DARPA/AFRL/NASA Smart Wing Second Wind Tunnel Test Results

11 Pages·2001·0.35 MB·English

by  Scherer, L. B.

useSPIE for Header
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ABSTRACT
oT yfitnauq eht stifeneb fo trams slairetam dna serutcurts evitpada gniw ,ygolonhcet Northrop Grumman .proC)CGN( tliub
and tested two 16% scale wind tunnel models (a conventional and a ‘smart’ model) of a fighter/attack aircraft under the
ASAN/LRFA/APRAD tramS slairetaM dna serutcurtS tnempoleveD - tramS gniW esahP1 . Performance gains quantified
included increased pitching moment (C ), increased rolling moment (C ) and improved pressure distribution. ehTstifeneb
M l
deerneiwatbo rof ,sselegnih deruotnoc gniliart egde lortnoc secafrus htiw deddebme epahs yromem yolla )AMS( seriwdna
e sgtinswiinwwatps y bdetceffe e beuutq r AoMtS mechanisms, derapmoc ot lanoitnevnoc degnih lortnoc.secafrus
sihT repap stneserp na weivrevo fo eht stluser morf eht dnoces dniw lennut tset demrofrep ta eht ASANyelgnaL hcraeseR
s’retneC)CRaL( 16ft Transonic Dynamic Tunnel (TDT) in June 1998. lufsseccuS stluser deniatboew re: 1) 5 seerged fo
esiwnaps tsiwt dna21-8 % increase in rolling moment utilizing a elgnis AMS euqrot ,ebut )2 21 seerged ,fnooitcelfed and
10% increase in rolling moment due to hingel ess, contoured aileron, and 3) demonstration of po tical techniques for
measuring t seiswitwnaps dna detcelf.eedpahs
1. INTRODUCTION
The purpose of the Northrop Grumman Smart Wing Phase 1 program was to explore the use of integrated smart materials
into an aircraft structure to provide unique capabilities resulting in significant aerodynamic performance improvements.
Further, the program was to demonstrate these capabilities during wind tunnel tests at NASA LaRC 16ft TDT wind tunnel.
wm eal irlnnvgaAaeorcrre pv e odbencun oenf.ri1efeR
POINT OF
FLOW SEPARATION
POINT OF
FLOW SEPARATION
DDNEAG NR IESHVCOITSIRETCA R EAW RHO1UCLGFIF
SECA FLROURS TDNEORCUO TSNSOECLEGNIH
owT cimanydorea stpecnoc erew desucof ,no hcihw yleuqinu dezilitu trams .slairetam ehT tsrif saw eht htoomsderuotnoc
trailing edge control surface. The drawback of the hinged control surface is the sharp discontinuity at the hingeline. This
discontinuity can cause the flow to separate prematurely, as shown in Figure 1, causing a degradation in performance which
semoceb erom decnuonorp ta rehgih lortnoc ecafrus .snoitcelfed ehT ria wolf htiw a htooms deruotnoc lortnoc ,ecafrusni
contrast, stays attached longer, significantly improving the overall pressure distribution. SMA wires were incorporated into
a elbixelf erutcurts no eht tramS ledom palf dna norelia ot edivorp eht htooms deruotnoc lortnoc ecafrus nehw .detautca A
etelpmoc noitpircsed fo eht AMS gniliart egde lortnoc secafrus si debircsed ninitraM 2 . ehT dnocescimanydorea
improvement concept utilized an SMA torque tube, Jardine 3 , ot edivorp gniw tsiwt draobni ot .draobtuo gnisaercnI ehtlacol
angle of attack on the wing outboard area provides for increased lift and pitching moment. This adaptive feature would
have application for take-off and landing configurations.
32”
32”
37”
37”

SL E GT DLNRDO AIANMENWMAROS UI2GTINFE VNOC
owT dniw lennut ,sledom nwohs ni erugiF ,2 erew detset ni eht ASAN yelgnaL TDT dniw lennut .ytilicaf enO ledom sawa
baseline conventional fighter-type wing, utilizing standard hinged trailing edge control surfaces with a trailing edge flap
and aileron. The other model was the smart wing, identical in shape and construction to the conventional wing, with the
noitpecxe fo AMS eriw detautca palf dna ,norelia dna a AMS euqrot .ebut owT stset erew detcudnoc rof mumixamksir
reduction:
Test 1 - First demonstration of smart wing 4 - gnitadilav cimanydorea stifeneb rof a yranimilerp foorp fo,tpecnoc
demrof rneip yaM.6991
Test 2 - Second iteration of integrated smart wing concept demonstration - further quantifying aerodynamic
stifene b g dnynnigoahoi slt/ioaslnrfbhgfa ceot,etysst naeedipuesmsriofrep ni enuJ.8991
noitarugifnoC Deflection or Lift Roll stnemevorpmI %
t sginwiTW DD CL DD Cl tfiL lloR
).geD(
Flap Only (Test 1) 5.7 1850.0 3910.0 %7.9 %2.01
Flap and Aileron Combined 5.7 6190.0 7830.0 %6.71 %1.71
(Test 1)
nor eyl)li1nAtOseT( 5 510.0 %0.8
nor eylli nAtOseT()2 01 9810.0 %5.01
Wing Twist (Test 2) 3 4430.0 3910.0 %0.8 %0.01
5 2050.0 6920.0 %5.11 %6.51
Wing Twist (Test 1) 4.1 6040.0 8120.0 %0.01 %8.21
Combined Aileron & 01+ o Aileron,
Wing Twist (Test 2) 5.4+ o Wing Twist 7650.0 130.0 %3.51 %3.71
STL U TESLSR E EEU YR3N TGDRNFINAUO FIMTWMUS
erugiF 3 sedivorp a yrammus fo elbaton cimanydorea stnemevorpmi rof eht elgnis stceffe fo ,palf ,norelia dna gniwtsiwt
dna eht denibmoc stceffe fo palf htiw ,norelia dna gniw tsiwt htiw norelia rof htob .stset hguohtlA eht sucof fo siht repapsi
the results of the second test entry, it is important to delineate the accomplishments that the integration of smart materials
nac edivorp ot llarevo tfarcria cimanydorea .ecnamrofrep ehT noitcelfed ro tsiwt si detsil ni seerged rof hcaenoitarugifnoc
including delta lift and roll aerodynamic performance improvements and provided in both absolute and percentage terms.
sA a lareneg ,tnemmoc ti nac eb nees taht egatnecrep stnemevorpmi fo eht trams ,ygolonhcet dehcaorppa ro dedeecxe 01%
rof eht elgnis ,stceffe dna degnar neewteb 51 dna 71 % rof denibmoc .stceffe tI dluohs eb detniop tuo taht segatnecrepera
evitale rot a e”fraesv“uen afmo 8 seerg e,dAOA g2 nrut dna a c iemraunsysder pfo 09 sbl rep erauq stoof 002~(.)hpm
A 10% improvement in lift and roll performances, in more tangible practical terms, would translate to:
• a 0002 bl esaercni ni ekat ffo ssorg thgiew )WGOT( rof a lacipyt 000,04 bl rethgif kcattatfarcria
• na esaercni ni ydaets gnillor etar morf 021 ces/seerged ot 231 ces/seerged rof eht tnelaviuqelanoitnevnoc
dna deruotnoc norelianoitcelfed
ehT tseT 2 palf ,stluser eud ot eruliaf dessucsid ni ecnerefeR ,2 era na suoivbo elbat noissimo morf erugiF .3 eroferehTeht
absence of any real flap data in Test 2 precluded a direct comparison between wind tunnel entries for this particular
.noitcnuf ,revewoH ni lla rehto ,stcepser tseT 2 evag gnigaruocne stluser dna devom eht ygolonhcet rehtruf .drawrof tseT
1 detneserper a tnacifingis noitartsnomed fo eht ygolonhcet dna ;stpecnoc ,sselehtreven emos snoitcirtser fo eht tset tespu
dna ecnamrofrep esora eud ot snoitatimil fo .epocs seussI dna srotcaf gnitcirtser eht tseT 1 noitucexe:dedulcni
• AMS eriw eugitaf smelborp ta noitanimret fo lortnoc ecafrus gniliartegde
• :yt inplo aiidlbtefactpeial mcf±ieLd ,5s.e7erge d noreli a± 5seerged
• Non- uniform deflection of flap and aileron, particularly in the spanwise direction
• mum igxn aitMwsiwt fo ,5s 2et.er1roghes dfo eht 5 eergleadog
All of the foregoing limitations were corrected in Test 2, which went according to plan except for a failure that limited the
d er toe cfhAp etMaalSltlfaodc 2 . Figure 4 highlights the achievements, results and benefits attributed to the program Test 2
.troffe spahreP eht tseb tnemevorpmi emac ni eht detadpu elgnis euqrot ebutngised 3 that permitted the target 5 degree twist
elgna ot eb .dezilaer ,revewoH ssorca eht ,draob lla fo eht trams lortnoc ecafrus stpecnoc detifeneb morfelbaredisnoc
design improvements over the configurations used in the first test. Model design enhancements unilaterally equated to
higher deflection angles and improved fatigue characteristics and more uniform control surface deflections.
tnemevorpmI Results Benefits
lor tenrooCM mr oefrionMU rtefhdignLiaH
no saerA lortnoC gnihctiP
tnemoM
& spalF AMS ecafruS
stnemevorpmI
snoreliA noitcelfeD
nlgeidsoeM&D rehgiH rtefhdignLiaH
Integration lortnoC weN noitcelfeD gnihctiP
Performance Surfaces se&lgnA tnemoM
ngiseD reetutqeiBtaF stnemevorpmI
stnemevorpmI
scitsiretcarahC
devorpmI
elgniS gnilloR
t sriewhTgiH
euqroTebuT tnemoM
selgnA
metsyS scitsiretarahC
ER U.G4IF T GRTWNASEIMEIW STV2REVO
le nsncuiTm acniynDos ny ta e6or1l1oT. gF2 AnSaALN
ehT ASAN CRaLTDT 5 saw desu ot tset dna yfitnauq eht cimanydorea ecnamrofrep stifeneb fo htob dniw lennut.sledom
The TDT is a unique facility in that it is designed primarily for aeroelastic research, and validating vehicle performance for
safe operation with respect to aeroelastic instability. It is important to understand that the models used for the smart wing
margorp erew ton dengised ot yduts citsaleorea ,stceffe revewoh ecnis rieht noitcurtsnoc saw ralimis ot na lautcatfarcria
s, e tgcrn cn i ereeismnhelfwa ao tcbf wfns nieeeoylroshldeec sti dd.ohronspwimudaao l esnh ’oTtTis DtesTcties 61 .tfyb
61 .tf dna ylisae detadommocca eht tset .sledom ehT ytilicaf nac edivorp elbairav lennut latot erusserp morf raen muucavot
eno erehpsomta morf hcaM srebmun fo 1.0 ot evoba eht deeps fo dnuos ta hcaM .2.1 ehT ytilicaf osla sah ehteuqinu
capability of using either air or R-134A high density gas, which greatly assists in the scaling of aeroelastic models. The
high density gas, however, was not required for testing on the Phase 1 program.
2.2 noitalla tlsen nIsdnlnu eaTldeo nMdnnuiTW
owT naps-imes gniw sledom erew evitatneserper fo a lacipyt kcatta-rethgif .gniw htoB gniw sledom erew lacitnedi ni,ezis
contour, and construction 2 similar to an actual aircraft wing with spar and rib interior construction and aluminum skins.
The conventional wing model had an electrically driven, remotely actuated trailing edge flap and a manually moveable
hinged aileron. The smart wing had a smooth contoured trailing edge flap 2 fo eht emas ezis sa eht lanoitnevnoc .gniw tI
saw detautca yb yllacirtcele derewop AMS seriw detacol ni eht reppu dna rewol sniks fo eht .palf ehT trams gniwnorelia
also provided smooth contour and was actuated in the same manner. The smart wing also incorporated a single SMA
euqrot ebutmetsys 3 esolc ot eht raps-dim hcihw saw detautca yb eht gnitaeh fo emorhciN eriw depparw dnuora ehtAMS
.eeubqurtot
Figure 5 is representative of the test installation and shows a plan view of the smart wing demonstration article as installed
ni eht tset rebmahc fo eht yelgnaL .TDT hcaE ledom saw detnuom yletamixorppa eerht teef ffo eht llaw no a rettilps,etalp
nwohs ni erugiF ,5 hcihw sevom eht ledom tuo fo eht llaw yradnuob reyal dna drawot eht retnec fo eht lennut erehwria
wolf si erom .mrofinu hcaE ledom saw ylmrif detroppus no eht ASAN yelgnaL napS-imeS elbatnruT metsyShcihw
extended through the tunnel wall. The turntable provided capability for the model to be pitched through an angle of attack
egnar fo 03± seerged fi.deriuqer
ledoM & ecroF tnenopmoC-5
eldnipS ec ntanleamBoM
reettatliPlpS
)S50TDT(
ledo Mo tecnalaB
retpadA
Splitter Plate
l llaeWnnuT
maertsn wgonDik owoeLiV citameh cwSe ipVoT
ER U5 GNI OFL YNIE EI TN6DLT AN1 NGDAL UINTSLTWAAA LNTL SENDIOM
Critical set up items also depicted in Figure 5 were the spindle and balance. The spindle was necessarily rigid to provide
etauqeda troppus rof eht ,ledom hcihw si derevelitnac ffo eht .elbatnrut ehT eldnips dna ecnalab erew detcetorp morfeht
wo l gf .ynrd”bi ieer ra soinosona a.llafnicc oan“ f itedoe te e cec Dldgn heeeinantsrdviliba oowauemrobshpltlof
gnirot indnonoiaMtatnem u3r. t2 tssneIT
Each model was extensively instrumented during testing to record test conditions and results. The model instrumentation
:seiroge teae cro htdtneidi vyildd a yodearebmbyolpme
• Aerodynamic force and moment measurement and pressure instrumentation
• noi td rcne ueacgolanttffinserwoiducwSt
• Safety and routine test condition monitoring
ehT cimanydorea gnirotinom saw demrofrep gnisu ASAN s'yelgnaL yb ,S50-TDT a evif tnenopmoc ecrof dnatnemom
balance capable of recording the lift and drag forces, and pitching, rolling, and yawing moments. Each of the models had
ruof esiwdrohc swor fo citats erusserp spat ot erusaem lacol ydaets etats citats erusserp sa nwohs ni erugiF .6 ehTerusserp
sebut erew detuor hguorht eht rettilps etalp dna otni eht "eonac" gniriaf erehw yeht detcennoc ot a trop no eno foevif
yllacinortcelE dennacS erusserP )PSE( .seludom esehT seludom erew detarepo dna detarbilac enil-no yb eht s’TDTISP
. ml eo0tr0st4yn8SoC
STRAIN GAGES- ROSETTE y = 7.060
STRAIN GAGE- FULL BRIDGE y = 9.060 OnlMyodSemla rt - BOTHC ONVENTIONALA NDS MARTW ING-U PPERA ND
WGENIIVKNOWOOLD TSICLHTES AEVNISTOZRI SNCLINOMETER y = 14.483 CMoonSdvmeealnr-att ni do nal LSOMWAERR TW SIUNRFGOA CNELY-U PPERS URFACEO NLY
INCWR POTENTIOMETER y = 19.901
IELS
TLEI THERMOCOUPLE y = 24.000 y = 28.000
ACCELEROMETER y = 28.310
y = 30.004
y = 31.500
IMS y = 33.690
TMI
y = 37.375
TTTC OELS TLEO
TTTD SITT
OMS OMT
IFAMSP TILTWT AWTFS
PTEF SAI PSMAFM TILTMR OFAMSP SAO SMTWA
IAAMSP PSMAAO AWTAS
IAT
OAT
FAMST
AFLAP AAILTSMAA Row E Row A Row B Row F Row G Row C Row H Row D
SNOITA CT ORELORPUSS ECRIPT ADNTNOSAITATNEMUR TE SRL6NUAIG NIRFE TNI
Smart control surface contour tip deflection measurements were obtained using both internal and external methods. The
external methods, however, proved to be the most reliable. Figure 7 shows the primary approaches used in determining the
.snoi tpcietlfed
lanretnI )yramir Pl(anretxE
• -sretemoitnetoP • oediV yelgnaL ASAN
lortnoC dna noitisoP ecafruS noitamrofeD ledoM
8,7,6
)DMV( metsyS
• gn i-Wsrosn eSSMEM
Twist
• gn i-Wsrosn etSliT • yelgnaL ASAN
erioM noitcejorP
)IMP( yrtemorefretnI
metsyS
RD UNYO ARTFAN OMON MCOEUIRST UA7GTINF EMUR TSNI
TNEMER UTGSSNAIIEWWMT
An internally mounted rotary potentiometer was used to measure flap position for the conventional wing. The aileron
noitcelfed ,elgna ,hguoht saw tes yllaunam yb .stekcarb esehT stekcarb erew deraperp erofeb gnitset nageb dnaerew
fabricated in ±5 degrees, ±10 degrees, and ±15 degrees settings. The smart wing used linear potentiometers at each of the
three control areas on the SMA flap and the two areas on the SMA aileron. The potentiometers attached to the tips of the
AMS palf dna norelia yb a elgnis eriw dna erew detacol tsuj drawrof fo eht tfa .raps ehT tnemhcatta eriw saw eerf otevom
within the flap through a channel in the RTV skin. Tip deflection measurement using linear potentiometers, however,
devorp ot eb yrotcafsitas tub ton lamitpo ni smret fo noitcelfed ycarucca dna .ytilibataeper ehT niam nrecnoc sawtaht
hngouiothctell afs ea,dwdedivorp yeht did ton edivorp tanemerus afeom eht tnuoma f.oerutavruc
A Q-Flex inclinometer was mounted on the spindle pitch mechanism and a Scheavitz Inclinometer was mounted at the root
fo hcae ledom ot erusaem eht elgna fo .kcatta rehtO lanretni secived erew detagitsevni rof rieht ytiliba ot erusaemgniw
noitator eud ot .gnitsiwt gnomA eht rehto secived erew SMEM srosnes depoleved ta yelgnaL dna tlit srosnesderutcafunam
by Advanced Orientation Systems Inc., both of which were attractive for their small size, which aided installation into the
draobtuo gniw .snoitces ehT SMEM ,srosnes hcihw demrofrep ot nihtiw eht deriuqer 52.0± seerged citats,ycarucca
detaroireted ot sa hgih sa 0.5± seerged rednu cimanyd .snoitidnoc ehT citylortcele tlit srosnes devaheb llew rednucimanyd
snoitidnoc tub osla dereffus morf ytilibailer smelborp htiw tcepser ot seicnapercsid ni eht derusaem epols neewtebevitisop
dna evitagen .selgna ehT yek esoprup fo lla fo eht lanretni srosnes saw ot edivorp etarucca noitisop atad rof ehtderuotnoc
surfaces and spanwise twist that could be used in wind tunnel measurements or in flight. (Developing and using sensors for
laer emit kcabdeef ot eb desu rof eht thgil flortno cmetsys si na noisnetxe fo eht suoiverp aedi taht dluohs eb rehtrudfeusrup
on a follow on program, and all of the methods in the left hand column of Figure 7 , though lacking in maturity, should be
degaruocn erof )r.ethntermupfoleved
evitscteelgfreaRT
ERUG I.F8 YA ESOLAEGLNDNEIADVLOM NOI T)ADMMRMVOE(FTESDYS
External methods, on the other hand, were much more successful on the program for measuring control surface position and
met s)yD sMnVo(itamro flee Ddooe MdliaVci type ol.g tnA saSeiLAhwNTt 6,7 saw eht yramirp metsys rof gnirusaemAMS
palf dna nor enloiiatisop dna gniw .tsiwt ehT doht esmyolpme a seires f oevit clell afemepsra-tor tsec rsyildluf edreaccalpta
several spanwise rows as shown in Figure 8. The discs serve as targets, which when illuminated with a light generate a
high-contrast image in a strategically located CCD camera. Image processing is used to automatically locate the targets.
Photogrammetry is then used (after calibration) to determine the motion of the targets in the pitch plane and hence compute
angular changes at the various spanwise target locations. The system is able to provide angular deflection in "near real-
time" with updates in angle approximately a few seconds after triggering.
noitcejorP érioM yrtemorefretnI )IMP( saw a dnoces lanretxe lacitpo citsongaid euqinhcet desu ot erusaem gniwecafrus
epahs dna .noitamrofed ehT IMP metsys desu na ,derarfni deslup resal ot tcejorp a seires fo ,decapsiuqe lellarap senilotno
eht rewol ecafrus fo eht .ledom segamI fo eht detcejorp dirg senil erew derutpac gnisu a DCC aremac dna emarf rebbargni
ecnerefer )ffo-dniw( dna no-dniw .snoitidnoc egamI gnissecorp seuqinhcet erew neht desu ot tcurtsnocer eht ledomecafrus
shape and/or deformation under aerodynamic load. Figure 8 shows PMI-measured quantitative wing surface shape
measurements for the smart wing in non-actuated and torque-tube-actuated conditions. The PMI measurements were
spatially continuous at a resolution of 0.055 in. per pixel (1.4mm per pixel), far exceeding that of the VMD system.
,revewoH ta sti tnerruc etats fo tnempoleved rof lennut-dniw ,gnitset IMP seod ton ssessop eht hgih eerged foycarucca
demonstrated by VMD, and quantitative real etmuiptt- usoi.e ltboinssop
006 peahS
,emmliforP
005 .003
004 .051
m
Chord, m 000023 0.0
.051-
001
.003-
0
001 90 80 70 60 50 40 30 20 100 90 80 70 60 50 40 30 20
anpS tnecreP anpS tnecreP
edtsiwtnU gniW edtsiwT gniW
MY)ERIT TMSEEPYRM(SURGE IF9NEFROREIITTON C MIYE EJALOSGRANPNAL
The third category of instrumentation used during the wind tunnel testing relates to instruments used to monitor model and
ytilicaf .ytefas gnomA eht srosnes desu ot rotinom eht ledom erew niarts ,segag sretemorelecca dna .selpuocomreht ehT
noitacol fo eseht srosnes si nwohs ni erugiF .6 ehT eerht sretemorelecca ta eht gniw pit erew dellatsni ot rotinomledom
dynamics. The strain gauge rosettes, mounted inboard and outboard on each of the three wing spars, provided real time
model dynamics and structural load monitoring. Thermocouples were installed on the SMA torque tube, naturally, to
monitor its temperature. They were also co-located with the strain gages to provide temperature compensation for the strain,
if necessary, and determine the effect on the internal temperature through out the model due to torque tube actuation.
snoitid ndeoncCaneu qt .esm 2Sea4Trg orP
A rettulf ecnaraelc nur saw edam ta eht trats fo eht tset rof hcae fo eht owt sledom sa a ytefas noituacerp ot erusneledom
cimanyd ytilibats roirp ot gnidrocer yna lautca gniw ledom ecnamrofrep .atad ehT ledom saw nekat ylwols otgnisaercni
dynamic pressures and Mach numbers, for flutter clearance, and angle of attack varied between 0 and 10 degrees. The
nosaer rof gniod siht saw ot raelc eht tset epolevne dna yfirev taht eht ledom devaheb yllacimanyd sa detcepxe ni ehttset
range.
ehT sledom erew detset ta owt tnereffid tnatsnoc lennut latot serusserp fo eno erehpsomta aes( level ro 0022 )fsp dna5.0
atmosphere (17,000 ft altitude or 1100 psf). Tunnel freestream dynamic pressures were set at 60, 90, or 120 psf. This
dluow dnopserroc ot hcaM srebmun fo ,2.0 ,52.0 dna 92.0 ylevitcepser ta eno erehpsomta dna ,3.0 ,63.0 dna 4.0ylevitcepser
at 1100 psf tunnel total pressure. Data was recorded at specified, fixed angles of attack (AOA) starting at -4 degrees and
increasing up to +16 degrees in fixed increments of 2 degrees except near the 0 degree AOA point, where the increment
saw 1 .eerged nuR ,sralop stolp fo cimanydorea stneiciffeocC( , C , ).cte susrev ,AOA erew detareneg ta suoiravlatot
L M
,seruss ehrcpaM srebmun dna dexif gniw tsiwt / l o.rstn noeoicctaafrruugsi fnoc ehT snaem ftonemerus aseamw sa:swollof
• ET palf noitcelfed elgna no lanoitnevnoc ledom saw tes htiw na cirtcele rotom dna detarbilac htiw ayrator
retemoitnetop
• neol igstn ac answetot l erehfdsekteelcidxiawiArfb
• tram Sgniw AgMn Sinl oieiagtrdc teedlnfae dgni wtsiwt d e. yb)dlieelrvracoucsbsieaatdep(mo
3. TEST RESULTS
A pot level noissucsid fo eht dniw lennut tset stluser saw detneserp ni eht .noitcudortni sihT noitces sweiver ni eromliated
each individual function tested. The wind tunnel results are discussed by adaptive concept, and in general, results for each
configuration on the conventional wing are compared directly to the equivalent configuration on the smart wing. This
gnipuorg fo scipot spleh yfiralc eht atad noitatneserp sa derapmoc ot a lacigolonorhc .mrof ,ecneH snoitcesbusera
structured into:
• ytilib aetnaielpeesraB
• Wing twist
• Aileron performance
• srotcef fdeenibmoC
3.1 Baseline Repeatability
ataD saw tsrif derapmoc rof eht enilesab ,snoitarugifnoc ,.e.i lla lortnoc secafrus ta a lartuen ro detcelfednu etatsneewteb
the conventional wing and smart wing. This was to ensure that the two models were identical in shape and to verify proper
.noitallatsni enilesaB ytilibataeper atad nwohs ni erugiF 01 hcaM( = 52.0 dna cimanyd erusserp fo 09 )fsp swohs stolpfo
,tfil ,hctip dna gnillor tnemom tneiciffeoc susrev elgna fo .kcatta ehT atad stolp swohs yrev doog snosirapmoc neewtebeht
conventional and smart wing, giving confidence that the incremental data obtained from subsequent tests to establish
esctnnaemmreovforr eprpmoif eht trams gniw dluow eb.etarucca
0.8
CL
0.15
)33 nuR( gniw vnoC
0.6 0.01 )3 1n1u Rg(n itwramS 0.1
]ge dA[OA 0.05
0.4 ) 3n3u Rg(n ivwnoC -5 -0.01 00 5 01 51 02 -5 00 5 01 51 ]ge dA[OA 02
)3 1n1u Rg(n itwramS -0.05
0.2 -0.02 1.0- )3 1n1u Rg(n i)t33 wnruR( agniwm vnoCS
A[OdAe g] -0.03 -0.15
-5 00 5 01 51 02 -0.04 2.0-
-0.05 -0.25
-0.2 -0.06 3.0-
Croll
-0.07 hctipC -0.35
-0.4
Lift Coefficient vs Angle of Attack tsnveiciff etonCe mgonMihctiP kcattA fo elgnA sv tneiciffeoC tnemoM gnilloR
ekl cgfanotAtA
ERUGIF 01 EYNTIILLEISBAABTAEPE R- TSET 2 Q( = 09 ,FSP HCAM =)52.0
3.2 Wing Twist
enO fo eht yrami rspevitcejbo fo eht dnoces tset yrtne saw ot esaercn iesiwnaps gniw tsiwt morf 52.1+ seerged ni tseT ,1ot
5+ .seerged sihT evitcejbo saw dehsilpmocca dna stolp fo eht tfil dna gnillor susrev elgna fo kcatta rof eht tramsgniw
untwisted and twisted at 3 o and 5 o are shown in Figure 11. These incremental improvements, however, were relatively not
as large as obtained from the first entry. A 1.25 o twist provided an approximate 12.8% rolling moment improvement at 8
seerged fo elgna fo kcatta elihw rof tset yrtne ,2 a 3 eerged gniw tsiwt dedivorp a %01 gnillor tnemom esaercni ta ehtemas
angle. Closer examination revealed that for the first entry, the inboard section of the wing was actually twisted more than
gnirud eht dnoces dniw lennut .yrtne ,ylfeirB rof eht tsrif yrtne a mednat metsys fo AMS euqrotsebut 3 saw ,desu enomorf
the inboard rib to the mid rib and another from the mid rib to the outboard rib. In test entry 2, only one torque tube with its
attachments connecting the inboard rib to the outboard rib was used. This approach provided significantly more twist at the
outboard wing tip than the first entry but not on inboard side of the wing where there is greater amount of wing area.
Inboard rib to mid rib accounts for 73% of the wing area for this particular configuration. Figure 12 shows the amount of
tsiwt gnola eht htgnel fo eht gniw gnitcelfer eht draobni tsiwt dna eht secnereffid neewteb tseT 1 dna tseT .2 erutuFkrow
might concentrate resources on the inboard section of the wing installation area.
0.8000 0001.0
0.6000 0050.0
0.4000 0000.0 AOA
000.6- 000.4- 000.2- 000.0 000.2 000.4 000.6 000.8 000.01
0.2000
0050.0- Run 126 No Twist
0.0000 AOA Run 129 Twist = 3 deg
-15.000 -10.000 -5.000 000.0 000.5 10.000 15.000
0001.0- Run 133 Twist = 5 deg
-0.2000
Run 126 No Twist
Run 129 Twist = 3 deg 0051.0-
-0.4000
Run 133 Twist = 5 deg
-0.6000 0002.0-
LC BLLORC
-0.8000 0052.0-
Lift Coefficient (C ) sAvOA gnilloR tnemo MtneiciffeoCC( ) .svAOA
L M
ST CT -EST FE SE2GFT.IRNE1WUI1TGWIF
PALF( = 0 ,SEERGED Q = 09 ,FSP HCAM =)52.0
00.4
00.3
aerA gniW latoT fo %37
00.2
Twist (deg) 00.1
)ged( 1 tseT tsiwT
metsyS IMP
tiF evruC 1 tseT tsiwT
00.0
0 5 01 51 02 52 03 53 04
).n in(apS
00.1-
E RN UO.G I2IET1FUU DB E ITS ORSITTIW SWNEAITAUMDPEQSSBRUOTT
3.3 enconraemlrioAfreP
ehT ecnamrofrep stifeneb dedroffa yb eht AMS detautca palf dna ,norelia susrev a degnih ,noitarugifnoc si ,spahrep enofo
the most critical results and achievements of the smart wing phase 1 program. Although the SMA flap failed 1 * during the
dnoces ,yrtne eht norelia detartsnomed eht ecnamrofrep sniag fo gnisu a htooms deruotnoc lortnoc .ecafrus erugiF 31swohs
the improved pressure distribution on the contoured aileron (dark area) versus a hinged aileron (light area) with reduced
flow separation at the trailing edge. A superior design addressed the first test limitations which limited maximum
noitcelfed dna mrofinu noitcelfed gnola eht .naps ehT wen ngised desaercni noitcelfed morf 5 ot 01 seerged dnadedivorp
rof a erom mrofinu .noitcelfed ehT stifeneb fo eht deruotnoc norelia revo eht degnih norelia ta 01 seerged fo noitcelfedera
shown in Figure 14. The 10.5% improvement, shown in Figure 14, is a 2.5% improvement above the first entry results at
the same test conditions.
PressIumrper,oved
Reduced Separation
T OELRPUSS E,RSPSENEVITCE FN F OEE3RR1EU LGIIAF
N U1R4( sv . 116, Q = 90, AOA = 6)
0.1
Run 41 Conventional Wing
Run 119 Smart Wing
50.0
0 AOA [deg]
6- 4- 2- 0 2 4 6 8 10
-0.05
Conv. and SMA
-0.1 see r pg =ae0ldF
-0.15 Conv. and SMA
seerged 01 = noreliA
-0.2
-0.25
10.5%
-0.3 Improvement
Croll
-0.35

gnilloR tnemoM tneiciffeoCsv Angle of Attack
NOI-TCE LSSFESE EED- RN0GE1EVDITC E NEF.ORF4RUE1 EGLIIFA
TEST 2 (Q = 90 PSF, MACH = 0.25)

1 * Detailed explanation is given in Reference 2. From Test 1, flap performance showed substantial benefits 1 etipsed rewolnoitcelfed
selgna d nmnarooiftiunbui-rntosnid
An alternative method of visualizing the aileron effectiveness is illustrated in Figure 15. Here the rolling moment
tneiciffeoc eud ot norelia noitcelfed si dettolp rof tnereffid elgna fo ,kcatta yleman ta a = ,6 ,8 dna 01 .seerged owT
important features are shown on this graphic. First, at zero degrees deflection, the rolling moments are very close which
dluow yltceridni yfirev eht ycnetsisnoc neewteb eht .sledom ehT dnoces si eht esaercni ni ecnamrofrep tnemevorpmihtiw
increasing control surface deflection angle. This shows that the separation reduction the contoured aileron provides results
in an improved lifting surface and a lower deflection angle to obtain the same lift as a conventional aileron.
0
51- 01- 5- 0 5 01 51
noreliAnoitcelfeD
50.0- 6 =AO A,71 1nuR
8 =AO A,71 1nuR
0 1 =AO A,71 1nuR
A, Og= An.6ivWnoC
1.0- A, Og= An.8ivWnoC
Conv. Wing, AOA = 10
ta ecnamrofreP desaercnI
51.0- snoitc erlefhegDiH
2.0- a= seer g e6d
52.0- a seerged 8 =
Cl
3.0- a seerg e 0d=1
53.0-
NOITAI RASAH SVP-ELNAEVITCE FNFOERE L.EI5RA1UGIF
(Q = 90 PSF, MACH = 0.25)
ecnamro fdr ee4nP.i3 bmoC
tI si tnatropmi ot etaicerppa taht ni lla fo eht evitpada stceffe dessucsid os ,raf ti si ylekilnu taht yna eno fo meht lliweb
desu ylevisulcxe ni noitalosi rof a etelpmoc .noissim erehT lliw eb ynam secnatsni erehw stceffe lliw eb denibmoc hcussa
during take-off and landing where the flap and aileron will be actuated simultaneously, or a combination of wing twist with
aileron might be used. Test 2 examined the combined effects of aileron and wing twist on rolling moment. Figure 16
demonstrates a substantial improvement of 17% (at 8 degrees angle of attack) when both twist and contoured aileron are
used in unison compared to a conventional aileron alone. This further demonstrates the benefits of using smart control
.secafrus
2.0
1.0
0 ]ged[ AOA
01- -5 0 5 01 51
ged 01 = noreliA .vnoC ,14 nuR
1.0- ged 01 = noreliA tramS ,431 nuR
17.3%
2.0-
tnemevorpmI
3.0-
BllorC
4.0-
E RE U SGGTD.AN INE6E IRFEN1RGLOMICNLFTOBNIOSMMIWRIOWCT
NOITCEL FNEODRE LTIR AADMNSA