUNNELiNVESTIGATIONOF EFFECTOF YAW

ON LATERAL-STABILITYCHARACTERISTICS

v’- SYMMETRICALLYTAPEREDWING WITH A CIRCULAR

FUSELAGEHAVINGA HORIZONTALAMD A VERTICALTAIL

By ArthurR. Wallaceand ThomasR. Turner

N’ACiA ~~‘‘WASHINGTON

LilT.AC.AWARTIME REPORTS are reprints of papers originally issued to provide r+d distribution ofacivancs research results to an authorized group requiring them for We war effort. They were pre-viously held under a security status but s-re now Unclasstiied, Scme of these reports were nc)t tecn-nical.iy indited. .LJ1 l~aVkJ been reo]..,~rjuced ~~~i~n~ut ~,fi.~l~e in crde.r co e;~)edi te (;~nel-:li C!i.5tril)l.l*.iCjll.

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JW7ABCM EESTRIOT~D BEPORT,.., -. ..

OR IIATllRAkS!CABIIJITY0HAJiACT31RIS!CIOS ..

.v- SYH141WRI(M,IJLY9!AJ?EEEDWIMG WITH A OIRGIJIU

I 3TJSS!MiGIlEAVI19G A HOEIZONTL..:..,,

I& Arthur R. Wallaoe ad

AND A VEETICAL ZAIL

Yhomas R. Turner

SUMMAEY

Tests were made in the LMAL 7- by 10-foot tunnel todetermine the effect of a horizontal tall on the lateral-stab51ity characteristics of a high-wing. a midwingo anda Iow-Vlng monoplane The model combinations conmlstedof a circular fuselage. an ~ACA 23012 tapered wing. andan NACA 0009 horizontal tail surfaoe. Each wln~fuselagecombination was tested vr~tha partial-span split flapneutral and deflected 60 and with and without a singlevertical tail. Tests were also nade of the fuselage withand without the ta~l surfaces.

“ The effect of the horizontal tail Ss shown in thepresentation of the results In the form of Increments ofthe rate of change In the coefficients of rolling moment,yawing moment, and lateral force with yaw oaused by win=fuselage interference. The coefflolen$s at high anglesof yaw for all model configurations are presented, Thedata are oompared with data from 61milar model combtna-tlons without the horlsontal tail.

Yhe a“ddition of tho horizontal tail wtisfo-antitoreducci tho variation of the wing-fusalage interference andthe chance In tha offoot of wing-fusolago interference cnthe vertical tall with vcmtlcal position of the wing ontho fusela~a. T+G pr}:sonca of the horizontal tail incroasodthe effective aspe~t ratio of the vertical %aSl by 20 ta 6(2peroents dependihg on the nngla of attask. For angles ofyaw larger than about 15° the horizontal tall slightly re-duced the effectiveness of the vertical tail.

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Conoid.erable data are available for the evaluationof the effect of aerodynamic interfez=encebetween wing,fuselage, and vertical tail on lateral-stabiltty oharao-teristics (references 1, 2, and 3). These data ”indioatethat tho vertical-tail effectiveness Is greater with thewing in a low position ontthe fuselage than with the wingIn a high potiition. klr-flow surveys in the region ofthe vertical tail showed that the change”in tall effective-ness with wlag position”resulted from a sidle flow themagnitude and direction of which were functions of wingpositiou (reference 4). Beoause the data of references 1to 4 were obtained for models without q horizontal tall,the question arises as to whether a horizontal tail willmodify these resq.lts. The horizontal tail has been knownto iqc~ease the effectiveness of the vertical tail by act-Ing.as an.end plate. A theoretical analysis.of this end.–platq.effect was made in reference 5..

.,Tho present report continues the .lnveatigation of

lateral-~tabillty characteristics by adding a fourth part,the horizontal tail, to the previous model consisting ofa wtng, fuselage, and vertioal tall. The purpose of thepresent report” is to determine to what extent the hori-zontal tail Influences the effect of winefuselage inter-ference on the vertioal tail and to determine eqerlmen-tally tho end-plate effeot of the horizontal tail on thevertical tail. . ..

MODUL MD APPARATUS

m-.ne tests were made in the LMAL 7- by lo-foot tun~elwith the regular sl~oomponent balance. The tunnel and thebalance e.redescribed In reference 6 and 7.

Tho model (fig. 1) was identioal with the circularfuselage &nd symmetrically tapered wing model of refer--ence 1 except for the addition of the horizontal tail sur-face. For tho midwlng combination the chord line of thewing was plaoed on the center line of the fuselage. Yorthe high- and the lo-wing combinations t~e surfaoe of thewing was made tan ent to the surface of the fuselage.

1?The

wing was set at O incidence with respect to the fuselageaenter line for all cases.

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3

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1.,-..,,..-..-nhe.3:1 gymmetr-ically tapered ..wlngused..!.%.t.,~e.$e???

was previously used In the investigation reported in ref-erence 10 It has an BAOA 23012 seation and the maximumupper-sufaoe ordinates are in one plane, w~th the resultthat thetohor~ plane has a dihedral of 1.45 . The Wing

tips are formed of quadrtits of approxlmatiely similarelltpses. Yhe sweepbaok of the locus of quarter-chordpoints IS 4.76°, the area is 4.1 equare feet, and the as-pect ratio is 6.1.

! The fuselage is oiroular in cross section and was1. made to ordinates given In referenoe 1. Both tall sur-~ faoes are of XAOA 0009 seotlon and have areas whioh ar-

bitrarily Include a portion through the fuselage, Asshown in ftgure 1. Yhe horizontal-tail area is 97.8 squareInohes and the span is 20 i~ohes, “whloh gives a geometricaspeot ratio of 4.1. The incidence of the horizontal tailwas 0° with reepeot to the fuselage center line for allaaaes. The vertlaal-tail area Is 63.7 square inches andthe span measured to the oenter line of the fuselage is10.87 inohee, which givesa geometrlo aspect ratio of 2.2.

The split flaps, of 20 peroent ohord and 60 peroentspan, were made of l/16-inch steel. ~or the high-wingand the midwing combinations, the flaps were out to allowfor the fuselage and the gaps between the fuselage andthe flaps were sealed. The flaps were attached at a 60°deflection.

The test procedure was similar to that desoribed in .references 1~ 2, and 3. Tests were made of the fuselagealone, of the fuselage with horizontal tall, of the fuse-lage with vertieal tail, and of the fuselage with bothtail surfaces. Slmllar tail var~atlone were tested withwlnrfueelage combinations representing high-wing, mid-wlng, and lo~wing monoplanes. All win~fuselage oombi-natlons were tested with and without flaps. The comblna-.tlons were tested at an les of attaok from

!? 00-100 to 200

with the model yawed -6 , , and 5°. A yaw range of-15° to 60° waq-investigated for most win~$uselage eombl-nat50ns at an angle of attack 2° leee than the angle ofattaok for maximum lift als0° yaw.

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A dynamio pressure of 16.37 pounds per square foot,whioh corresponds to a velooity of about 80 miles perhour, was maintained in all tests. The Reynolds numberbaaed on a mean wing chord of 9.84 inohes was about609,000. Based on a turbulence factor of 1.6 for the”LMAL 7- by lo--foottunnel, the effective Reynolds numberwas about 975,000.

RESULTS

The data are given in etandard nondimensional coef-ficient form with respect to the center–of-gravity loca–tton shown in figure 1. The results are referred to asystem of axes in which the X axis Is the Intersection ofthe -plane of symmnt.r.y of tho model ‘with a plane perpen-dicular to the plane of symmetry and parallel to the rela-tive wind. direction, the Y axis is perpendicular to theplane of symmetry, and the Z axis is in the plane of sym-metry and perpendicular to the X axis.

The coefficients for the fuselage alone and for thefuselage with tail surfaces are based on the wing dimen-sions, The coeffiolents are defined as follows:

lift coefficient (lift/q6 )

drag coefficient (D/qS)

pitchin~moment coefficient (M/qm )

lateral-force coefficient (Y/qs)

slope of curve of laternl-force coefficient againstyaw (acyiav)

rollin~moment coefficient (L/qbs )

slope of curve of rollln~moment coefffoient againstYaw ( acz/aw

Tawin&moment coefficient “(N/qbS)

slope of curve of yawin~moment coefficient againstyaw (bcn/W )

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A=”” change in partial derivatives oaused tiywing-fuqelage Interference. (Dealgnatea Incrementsof OZ* 0 Cn$s or CY*)

A= change in vertiaal-tall effeotlveness cauaed bywfn~fuaolage interference (Designate incr-monta of cl~s.cn~~ or CY*).

where

i

L

D;.1,

rolling moment

drag

lateral force

pitching moment

yawtng moment

il~mamlcFreasure (1/2 ma)

tunnel-air velocity .

air deneitr

win.c~area

vertical-tail aren

wing a~an

average wing chord .

effeative aapect ratio of vertical tail

anglo of cttack corrected to free stream, degrees

uncorrected angle of attack, degrees

angle of yaw, degrees

angle of flap deflection, degrees

6

The subscript f refers to the vertloal tails exceptwhen used with 6.

Lift, drag, and pitching-moment coefficients for thevarioue win~fuselage arrangements are presented in fl~ure 2. The valuee of “a, CD, and Cm ehown in this

figure wore corrected to free air , hut in all subsequentfi~res no correction were made.

The corroctlons were aomputed ae follows:

Aa = Sc57.3 6~ ~ L (deg. )

,

where

jet-boundary correction for wing (G.117)

total jet-boundary correction at tall (0.179)

wing area (4.1 sq ft)

tunnel cross-secttonal area (69.59 sq ft) o

r~.tio of dyna21ic pressure at tail to free-streamdynamic pressure; assumed to be unity

change in pitchin~moment coefficient per degree .change in stabilizer setting as determined intests

corrections were additive. .

The Iateral-stabillty derivatives for oommonentparts of the model appear in figure 3, which shows theend-plate effect of the horizontal tail on the verttcaltail.

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. ... .,Another method illustrating the effect of the hort-so-ntaltail on the vertical tail is to tr”eattihb”in--ereaaed effectiveness of the vertical tail ~a an increaeeIn effective aspeot ratio, as was done in referenoe 6. .This method oan”be used by employing a relation betweenaspect ratio and slope of the lift ourve. A formula forgiving this relatlon that gives one of the best agreementswith experimental values is given in referenoe 8.

Aa= a. ~

When solvod for the aspect ratio A, this equation gives- effective aspect ratio which will be termed Ae

Ae =2a

a. - lEa(1)

where

a. slope of lift curve for infinfte aspect ratio(0.1 por degree is a representative experi-~eutal value for en EACA 0009 airfoil)

a sloge of lift curve for vorttcal tail (a =

~cSf %f

per degree, where the crhltrary

selection of -.Sf is shown In fig. 1)

x ratio of seniperimeter to span of an elllpticplate of aspect ratio A

!l!hevalue of WJf was obtained direotl~ from the

force mea~~saents and also indlreatly from the yawin~moment measurements l)yuse of the equatioq

&C%f ‘-cwf Tf (2)

whese

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8

...

Cwf = c%(model with Pertioal tail) - CnW (model with-out vertical tail)

1) model wing span

lf modol tail length arbitrarily chosen to be measured.from uodel center of gravity to aerodynamiccenter of verticai tail (fig. 1) along X axis

In these computations the dynamic pressure at thetail wzs assumed to be equal to free-strem dynamic pres-sure. The values of Ae obtained by the foregoingmethod are given in figare 4. The increase In effectiveaspect ratio of the vertical tail caused by adding thehorl~ontal tail is shown in figure 5 as a ratio of ef-fective as~ect ratio with and without the horizontal tail.The theoretical value of the ratio, computed by methodspro~ented in reference 5, Is also shown In figure 5 forcomparison.

Inamuch as the results given in figure 5 are pre-sented as ratios, they are believed to be valid for anyreasonable methods for obtaining tail area, slope of thotail litt curve, tail length, and dynamic pressure ntthe tail.

Zhe i~crements of partial derivatives with respectto tne angle of 3“&Lwof rollin~moment, yawin~moment, andlateral-force coefficients A= due to wing-fuselage ln-torferonce and La due to wing-fuselage interference onthe vertical tail aro shown in figures 6 to 11.

Yhe iocromont k= is the difference betwoon thoslope (c

% ‘ c%’ andCY*) for the wing-fuselage combl~a-

tlon with”the horizontal-tail and the sum of the slopesfor the wing sailfor the fuselage with horizontal tail, .each tested separately. !l!hueAl is the change in

Cw’

‘%, and Cy cnused by wing-fuselage interference for

tthe model WI bout the vertical tail.

The increment Aa is the difference betweem. theslope produced by the vertical tail with the wing hnd the

- .-. sla~e.~roduoed .by,th.evertigal tail without Ijhewing.!CheInorement A= ifl,therefore, the ehango in effec- -

t%veneem of the vertioal tail oaused by the addition ofthe wing to the fuselage. The slope for the oompletemodel may be obtained by a summation of the slopes fortho oomponent parts and the Increments oaused by inter-ference. If, for example, tho value of

c%for the

complete model is desired, the following equation may beused:

c%= Cnw (wing) + CnW (fuselage and both tail surfaces)

+ AICn$ + A=CnW ,

Values of CI* and CY$ for the complete model may be

obtained in a similar manner.

The ~alues of CIW, CnW, and CyW used to oompute

Al and A= were obtained from temts at -5” and 5“ yaw

by assuming a straight-line variation between thosepoints. Thi~ assumption has been shown in reference 9 tobe valid except sometimes at high angles of attack.l!atledsymbols on the ourves of figures 6 to 11 Indloate ‘values of slopes measured from curves of figures 12 to 15.The arrovs in figures 6 and 10 indicate the direction ofdlvorgence after the stall.

The laternl-etability characteristics of the compo-nent parts of the model at high angles of yaw are givenin figure 12,and the characteristics of the three winefuselage combinations with”various tail arrangements athigh angles of yaw are shown In figures 13 to 15.

DISCUSSIOIV

General Comments

The lift, the drag, and the pitchin~ntoment ooeffi-aients of the several model combinations are shown infigure 2. As is to be exp~cted, the higbing oomblna-tions have more statio stabillty in pitch than the low-wingcombinatiens. Inasmuch as the teets were made withcutwing fillets, the data for the low-wing combinations show

10

the effeet of the burble d the win~fuselage juncture(re$erence 2). The:pitch ing-moment-ooef fioi.ent curvefor the midwing combination when &f = 0° has a hump

between -2° and 6° angle of attaok. Comparison withfigure 6 of reference 1 shows that the hump probably isa wln~fuselage effeot and not the offeot of the wing ..wake on the horizontal tail.

Lateral StabllitT at Small Angles of Tnw.“

# Component.p&ts.- The win~alone data ~l~en in “

figure 3 were taken from Sigure 3 of reference 3, whiohgives also a brief discussion of the wins aerodynamiccharacteristics.

The addition o% the horizontal tail to fuselage alonehao verr little effect on lateral-stability .characteris–tics, but the addition of this surface to the fuselagewith vertical tail has a pronounced effect (fig. 3). ~heeffectiveness of the vertical tall is increased by theend-plate effect of the horizontal tall. This increasedeffectiveness is shown in the Cl

w’%$s .and CY$ cur?e~

‘ “(fig. 3). The inorease “in effective aspect ratio of the.. vertical tail resulting from”the presence of the horizont-

al tail, computed from equation (1) , is shown in figure 4.Figure 5 shows the ratio of effective aspect ratios withand without horizontal tail for comparison with theoreti-cal valuo talcenfrom reference 50 A considerable varia-tion of end-plate effect with engle cf attack is shown.Inasmuch as the results i~clude the interference betweeuthe fuselage and the vertical tail as well as the end-plato ei’feet, it is not certain whether it Is the end-plate effect or tho fuselag-tall interference that varieswith angle of attack. Although the presence of the fuse-lage represents the practical case, few airplanes have aoroso-noctional area as large at “Ghetail as that repre-sented tiythis mc~el; “hence, for theoe results on exaggera-tion of fusalage--tiailinterfere~ce 13 to ks expo~ted,whatovor effect ths intorferenco ma~ h~ve. ir C;pplicationto design ths angle of attack as given in figures 4 and 5should be considered tail bngle of attack rather thanangle”of attack of the airplane.

New data were taken for all fuselage-tail r~sultsbecause.~oor correlations resulted when an ahteupt was

made to compare the-fuselage-tail data of references 1,— “ 2’L-’-&nd‘3--withthe addi+lonal data.talzea..forthis report.

Mot only were comparable fuselage data taken Urid-e&thesame conditions, but 81s0 an improved prooedure for testsand an improved”method of measuring the yav angle wereused. This fact aocounts for the d.ifferenoes existingbetween sqme’o= the fuselage data In the present reportand data in previ Qus reports of this series.

s~ ith horizo

ti.- .T.hevalues of AZCI, 8 Al% g ma A1CkV wm,e

changed only small amounts # $y the a ditlon of the hori-zontal tail to the w3ng-fuselage combinations , as shovnby a comparison of ~lgures 6, 7, and 8 with figures 4, 5,and 6 of reference 3. In general, however , the interfer-ence was decreased.

The wing-fuselage interference with horizontal tall “in place contributes a%out 2° effective dihedral for the

. high wing, 3/4 0 for the mldwtng, and -l~” for the lowwing (in fig. 6, a Cl

$of 0.0002 being considered equiv-

alent to an effective ihedral of 1°, reference 9), withflaps retracted. With flaps deflected 60° the effective -dihedral is increased 1/2° to 2°.

The values of Aicn*

are, In general, negative;

therefore the horizontal tail Increased. the weather cockstability (fig. 7). With flaps deflected 60° the valuesof AzOnW are also negative and, for the low wing, the

tendency toward veathercoek stability is considerably

L

inoreasid.

The values of ALCq with flaps retracted are posi-

tive for the wing tn the”htgh and low positions but nega-tive for the wing in the midposi%ion (fig. 8). The wtngaats as a modified end plate when in the high or lowpositions and thus increases the side force produoed bythe fuselage. When flaps are deflected 60°, the valueof AzCyW is nearly zero except for the low-wing combl-

natlon, for whioh it is more posttive than when flaps areretract eclm

Mfec t of wln- fusela~e Interference on vertical taalwith ho?i~ontal tail dn ulaae.- The increment A=CJI. tS

*

12

rather small and erratic (fig. 9). The”values sta? withinabout 1° effective dihedral for the unstalled range ofangle of attack. This result Is In good agreement withA=C Z

vwithout horizontal tail (reference 5).

The increment‘acw

is negative (increases we&th-

ercock otability) for the low wtngs but becomes less nega-ti~e (less weathercock stability) as the wing is moved upto the aiddle position and becomes positive (decreaeesweathercock stability) for some of the high-wing conditions(fig. lo). This same trend for AaCn~ is true for thehorizontr,l-tail-off condition, byt it should be noted thatthe difference between numerical values due to verticallocation of the wing IS only about one-half as”great whenthe horizontal tail is In place.

Tho reduction of the difference in the interferencebetween high- and low-wing models 1S agaifiapparent inAaCy

*when the horizontal tail is present (fig. 11).

The early break In the low-wing curves of blcl~ #

A1%WS anfiAaCnW for 8f = O“ at about 10° angle o;. attack is caused b~ a burble developing at the wine “

fuselage juncture as explained in previous reports of thisseries.

L’ateral Stability at Lnrge Angles of Yaw

&selajae and tail combinatioas.- Although rathererratic, the Cl curves (fig. 12) are cons~stcnt la thatthose combinations which have weathercock stahllity havemore effecttve dihedral at low angles of attack; whereasthose combinations which do not have weathercock stabilityhave more effective dihedral at high angles of attack.Although the horizontal tail improves the effectivenessof the vertical tall at small eagles of yaw, a more suddenbreak occurs in the curves at angles of yaw greater than10°; therefore, at large angles of yaw the vertical tailis less effective when the horizontal tail is present, assho~ by the Cn and Cy curves.

uefie mow.- The addition of the horizontal tailto the complete model has cnly small effects on Cl com-pared with the effeot of a change in the vertloal position

m 1

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of t+pewing or the addition of tho vorticd tall (fig,-13)*The offect’of thos”b-modol chan”ges‘on’ C~- haS pro-

vioualy boon disouesod in roforenoo 3..

Who curves for Cn with the wing present (fig. 14)show again that the end-plate effect of horizontal tallon the ~ertical tail Is detrimental to the restortng momentIn raw for ap.glesof yaw greater than about 25°. The ad-dltioilof the wlnge to the fu’’rlage garo 8 substantialIncreaoe In tks reato~ing mozii..% in yaw at large angles ofyaw sailincrae.sedthe vreatheroook stability at small an-

.gles of yaw (figs. 12 and 34).

A pronounced break In tho C= aad Cy curves of thefuselage with the horizontal tail that occurred at a hi~hangle of attack (fig.,12) between 26° and 30° yaw vanisheswhen the wiuR is added., The break may have been caused bythe installing cf the horizontal tail as its resul~antangle of attack is reduced by yaw - that is, the anGlo ofattack measured in a plane Farallel to the plane of sym-metry of the unyawed model. When the wing 2s pre~ont, thedownwash pl*ohablyprevents the ta~l from stalllng for anyportion of tho yaw ranga.

Curves for Cy (fig. 15) show lar~er ~alues at largean[;lesof yaw when the horizontal tall is pbsent; thiofact IU in agre,cmentwith what has been shown by curvesfor Cn with l%o wing and ourves ::011 On “~d CT -itkout

the win;;.

CO?7CLUSIOHS

!thoresults of tests of a model aonsistlng of a oir-culal*fuselage, tall Su-faces, and a wing tn high~ middle, “and lov positions indicate that%

1. The offootive aspect ra+lo of the var%~cal tail aadetermlaed from lateral force on the vertical tail waa ln-oroasod from 20 to 60 percent by the addition of the hori-zontal tail, depending on the anglo of attack.

2. Eor angles of yaw greater than about 15°, the pros-”ence of the horizontal tail decreased the restoring momentin yaw contributed by the vertioal tail.

3. The vert:cal-tail effectiveness inoreesed as thewing was moved from the high to tho low position; the low-wlng combination therefore had the most weather coak stabil-

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Ity. The addition of the horizontal tail reduced thechange in vert~cal-tail effectiveness with wing positionabout 50 poroent ~ with the r esult that the high- and Iohwing models possessed more nearly the same weather cock “Bta3ility.

Langley ii~morial Aeronautical Laboratory,Hational Advisory Committee for Aeronautics,

Laagley ~ield, Va.

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REYEHENCES-, --, .. 7.-,. -,..- .. . -.-.,.,,. L-.‘---

1. Eouse~ Rufus 0., and Wallaceg Arthur R.# Wind-TunnelInvestigation of Uffect of Interference on Lateral-Stabi.llty Oharacterlstlcs of ~our EACA 23012 Wlnge,an Blllptical and a Oircular l’uselage, and VertioalFins. Rep. Mom 705, MACA, 19410

2. Recant, Isidore G., and WalLace, Arthur E.1 Wind~Tunnel Investigation of Effeet of Yaw on Lateral-Stabillty Characteristics. III - Symmetrically TapereclWing at Various Positions on Circular Yuselage with.and without a Vertioal qail. T.I!.~Oa 825, HACA, 1941.

3. Recant, I. (3.,and Wallace, Arthur R.: Wind-Tunnel In-vestigation of flffeot of Yaw on Lateral-StabflltyCharacteristics. Ili- Syrametrically TfiperedWing witha Circular Fuselage ~~vtng u Wedge-Shaped Rear and a oVertical Tail. liLCAA.E.R., March 1942.

“ 4. Recant, Isidore (3.,and Wallace, Arthur R.: Uind-T&nelinvestigation of the Effect of Vertical Position ofthe Wing on the Side Flow in the Region of the Verti-cal Tail. !2.M.Ho. 804, l’?ACA,1941.

5. Katsoff, S., and Kuttarperl, William: The End-Plate~ffect of ~ Horizonta~-Tail Surface on a vertical-~a~lSxrface. T. N. No. 797, EACA, 1941.

6. Harris, Thomas A.: The 7 b7 10 root Wind !l!unhelof theNational Advinory Committee for Aeronautics. Rep. iiO.412, NACA, 1931.

7. Wenzi.uger, Carl J., and Harris* Tkomas A.: Wind-TunnelInvestigation of an N.A.C.A. 23012 Airfoil with Various .Arrangements of Slotted Ylaps. Rep. No. 664, HACA,3,939.

F’8. Jones, Robert T.: Oorrectipn of the Lifting-Line TheoryfOi*the Effect of the Chord. !P.E.No. 817, &ACA, 1941,

9, Bamber, M. J., and House, E. O.: Wind-Tunnel Investiga-tion of Effect of raw on Leteral-Stability Character-istics. I - your N.A.C.A. .23O12 Wings of Various PlanYoras with and without Dihedral. T.N. ~Os 703, MACA,19Z9.

.

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Fig. 1

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k~ ,~L--- MW.N cHoRD 9.84E=

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Figure 1 . -Drawing of NACA 23012 wingcn-cularfuselage and horizontalN ACA 0009 sectmn.

In cornblnqt!on withond vertcal tuIl of

iiF

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T r I I‘(a) Wing po

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-+ NA CA 230/2 wing wjf.bcircularfuselagea?dhorizontalund verticaltuils.

~1 Angle of attack, CL, deg-8 + 0~4___t__8~ --. 12 ._+._ 16 20

1

N ACA

L0- 6+ v(de@(deg)

a—Q——o CjA— Go o1.6 \0——60 5 f

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-4 m II

1 I I I

Figure 2.–Co Winued.Angle of attack, cx, deg

‘6 -8—-4 ,~o~+ 8~/2_ /6— 20

‘ig. 2b

NACA F

64= 65-. &,. ... * —.. —.. . ..—.~ * \ –.1

La

o—❑——() 5 \

1.6 A— 60 0 -.3O——&3 5

7 \

1.4- /

\

I

\1.2

P/.

2

1.0 6

/ w

OJ

X3

/

$

.8 / r

+E

I

Q) I.—c1 /.-

; ‘G 4,4

/alo

d

(J

.++

B .3

+.--1 %

.2 / /A-

/ ,2

4L f( r/

.v “

0

@

.1

4 ) ~

/ ,

/ “E m

-. 2 0

/

(c) Wing position low

-+ I I I I I -----

figure 2.– Concludede“lAnqle of attack, cc, deq \

-lg. 2C

II

—w

Lia

Fig.’ 3

I

NACA

C2+

P’li IFuselage Wing

Q Fuselage alone + fj+=oo-m Fuselage with ver+ica I tail x & = g~’A Fuselage with horizontal-tail -------- High ~~

.0 Fuse [agewithboth ta i f surfaces Middle

-.001

,00I

g

%y

—

uo

~ool

.006

()

.00+

cYyk

o

-,004

-a -4 o“ 12 ‘ 16 20Angle o-f diack, oc~ d eg

Agure d – Vimuhon of Czy, Cop, and Cy wifh ang/e of a}fuck.

NAL_AI?t30/2 WI”

7

done, fkehge, und fu~eb e WI% hi/7surf~ces. (Du u For wing from reference d

‘!d

NACA Figs. 4,5

I I I I I

From Frcnn H~r IZmtsl Tqil :..C’n Cy

❑ Q ,Off_ _ @_A= on

0’.-+p 3’0

L

+ - ._$. tJ —_ .0- 2.0 < t,

:“,, . .. . c) t)

~

~ Lo

:+w

o Ln

-0 -4 0 4 8 “12 16‘N20 ‘w

Angle of attack oti fuselage and ~horizontal tail, OC’ deg

“%%k ?$%:: a’

IIx

ct ratio of vertical tall n(EometrLc aspectratio,2.2.7

~nu

.-

2

.

.... ..

0 From CyE From Cn

2.2

L8Point campu+cd frdm(

. 4 / reference 5.El

Li Q\

1.4-

FT+Flll

.“

4A ‘-f’--’i-+-T-[ 1+FiiI,u

-8 -4 0 4 6 12 16 ZoAngle of attac K of fusetaqe and .

horizontal tall , OC’, decj

Figure5.- Ratloof erfective aspect ratio ofcular fuselage with horizontal tall to that

VOrtiCd tail on the cir-without horizontal tail.

,.,’

NACA

—.-

i,...

1;i .:

-,, ,1,

.. ,,

m

.001

., .,

-.00 I

High Middle Low tailo El b On )‘+ m x off I

.001 I I /

--,+._+._+__+_

1 !.

x- 1-%-—x-–F––

–,001 ~ I ~

(b) 6+=60: NT,

-6 -4 0 4 s /2 /6 20Angie

2of ati-ac~, ~~decj —

IIFigure 10.- Effect of wing-fuselage interference on

CWdue to vertical

tall . NACA 23012 wing with fuselage and horizontal und vertical tails.(bata with horizontal tail off from reference 3.)

.00

I+––-t–--+––

(a) 6~=’0°. 1!.00+ —

VVinq location Horizontal

a n b On.00+ + 3 x

—s..--+-”--+- -7-

-c -4 0 .g

Angle of attack, ti~~eg1

Figure 11.- Effect of wing-fuselage interference on Cy@

due to verticaltail. IIACA 23012 wing with fusela~e and horizontal and vertical tails.(bata with horizontal tail off frm reference 3.)

Fig. 12

!.

F~ure /2.– W,va}io~ of C’, Cn, und CY wi+h angle of yaw. Fuse/uPeand f use/iye wlfb lid surfaces.

—

Fig. 13

I

!>

6-Ic,-—5!Y

t’c0

?-V

.-

,04/

I /Her. OC’

/Wing / \

Iocat ion &i I (cl;g).03 _@ High /f

❑..Middje On ,.14 .- .. - /_ Q Low on 15

,, .’P

+ High off 14 / /~ ‘m/ //, “

—x Low off I5 /.Oz A ~

_/ - “(Y ---- ---= .-

Ved.tai I on /“/

_-----Vet-f. tail off /~

.01

0

-,0 I

pI a

m

--,OZ 22,1

-/ 6 –B e /6 24 32

A“ng Ie of40 48 &

yaw, +’, deguo

(a) 6f = OO. zFigure 13. - Variatton of rolling-moment coefficie t wl

NkCA 25012 wing with fuselage and horizontal an~ verk!cltwfails. u

,03

,02

.01

0

-,0I

-,07.

[

Wing Her. OC’location tail (deg) -— .-

~ High 13❑ Midd[e (% 13

/ ./ “ .

,P’ %D Low On 12 4+ High off 13 /

x Low off IIw/

Vert.tail on//

–---–vert. tull Off /~/ d /“u-,“

/ }

. Ff-

-

/ _ _

/

, m-. --G- --_a

“- -b---w .

-+& b

-. ~-

--y--1

x—

I A- 1 I

17--L I

-/6’ ‘B 16 24 32 40 +8

A~gle OF ~~w, W, deg(b)br= 60°.

r’i~e15.-Concluded.”

Fig. 14

J

,02

0

-.OZ

-lo+

-,06

-,08

02

0

.02

.04

-.06

❑ Middle On 14p LOW On 15+ High off 14

-—

-i 6 –8 o 8 /6 24 32 40 48

(a) 6f= OO.Angle of Yaw, V,deg

Figure i~.- Variation of yawing-moment coefficient with yaw.NACA 25012 wing With fuselage and horizontal and vertical tails.

H~ High On❑ M[ddle OnP LOW of-l+ High Offx Low off

I Verttuil

u

11

.

--c-- -–– --.+-—--w--- -—-+-n

----A-———---- .+

cc’(::9)

13 \12 1

,

13 1 / 2II

~ “

-–––Vert.tail ~;f

-i6 -8 0 e 32

Angle of ya~, w,.ci~g40 48

(b) 6f= 60°.Figure l!+.- Concluded.

NACA Fig. 15

m

.

.32

,24

.16

.00

u

.32

.24

J6

.00

0

-,00

B“Wing Hoc

location tall~ HighQ Mlddie %p Low+ High ::fx LOLU off

Vert tui I-----Veritail

II

-/6 -e o 0 /6 24 32. 40 x48

Angieu

of yaw, W, d% o—(a) 6f = OO.

Figure 15.- Variation of lateral-force coefficient with yaw.I?ACA 2JO12 wing with fuselage and horizontal and vertical tails,

n—

w

I, x

/$Ioy%m ?jfi(d;g)

—~ High OiI 13❑ Midd(~ On 13 /’

_ Q LouJ On 12+ High off 13

— x Low off IIVert.tail on

-——

~n- - *5

-0--: =;“

-/6 -8 0 8 /6 24 32 40 40

Angle of yaw,~,deg(b) bf = 60°.

Figure 15.- Concluded.

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