Investigation of Stability and Control Characteristics of an Airplane ...

1 ..

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

TECHI\ICAL :\OTE

No. 1208

OF AN AIRPLANE MODEL WITH SKEWED WING IN THE

LANGLEY FREE-FLIGHT TUNNEL

By John P. Campbell and Hubert M. Drake

Langley Memorial Aeronautical Laboratory Langley Field, Va.

Washington

May 1947

https://ntrs.nasa.gov/search.jsp?R=19930081890 2018-03-18T21:32:43+00:00Z

~-~-------- - --

NATI ONAL ADVISORY COMMITI'EE FOR AERm;AUTI CS

TECTINICAL NOTE no . 1208

INVESTI GATI ON Of STABI LI TY AND CONTROL CffiLRACTEnI STI CS

OF AN AIRPLAIIJE MODEL WI TH SKEI{ED WING I N TEE

LA}'I~GLEY FREE -FLIGID' TUNNEL

By J ohn P . Campbell and Hubel't M. D:cake

SUi: 1M/illY

_'ill j.nvestigation to determine the otabili ty and. control char ­acteristics of an ai;:-plane model with a skewed 1vlng has been made i n the Langley f r ee -fligh t tunne l . 'J..'hG 'vinC of the ni.odel 11"ae pi voted in such a .ray tha t it conl d be r otated as a uni t 1{i th respe c t to the fuselege so tha.t one sic.e of the ving \Tas Sl-Tept fonTard. and. the other sid.e svrept back . Hi th an d.rrnngement of this type the wing of an airpla..11.e could. be set at right anGl es t o the fuselage for take -off , lanclin3 J and lo~," -E::;?eed flight ana COUld. be rotated to some lar e angle of Ske1-l t o perltli t flight at high speedS .

I n the investigation, fl i ght teats , foX'ce t '3sts , and demping­in -roll tests "rere made on the mod.el vri th the wiag set at angles of skei'l from 00 to 600 • This investigation iias 0 an ex-.rloratory nature and was intended to provide only a preli minar: e.:1cl quali ta ­ti ve indi ca ti on of vThethe::.' such a o.eeign coulet b e f l own .

The resul ts of the invGstigation ind.i catecl that it was possible to sleevr the wing as a lUli t to angles as g!'eat as 1~00 ,,,i thout encOlmterlng serious sta,b ili ty EW."l.d control difficulties . At an. angJ.e of skev of 600 , hovrever , the aileron control be came l..ms8,tisfactOl'il:r weak . The ailero rolling effecti veness was not l'educec. by s::8I'Ting the wing fTom 00 to 400 because the damping i n roll decreased app:toximately the srune amount as the aileron rolli ng moments . The force t e Ats s~Wvred ·t..l-J.at for a skew- angle of 400 the ailerons produced l ar Ge pitching moments , but i n the flight tests no pitching tendencies were observed in aileron 1'011.'3 , apparently because the lift fo!'ces on the wi n produced by rolling introduced pitching moments that were equal and o~?osite to the ai l eron pitchi ng moments . The model did not eJ~ibit the undeSi rably l arge variation of effective dihed.::.:'al vTit h lift coeffici ent. that is characteristi c f vlinGs 1-11 th large amounts of svTeepoack or sveepforward . Skewi n,S

the .ling as a unit, hOi',evel~ , did introduce large changes in l ateral trim '11hi ch varied. 1fi th lift coeffi cient and skew angle .

2 NACA TN No . 12 a

INTRODUCTION

Theoretical and er'perimental investications hove sho'WIl that c0mpressib 11 ty effects on a '\-."in£:; can bEl deleyef:. 0~- slTeepinG the wing fOl'vrard or' backl-;ard. I n orc'er to obtai:l a l a:.'ce incrense in the Mach numbor' 3.t w'hich ~om:pressibility effect s occur; the nse of angles of svreep of 1}00 or more is necessary, bu_t these large ansles of sweep introduce s erious stability and. control problems at moderate and hj.e..h lift coei'ficieI~ts . For ex:!nple) 181'136 e.nGles of BI'Toep produce undesil'ably la1:'Ge variations of effocti ve dih3dr31 and pi t chins -momont coefficient "lith 11ft coe~ficicnt . Also a rapid increase in drag OCCU.1'S at motierate lift coofficients ",hich is detrimental t o take - off and climb perfOy·w.1l:lCe a...l1U which compli­cates tho l anding probl om .

In ordor t o cain the advantages of S~'Toe:l at hi6h speeds without experiencing the d.iff ict'.lties introduced 'b~" 8':1Oup at 10'l'! speeds J it has bee::1 proposed t ha t an airplano be 6q,uiPP00_ v;:l.th a ",-:tng pivotally att a.ched to the fuselage so that it can be s e t ot rie.,ht angles to the fuselage for take -off J landing) and lOi~-spoed flight and a t some angle of 31.,rOOP for flight at hi{;h speeds . In one suggested design the Hing is s}:ewuc1. or pi voted 8S a unit so t hat oni3 sid.o of the wing is s\·wpt fon/erd and the other sicte SI-TOpt back . In order to QscortDin the low-speod stability and c')ntrol character ­istics of such a ctOSir:;rl ::m inv )stigation hes boon concLucted in the Lanelcy froe-flig..~t tunnol. This invostigntion c oneistod in flight tests, forco t osts, and dampinG- in-roll tests of a IDodel equipped '\-ri.th 1:1 pi voted vTing that could bo set at variol'.s 81161 s of skew from 00 to 60 0 .

This invosti ga tion vTas of an oxplorator 3' na turo and "JaS intended to I'rovide only a prolimi.nary and qu<-< li tati.ve in'~.icQtion of whether such a desien could be flOl'T!1 . No attempt is thol~ofore made in this paper to give a complute and cOID::?rehonsi vo c1..isc~s:Jion of tho. sta ­bility and control problems involvod in SkO'Ioloc'··vTinc; designs . Tests at higher sca le of skm'Tod-vrlng moduls more r opros.:;ntat1ve of high­speed airplano dosigns will pr obably bo n ooded poforu an acclrrato and detailed analysis can be 1llCdo of tho stabil:i:ty ;IDd control characterj.stics of this typu of airplane .

Sl11BOLS AND COEFFICmNT8

Th0 forcos "ond coofficionts i1OrO moasu.ror.l 'vi t} roforonco to tho stabili ty axes . A diagram of those GIGS Sho -yri!1: t"_c posi ti V0 diroc­tions of foJ;'cos and momonts is presented QG fi sv.re 1) "Thich sho,\·TS the vTing in the symmetrical (ullokovcd) condition . For Gkowod condition tho axe s arc determinod by the fusolage r athcl' 'c!l(m -DY the wing .

NACA TN No . 1208 3

M

L

N

s -c

b

v

5 a

(_L~ft \, lift coefficient -\, qS /

drag coefficient (Dra g\ . qS /

p'itching moment

rolling moment

yai-ring moment

pitchinG-moment coefficient

rolling-moment c r)of fj cient

ya~~ng-moment coeffi c i ent

/ Lateral fo:~c o\ l a t era l-force coeff ici ent .-----.----. )

\, ' CIS I

vr1.ng area> sqnar e feet

mean aerodynamic chord of t1Ilske1ved irinG, feet

span of lLTlskel{ed wing, feet

d~mami c pressure , pounds per square foot

airspeed, f eet per second

mass density of 3ir , slug per cuM.c foot

anel e of sidesli p: decroes

anr,le of yai", c_egrees ; for force -te~t cl.J:::to equa l s -~

skevT anGle ( anel e throu:-;h vh1ch ",i<; io rot8tec: vlith respect t o t he fusel aGo )) deGreos

aileron def15ctlon) degrees

el evator deflection, degroos

ruc:der defl ction) degroes

~.

x y

z

pb

2V

p

NJ\.CA TrJ No. U:oG

angle of attack, d.egrees

longitudinal axis

lateral axio

normal 0. xi S

ra .ius of g)rrution of model about X- ax-ts , feet

r a d.h1S of gyration of mod.el about Z- 8.yis, feet

aileron-roIling-effectiveness fa ctor or helix angle generated by ,·rinS tip in rol l , r aCLi.ans

rolling fl n{jl.uar velocity) r a dians per second

effective- dihedral par aI!leterj r ate of change of rolling­moment coefficient with angle of sicle slip, per d.egree (COL/ dB)

On d.irectionaJ.-stabili ty parameter; rat e of chance of Y8l-ring-f, moment coefficient '"I'li th angle of Did.esli]? , per

degree (60n /d13)

d.':lllIpinc;- in-roll par ameter ; r ate of change of rollinB-moment coefficient I ,d th rollinG··an c:ular-veloci ty f actor

/ 0C~') / -\ d~~ .I

A1?PAl:1P.TUS AND HODEL

The flight tests and f orce tests \,'ero conducteo. in the Lun~ley free-flight tunnel, a complete d.escripti on of whi cb :L s given in reference 1 . A photograph of the tunnel tes t section in. th the model in flight is prosent ed. as fieure 2 . All force .tests were ma de on the fl'ee- flight - tunnel ba l :mcc (reference 2) i,'hich mea surGS forces and moments about. tho stabi l ity axes .

The v;:.: lues of tho d.amping- 5.n -roll o.eri v2.ti VG c~ were p detormined by rotation t esta in tho LElnl~ey 15- foot :free-spinnin.s tu~~el by tho method d.escribed in r oferenco 3,

- - - ~-------~-

NAeA TN No . 1 208

A s ~etch of the model used :i.n the t ests is pr esented as fi l::>lITe 3, and [3 photoer .::lph of the model is presented a s fi gm-e lj. . The moo.el fusel2.{se consisted of a boom on which was mounted the rotDt able w:i.ng .snd fi~:ed horizontal and vertical t ails . The 1vine vas mounted by a pivot at t he 50 -percent - chord point so that it could "be skewed t o anY':lngle up to 60° . L'l all tests the i·ring \\'8S

s kewed. so t:1at the right 3io.e "res Sr18pt b~ck and the left side Si'18pt f orward, a s s ho"\-IIl in fi 81lre 3. The dimensional and mess char a cteri stl cs of the model are rsi ven in te blo 1.

TESTS

Force tests 'Here macle to determine the stati c stability nnd a iler on control characteristics of t he model ,\Tlth ske,·; s n r les of 0°, 200

, 400 , an.d Goo. For ce tests ':-rerc also made t o det81w.ne the 101lcHudinal stp,bility of the model vn:t,h horizontal t ail off for skei; an{Sles of 00, !}.()o, and 60° . ro control- effectiveness tests i-TOre made for clevator ana. rl-'.ddcr . All force tests • .. rere made at a dynhl:li.c pressure of 1.9 pounds pCr 8~1J.are foot "Thich corresponds to a t est ROuTIlOIds nUrlDCl' of 179,000 <c! t 0 0 s kow' based on the mean aer odynamic chord .

All r olling-mom.ont and m-ling-moInent .B t u obtr..: i::-lCd in the

5

t es ts ilith d.ifferent skev! angleo are ~asod on tho sp"'n of' tho l.ms;:m·!od l;inS and all pitchi.ne -moment d.a t.a '.r e bQrJod on the mean aerodymunic chord of the Ullskouod ,ving . AE. t est da t a are r oforr ed to t he nor:nal contor of gravi t y at 20 p ercent of the mO"Jl 30rodynoDLi.c cho~~d of tho UJlskCl-TGd .vinc; unless othenr.!. S 0 notocl all the figure s . If it is desired, t he momonts can l)e based on the spFm ana. chord of the sk81'led wing by using t ho valuea g.i ven i n t ilble I f or tho differont skew an ·los.

Rotation tosts vore IDede to detormine the damp: .ns in rnll £'o!' the modol 'Hi th ske,\.; an glos of 0°, 20 0 , 400

, end 60° &t D Uft coeffic~.ent of 0 . 5 . All d3IlIpine;- in-:coll t.osts "Toro made a t a dJ~malc prossure of 3 .0 pounds per square foot which corrosponds to a Reynol.:1s ll'l.l.l!!.l)er of 226;000 at 00 sko,-l.

FliGht tests of tho modol ,vi th tho centor ("If c,ravi ty a t 0 . 20 mo::m aerod.y.namt c chQ}:,d of the 1..mslmved l~nG 'Ivere Inacl0 at a lift cooff'::' ciont of a-p})ro::dTIlr':ltely 0 . 6 for skou angle s of 0°, 10°, 20°: 300 , :.0 0

, 50°; ana. 600 • I n addition, for the same center ­ol- (;!:'crvity location, fli g;l1t tests ,vcre 1lUlde OV01~ a lif t - coefficient range fro:n 0 . 3 t o 1.0 f or 0 0 aml 400 ske"l. Furth,:;r flight t ests "lor e Il13de 'I·Ti t h 40 0 skev at lift coeffici ont s from 0 . 6 t o 0 . 9, Qnd tho center Oi Gr avity vres movod suc cessively back t o 0 . 25, 0 . 30,

6 NADA TN III o. 120':

and 0 . 35 mean aerodynamic chord of the unekewed Wlng . In the fllgpt t ests, abrupt deflections of approximately ±18° aileron ( total 360 ), ±5° rudder, and ±5° elevator were used for controlling the model. P.eference 1 describes the flight - testLTlg technique used in the Langley free-fli~lt tunnel.

BESUIJTS .AND DISCUSSION

Force -Test BestLlts

Lon3itudinal stability.- The results of the force tests nwde to c.e'feriiilIi.e- the loneitudinal stability char acteristics of the model are presented in fiG~0s 5 t o 7 . The data of figure 5 show that as the skei" angle of t ho model '·7~~ S i ncreas ed from 00 to 60°, the static l ongitudina l staljilitY (63 i '..1dicated by the s lope of the pitching-moment curve ) progres sively decrea sed u..l1til at 600 skev the mode l "Ha S longi t udi nally Ul1stable at lif t coeffi ­cients abo\'e 0 .'7.

The data of fi gure 6 show th1;l't "\vi th the horizonta l t ail off the r eduction in longit udinal stability with increasing skew lJaS

even more pronounced. A cOl.'lJ;Parison of the dat a 01 figures 5 and 6 indicates that the add.i tion of the hori zor;.tal t3.il groatly reduced the voriation in 10nc:Ltudinal stability over the lift r engo fo:c the 400 and 600 ske"\\red 'dngs .

Tho data of fi gures 5 and 6 ShOVT ·the e:x;pecte(l r eo.uction in lift - curve slope vTith increasing slec"\". The slopes of the lift curves are ap:proximateJ.y proportional to the cosine of the angle of skew.

The force - test da ta of fi gure 7 show the longitudinal sta­bility and trim chara cte:ci ctlcs of the model 1dth J+Oo sk(3"\v for the most f Ol'"W3rd and r oar ·rv. rd center-of - gravi ty locations usod in fli ght tosts . These data sho'\', that for the most r earward center ­of- gravity location tho model vms either statically longitudinally unstable or about neutrally stablo over the entire lift-coefficient range .

Later a l stabllity.- The results of t~e force tests made to dete:r.mine the later3l - stability chara cteristics of tho model at a lift cooff _Gient of 0 . 6 at z ero anGle of ya \'T ar e shov711 in fi f,'Ure 8 . l 'he offects of eke" on tho directional-stability parameter Cn13 and on the effecti v o dihedral .para.meter C~ 13 as determined by the

slopes of tho curves bati-Teen ±100 in flcvxe 8 can be summ<:;rized as follows:

NACA TN No . 1208

81m if angle

0 .0034 -0 .0006

I .0033 ' - .OOO~

__ ,-_ __ • 0_0_2~ ___ 1_-=.' 0006

The values of C ~ p f'or the model are consJd0rallly smaller

than the values norma ._ly encolli"1tered. a t 11 lift coefficient of 0 .6 on wings "liti. l ar ge 8!lountS of S1{eep~8ck .

The lateral-tri.m changes c/':used . y sko'Fin,g the i·dnB from 0 0

to 40 0 ena. 60 0 can be seen from the dcta for zero c.n~le of ;yavi in figure 8 ~d f r om tl 6 laterel - con~onent dat n of f::.gu;'es 5 and 6. 'L'he data of figu.re 8 for a lift coofi'icient of 0 . 6 sllo'\,- t h ·.., t u:i. th inc:ceasin3 SkSI-T an increasing positive (right ) l E:lteral force and an increasing negat,j.ve (lef t) r ollins ornent occUJ:-red . Changing t:t.e ske\~ from. 0 0 to 400 cAus ed a neE,ative ( left ) ':..i·ring moment but a f rrther i ncrease in ske".T t o 600 :proc.uc'3d a posi ti vo (rieht) ya'\dn g nomont. The 6..ata of figul'cS 5 ... ,nd 6 sho';7 t· mt th3 c an3es in lat eral t r im varied considcral ly .lith engle of att6ck .

I f the pi vat point of t he \. nl} ,lers shIfted forvT~rd the chan ge in rollins moments vr.i. th ske.; at lOll lift coefficiOl ts ifould be rcd wed because the ar ua of the l e: t '\Jin3 '·.'ould bo i ncr eased vrlth increasin,3 skOi,~ .

LnterEll .control. _ - T 10 results of the for ce tests B :ldo to det ermine the a ileron effectj.vencss aro shov71l in figures 9 and 10 . These results show t ha t t he a iloron effectiveness in producing rollins mOIllent was somowhut reduced by slcovTine; tb.o ,d.n6 to 1~00 ane:. "TaS e:coatly reduced by Sb:liv:1.ng tho w1,nc to 600 • "w110n deflected the ailerons at the 1\.0° and 600 olco~·; aJlt;les also ? T O ucod. s1z:3ble pi tchin13 moments . Thei telling m.omont :p:roduced by the a.ile:..~ons ,,;as rJ.uch greater at .00 SkOlv Lngie t han a t 600 , appar ently, bocm:so of the reduced effectiveness of tho ailerons in :producing l :lft at 600 ako'li.

The da ta in f i (S1.1.re 10 J which 8hm'T tho independont contrj.1JU­tions of t he r i8-ht EU1d lol't a ileron s t o tho cor d;ynamic m.omonts, incliceto that the l oft ( loading) aileron was most affocted by skew 811g10, and th:'!t a t )~OO and 600 skew its offecti venoss ::'n producing r olling moment was somewh:3 t l ess than that of the r:tght (trGiling)

8 N!-\.CA TN No, 1208

a ileron. This difference can be attributed to the increase in "i:-lnG area ahead of the trail.ing a~ leron.

Rotation-Test Results

The resllts of the damping-in-roll tests are shown in figure 11 tOlScther ,d th the ro11 inG ef'f'ecti veness of the ailerons based on the rolling-Llomcmt data of figure 9 . These d.ata show' t hat the damping in roll "?as r oa.uced by skew t o such an extent t hat the ailerons remai:!lec eff ecti ','e to eng.les of skew gr'eater than 400 •

The clalJl1)ing-in-roll ' c.a t a b i.A s ecl 0:1 the projocted span at ea.ch IJke~T anglG (rather than on the spqn of the i·'J. ~lv .. 71th 00 ske .. ,) are presented in fi g'JrG 12 for the rmrpose of com:purinr, the data 'in th ce lcclations made 'oy the simple r e1.::.":.ion I'~efJe~ltGti. in reference 3. This fl cut'e ShOH8 that JC:1G dan;ping in roll v[.ried cpproximately a s the cosine of tho S};:O"'l·l an ::;le, as WDulct be expected freID. the data of r(;ference 3 for cO!lventional1.y svTOpt -back idnga .

FUght-Test Rosults

In tho fU ght t ests of the model .. ;ith t he contor of gl.'uvity at 0 . 20 meon aorodJ:1a:.m.c chord the genoral flight cho.racteristics VTero sCltlsfactory and. rcrnn.1ned oss oriGially unc.13n30d. a s the "rlne ,vas SkOi-T3d froLl 00 t o 400 by 10° -inc:;:'oL'lonts . \-lith the ir:ing skewed. 500 tho flight chara ct eristics of tho mecLol i'Tor o se.tisfactory ey.cept t hat the ileron effectiveness , :as noticeably reduced and somo diff iculty "'I·TaS conso~uently oxper:tcnced in controlling tho model. ii Hh 600 SkO"'li the aileron offectiveness "\',8S even further reduced C).nd .. ras inado~uatG for :rnaintain1ng sustained :flights of sufficiont l ength to POJ:'!l1it j udginG the ot or stabiHty and control charectoristico of the model. Thes e flight-tost r esults nr o in agreoment with tho f orco-test r osults of ficures 10 to 12 in regard to the reduction in aileron offecti veness with 600 ske • .,.

Io prono'.mced chc:m~o s in stability and. control vore apparent with ske"1 cm glos up to 1.:00 but sizable ch8nges in l ateral tri m -v!Bre noted, 1-:hon t ho 8kew angle "as increased f r om 0° t o 1~00 while tho fli Cllt lift coefficiont 1-18S held cO:J.sJwnt at ab01..t 0 . 6, uso of a t0ta l of about 170 riCht aU.cr on trim and 30 l~i eht J.'ud.dcr £lotting was necossary to rnaint3in l.r:Ger al trim, (t .... lOt i s, to koep tho vTing loyel and. tho fuselaGe at z oro sideslip ). As the fli&~t lift coef­fici ent ,':as increCt seo. at 400 ekev!, howver, :progl.~esfJivel.y 8Il1:ll10r amount.s of [.dleron and rudd.or trim were re~uircd 1.1Iltil at a lift coeffici ent of 1.0 no trim "WUG needed. Thoso trim ch3nec13 lIDrD

indicated by t ho forco ... tost rosults of' figures 5) 6, anc:. 8 .

NACA TN No . 1208 9

lihen the skew angle was chanGeQ from 00 to 40°, only slight changes in elevator setting (not over 30 ) 'vere required to maintain the same fUght lift coefficient . itli th 40° s1;:ew , hm-rever, the glide -path angle "Tas 2 0 or , 30 higher than wi th 00 skew over' the lift -coefficient "-'ange from about 0 . 3 to 1.0 . Thi s result is substantiated by the fo"-'ce -test Qata of figure 5 "Thich indicate

'that the Qrag at 4c o slee", is higher than that at 00 ske',T at a gi ven lift coefficient . The force -test results in fignre 5 8hmr that at lift coefficients Greater than 1.0 the drag at a giv'en lift coefficient (and hence the glide -path angle ) f or the 40° skew angle became increasingly. gl'ea ter than that of' the unslceiveQ ,ving .

In the flight tests of the 400 skevred 'lorinG satisfactOl'y flights were made over a l ift -coefficient range from 0 . 3 to 1 .0 with static margins (static 10ng5. tudinal stability ) from lar ge values t o very small values . The fli@1t characteristics appeared to be slightly better at the ImTer lift coefficients . No pronounced changes in the longitudinal stability characte"-'istics viel'e noted as the static margin was progressivel y decreased by moving the center of gravi ty from 0 .20 to 0.2) and. 0030 mean aerod:Jl18Jl1ic chol'Q . 'I{i th the center of gravity at 0 . 35 mean aerodynamic chorQ, however , the model appeared to be lengituc.inal ly unstable and continv.ou8 application of elevator control Ifas r6quired t o keep the mOQel flying . The fo"-'ce-test d.ata of figure 7 indicate static longi ­tudinal instability for the forogoin6 conQition .

In the fli sht tests no pitching motions 'with aileron control wel~e noted for any sko>-r angle , lift coefficient , or center -of ­gravity location. This result appears to disagree "lith the force ­test results of fiGUres 9 and 10 vThich showed 8i zable pitching moments ,vi th aileron deflection for lj·Oo skeYT . This apparent Qiscrepanc;y- is explaineQ by the fact that Quring a steady ail eron roll , lift forces Que to rolling are proQuced "'hich are equal and opposi te to the lift forces proQuceQ by the ai1e:.:,ons . These lift forces due to r olling produce pitching moments that arc equal and opposite to the aileron pitching moments and hence oliminate the pi tching tendencies in a steady a iler on roll . No flights were made to determine the effects of these a ileron pitch:i,ng moments in a steady sideslip .

CONCLUSIONS

The results of the investigation in the Langley free -flight tunnel to Qetermine the stability and control characteristics of an airplane mOQsl wi th a skevled wing are surmnarlzed as fol1mrs :

10 NACA TN No . 1208

1 . In general, the results indicate that an airplane vdng can be skewed as a lUlit to angles as great 88400 w1.thout encountering serious stability and control difficulties .

2 . Longitudinal stability and control:

( a ) The longitudinal stability and control charactor- . i stics were satisfactory--in flights mad.e ,vi th 40° skew over 8 lift­coefficient ranee from 0 . 3 to 1.0 even for very l ow values of static margin .

(b) Only a slif.~t chullge i n longi tudinal trim occurred. 11rith increasing skew' but an appreciable increase occurred :i.n the glide angle reqvJ.red at a g:tven l:lft coefficient .

3. Lateral ste bili t :r:

(a ) The values of effective d.:i.h~al for the wing skewed as a unit were consideraoly l es8 thon those encountered on 1-lings "dth largo amounts of sweepforward. or sweepback .

(b) Sleewlng tho wing caused. sizable Cr.8ngeS in t he latoral trim ,<lhich varied. with lift coefficient and skew angle .

4. L~tGral control :

( a ) The aileron control effectivoness Wa S only sligJ.tly red.uced by skew for angl es less than 400 because the dampinS in roll decreased opproximately tho se.me amount as the aileron rolling moments. At 50° skew, however, tho ailo:ron coritroi effectiveness "res noticeably r educed, and e t 60° it was so weak that sustained fliGhts could not bo !Jl.&de .

(b) The force tests indicated that for 400 Sltew angle the ailerons pr~:)(luced large pi tching moments . In the flight tests , howeve r , no pitching tendencies were observed in aileron r oJ.ls, appa~Gntly because the lift forces on the wing produced by rollin3 introc1ucedpitching moments that were equal cmd oppOSite to the aileron pitchil13 moments .

La.Tlgley Hemorial Aeronautical Laboratory National Advisory Committee for Aeronautics

La.Tlgley Field , Va ., July ~3 , 1946

NACA TN No . 1208

REFERENCES

1 ~ Shorta l, Josep 1 A. , and Osterhout , Clayton J .: Preliminary Stability and Control Tests in the NAC/\. Free-Flight Wind Tunnel and Correlation with Full-Scale Flight Tests . NACA TN No. 810, 1941 .

2 . Shortal, Joseph A . ) and D-£'aper , John \{ .: Free -Flight-Tv.nnel Investiga.tion of the Effect of tlle Fuselage Length and the Aspect Ratio ancl Size of the Vertical Tail on latoral Sta ­bili ty .and Control. NACA AIm No. 3D17 , 1943 .

11

3. Bennett, Char les V., and Johnson, Joseph L.: E:>..-perimental Detel~nation of the Damping in Roll and Aileron Rolling Effectiveness of .Three Wings Having 2° .. 42°, o.nd 62° Sveepback. NACA TN No. 1278 , 1947 .

NACA TN No . 1208

TABLE I

DD1ENSIONAL AND MASS CP'AHP.C'l'EHISTI CS OF MODEL USED

I N SKEWED -1trING I NVESTIGATION

Weig...ht, Ib • • 4.73 to 5 .03

Wing: Area , sCi ft Span , ft

0° ske1'T 40° skew • 600 skeVT •

Mean aerod;ynamic chord , ft 0° skew 1~00 oke'" • 60 0 Skel{ •

•

Aspect ratio (0° skew ) •

•

SI'ieepoack of 0 .25 -chOTeL line, deg Dihedral, deg • Taper ratio (ratio of tip chord to Root chord, ft (0° skmoT ) Tip chord, ft (00 sJ:ew ) Loading , 10 per sq ft .. , .

Radii of gyration (for 0 0 skev ) : , ft KX' . kz ' f t .

Ailerons: Type . Al'ea

Sq ft Percent S

Sp<:l.n, percent 0 '.

• ,

•

.root. ChOl'OJ

2 .6'7

l~ .00 3·07 2 .00

·70 ·91

1.40 6.0 3·0

o 0. 50 0 ·90 o . 1~5

• 1 . '(8 t o 1.69

•

•

nATI ONAL lllVISORY COHHI 'l'FKE FOR AERONAUTICS

0.625 0 . 844

Plain

0 .19 '7

44

I

NACA 'I'N No . 1208

TABLE I

DII~NSIONAL AND MASS CHARAC1~RISTICS OF MODEL USED

IN SICE'w"ED -HI NG Il!TVESTI GATIOllJ - Concl uded

Horizontal Tail : Area · . . . . . . . . . . . . . . . . . . . . . . . . . . . . · . . . .

• • • • f • • • • • • • • • • • • • • • • •

Sq ft ••• Percent S •

Aspect ratio Tail length , hine;;e line t o center of gravity , ft ••••

Verti cal tail : Area

Sq ft ••• Percent S •

· . . . . . . · . . . . • • • • . . . . . . · . • • • • . . . . . . • • . . .

Aspect ra tic .. • ...•...•...••...... Tail l ength , h inge line to cen tel:' of gra v i ty , f t • • • •

NATIONAL ADVISORY COMMI1~ FOR AERONAu~ICS

?O

2

o .!~ 15

2 2

13

NACA TN No. 1208

x

Wind dJrect,on

y

\

I

~¢ MI

~ - --_ ----r_-rl

-."

Fig. 1

y

z

z NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

Figure I. - The sfabJldy system of axes IS defmed as an orthogonal system of axes having fhelr ongm at th~ cenfer of qrovlfy and In whIch the Z- aXIs IS In the plane of symmetry and perpendlcu/ar to the relafwe wind) the X-axIs /s In Ihe plane of symmefry and perpendIcular to fhe Z-axIs, and fhe Y-axIs /s pl!rpend1cuJar ta fhe plane Df symmefry. Arrows indicate post/we dlrechons of momenfs) forces J and contro/-surfoce deflecflons.

NACA TN No. 1208 F ig . 2

Figure 2.- Model with skewed wing flying in Langley free-flight tunnel. 40° skew.

NACA TN No. 1208

I

Skj-V C1.n9 /e~/\ (vane d from O'to 60°) ",

3: " 0.50 chord ",

_~"~r-~

/Vormal Cenfer-oF-gra VI t)' location ---_ __ --=.../

1"-----Z4---~

NATIONAL ADVISORY COMMITTEE fOIl AERONAUTICS

F lgure3. - Three-vlew s lr e fch of modC?/ used In ad)tJs ta b/e-.5 tewed-wl/ly I nye5!190!IOIl. (~/I d/!l1el!s/ O//5 /// 1/)C~tf>5 .)

Fig . 3

J

-- -- -- -- ~----

Figure 4.- Three-quarter front view of model used in skewed-wing investigation. 0 0 skew.

~ o >-1--3 Z

~ ,...... C\:)

@

~ 1-'. aq

f.P.

- ~.--~~--.~--~-

L

NACA TN No. 1208

.2 -+-

~ ~ ~'-5 s:::5 ""' 0 C:-+-.. I~ ~D .c:::~

0 :- _

~:-~ ,-;:-~, y - ----< , :...,-~ ~~

"t ~ -', ,y-_A' , [J~~ '~': ::-' ;r- -'

~ ?\

'~ r.l. , " -c~ ~

,-2 ~

--..;, - --<: '

ct(J

.8

& ,6 ""' -+-.. ~ (j

~4 § ~.2

C3

o

~'

-,4

1.6

(jI.Z 0 / $

~-) (,ynl . --~ :~.;;J

.~

.~~~~- .

+-§3 (j,B

s;::: c;..::

.If /~~, V" /)~

PI /1 J ~/ II / ,

~ 10

~ .4 ~) I /

~) /

-.......J (V/o/' ,~ .~ r-( ,

~:r=u o

o

_ 'U' .~~ .

8 /6 ' 24 32 Ary/e of attock) CX, dt:9

.r o fd7J) 0 - - 20 o ----- 4(J L'::.- -- 60

F ig . 5

NATIONAL ADVISORY COMMITTEE FOR AERONAUT ICS

40 .2 0 -:-2 P/tch/[Iq -moment coeffICIent) em

-4

fiq(Jre 5 - Effed of sAew anq/e on the aerodynamIc character/sties of the comp7ete sket--/ed - NIf)9 mode/

Fig . 5 cone.

4-sue ~--... ~~ 0 I CJ fS'-~~

~~ -.4

.0 4

1::>.::: §~

0 §a3 <:)

'e:+::: g''t-:0-.. 0)

~8 -:04 ::>::

-:08

.C6

-.08

-fl.~ .s=-~~ .>=-= ~ .):" - -l'r -, .:.-- - "t.

~ .~ --<->- - -> ..:.' L:.J

.y--<r-D_"'~ ' , . F='Q ~~ ~-Q F~! ~ '\ - ' . ..:c&:-o

'~ }, 1\ '~>-~

f'-- - ...

--<: -- y - :>--.._.., t/ ~

/ /'

~~ /i> I-----i )---;j ' -I--==(:J==(

v..:..., 0 I-O~ .~ r\ < ~-=

NATIONAL ADVISORY C01414ITTEE fOR AERONAUTICS

o 8 16 24 32 Af!q/e of ot/ock) ct) deq

hr;ure 5. - Concluded.

NAC A TN No. 1208

40

r (deq)

0 - -- 0 0----20 0------ - 40 6--- - 60

NACA TN No. 1208

o 8 AS 24 32 .4 Angle 01' attack." ~ deg

.t (deg)

o---() 0------40 t:,. ----60

Fig. 6

. 2 0 -.2 PlfchlnQ -momed ca:llie;-ent~ em

Dgu re 6. - E lfeel 01 sKe W' angle on tfJe aerodynamic chara.cter/,s/Ics 01 the sl:ewea-wlng model with the: hon zontal tail removed . NATIONAL ADVISORY

COMMITTEE fDR AERONAUTICS

Fig. 6 conc .

4

,08

-.08

.~-~ ~--=-< :~-=< 'r--< . - --; 'r - -..? -4. -~ . =='. ~

'r- -_( '7-- - .>- ~ ~

Ir -~ '>, .-(. .-

-:~ / /-l'

~

d /

~

.~-~ / < ~=- .--'j ~~ ",/

~.

NATIONAL ADVISORY CO ...... ITTEE FOR AERONAUTICS

I I

o 8 16 24 32 Angle of' a tlac/:., CC, dc:g

figure 6. - COl7clud ed .

NACA TN No. 1208

j (deg)

0---0 0---40 6- - - - - 60

-~-- -- -- --- -- -- ---.-- ---.. ----_____ ~ ~ ___ J

.c.

J .1 ,--. ~ ~ \J

~ ~ 0

B ~ ~

~ -.1 ()

~ I

~ ~ :-.2 ~ ct

-;3

n a ......... ~

1......--

o

1---l-o-f---_ fa c.g. J L.--r.l

-.......... I ..... 0e (percent:

1-0 -

I-- tJ (deg) M,A.C.): - I- - \ 0- 0 20 ........

~ ~D--O 35 1'-0 , .F1-'" 0----- /2 20 ---1-0. C>---/Z 35

-0.. t---

~ ~ ~---

1-'---V~

fb" p

r:-. .... - --, v \ \A 1\ tv' \ ........ ...... ......

-- r--_ \

1"<>- -- <> Ih 1'- .........

0--' ~"

Q, .......

" I. ' ,(::

<> ) .......... /

.2 .4- .6 .8 /.0 /,2 /.4- 1.6 Lift coeFr/cleht) CL NATIONAL ADVISORY

COMMITTEE FOR AERONAUTICS

F/qvre 7. - £.ffecr of" ~/e J/a tor de,CJecf/on and cei?t~y- of'-9ra.vlfy loca;lon on th~ stahc long;/vdlna.J sta.bllity ctncl tn/?1 of tnt!. complete .s/(ew~d-wtr)9 model with 40 0 slrew_

~ o ::t> t-3 Z

~ I--' i:'V

@

f"Ij ~.

aq

-..::J

Fig . 8

.4 Q)G ~-I-..~

~~ 0 , G 1L...l.-

~~ ,f.\ ~Qj '§8

-.4 "-l

.08

10'k

-:08

-h.. .04 c: §() ~1€ 0 IQ:! ~<j ~~ -,..;qj

~8 -.04

,0

If

-.08 -30

(.

'1 L ----~

,;7->- ) --- ~--- --- tJ-

NACA TN No . 1208

r a (deQ) (deq)

0 --- 0 3 o -- -- -- -- 40 7 /::,. ---6D 14

~ ........... r- -6._·1 --<_6

.r ~ ...(A\

~;~ P'

NATIONAL ADVISORY COMMITTEE FOIl AERONAUTICS

-20 -10 0 /0 20 30 Arqle of YON I 0', dey

fJqure 8 - Effect of skef...J anqle on the loterol stobillty and trim characteristics of the complete skewed - J,Jtn9 mate/. 0. oppronmofely 0. b.

NA CA TN No. 1208

~

~ ~\0~ .J C)'1-..' ~~ 1<U

\j-,'-.. C:::\J 0 ~~ ~<u ct8 -;-J

/./ I~

/ / 4' / /' L ('h

/'

-H

6-- -y-- -/

V <t-

v

J \.-~ .~--

?t- ~

/~ R -6

/ff I~ t-- .

- .~

l;s-- - . r-- ;; L-A / . /

k /

/.

c7 NATIONAL ADVISORY COMMITTEE FOA AERONAUTICS

40 20 0 20 40 Tot-a/ aileron de-fJecflon,deq

Leff oheA RIghi sf/ex

t (degJ

8=---4~ 6.----60

Fi.1..ur e 9. - Erreo! of' oxew anqle on aIleron effecf/I/c?/7ess of' The okewed-w lnq m o de!. Bofl; al Jero17o deflecfed wllh ec;.ua/ up atfd dowl7 -{ravel. CL ::: 06.

Fig. 9

F ig . lOa

-+--. .I ~~ ~~ §t:' a 19.)

~~ ;:,,;~

~§ -I

-+..... .01 ~~ ~<:.J ~-+-::' ~c:: a IS:!

~ ~ -;01

.oz G 4-..' c: .01 S!J ~ ~ ~ Cj

0 -+--c:::

~ § ~Ol I

~ ::3:::: 15 Cl;2 :-0.2

9 9------ -k-Y -9 cr .... C}-- ---0- 6

:., .... ...{ .) !:;:f ,... ..... ' L'J

~¢ n/ .6

1/ /' V

/ / :J lV

/L V ;'

/1 / L /0

t¥-- NATIONAL ADVISORY COHHITTEE fOR AERONAUTICS

20 /0 a 10 20 AIleron deflectIon"" deq

Left s lick kJ9hl sltck

(a) J' ::: 0".

NACA T N No. 1208

AIleron O--Rlght O--Left

Figure 10. - Effect of she /AI onq/e on the OJ leron efPecf/ye-!leSS of the skeNed - v./Jnq mixiel. A!/erol15 defleded mdlVldually. CL =0.6.

-----

NACA TN No. 1208

'k

~ ~~~ .I

~'" I ~ ~ 0

~~ :-k~ ct\j --:/

"."'- " tr= t~ r--t;1

-~ A/~

/ ,/

./ ':/ (\/ W l:)

NATIONAL ADVISORY COHHI'TTH fDQ AERONAUTICS

20 /0 0 /0 20 Aileron def/ecflo/l) deq

Leff she k. Rlj'hf 5f!ck

(b) b= 40 ~ . Figure /0. - COl7fl l7Ued

Fig. lOb

Aileron O-R/qhf 0-- Leff

Fig. lOc

.2 ~ /

/ [l

/ v>-.tV fY ~'T

_R r- '1 ~

,~ -t;:J

r ~ r===8 a----: ~ v T \.

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

20 /0 0 /0 20 A;/C'/,o/7 de/l{?ch ol7) deq

LeH c5 fJcx R'9hf She/<.

(c) J = 60°,

flrjure /0. -Coflc/ud C?d .·

NACA TN No. 1208

AIleron -RIC/hi --Lc:f'f

------~ --~

, I

NACA TN No. 1208 F ig . 11

..

..

(r-I---~

~.~ , ~ ~ ~

~ ~.~

~ ............

~ ~

()

- U r--r---I--:-:--6

./6

-~ ~ I---- ---...... ~

10 .!t) 3'J 40 50 60 SAe hi angle ) 1; deg

(C)Ca./calafe..d Troff) I.:Jh = C? - NATIONAL ADVISORY ~ e) f) (r COMMITTEE fOA AERONAUTICS

FIgure //. - £/'/ec.l or SAeW a.ngle on damping )n /01/ an d aileron roilIng eTTectIYe/)ess.

"---'o_~ __ ~~ ____ ~_~ __ ----------

Fig . 12 NACA TN No. 1208

-.5

(.'\..

o o

- r--- -=:::-

/0

LColculal-ed LExoerJmenfa/

~ --- .~~ --...

~ --- t:::--~ \.~

~ ~ ...........

.......

~.

20 dO S kew angle)

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

40 r ) deg

so 60

H9l1YeI2.- Var;arlon Or damplng- 1/7-ro/1 paramekr CZp w;/i, sfew angle. £ypeYlmenfa/ dC/fa ,cor eoc.h skew Cln91e IS based 017 projecled span(b co).r J. C alculof-ecl Cvrye waS" oil-cun ed be; use of fOrmu/a Trom rererence 3 ) C~ == ell CS =:ofoS t.