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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
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8 /6 ' 24 32 Ary/e of attock) CX, dt:9
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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 <:)
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-fl.~ .s=-~~ .>=-= ~ .):" - -l'r -, .:.-- - "t.
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--<: -- 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.