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NASA Contractor Report 159389
(HASA-CR-159389) RESDLTS OF t~SIGH ~TUDIES N84-33403 AND WIND TDN~EL rESTS OF AA ADVANCED HIGH LI:T SYSTEH Foa Ah E~LE~Y EFFICIENT TRANSPORT (Douglas Ai~craLt Co., I~c.) Unclas 373 p HC A16/ftF 101 CSCL 01C G3/05 24033
Results of Design Studies and Wind Tunnel Tests
of an Advanced High Lift System for an Energy
Efficient Transport
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Wayne R. Oliver
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McDonnell Douglas Corporation Douglas Aircraft Company Long Beach. California 90846
CONTRACT NASl-14744
DECEMBER 1980
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NASA Contractor Report 159389
Results of Design Studies and Wind Tunnel Tests
of an Advanced High Lift System for an Energy
Efficient Transport
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Wayne R. Oliver
Douglas Aircraft Company
McDonnell Douglas Corporation
Long Beach, California 90846
Prcpdrt'd for I.angl .. y H .. search C .. nt .. r und .. r Contract :'Io:\S\-I4,I-1
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ro~~DRD This document presents the results of a contract study perfo::med for the National Aeronautics and Space Mninistration (NASA) by tre Douglas ~..iLcraft Comr~ny, McDonnell Douglas Corporation. This work was part of Phase I of the Energy Efficient Transport (EET) project of the Aircraft Energy Efficiency (ACEE) program. Specifically, the study was one task in the contract Selected Advanced Aerodynamic and Active Control Concepts I:evelopnent. Tte activity included tre design and testing of a low-speed, hlgh lift wind tunnel maCel incorp::>rating a high aspect ratio supeLcritical wing and advanced high lift system. The model included variable c<lmber Krueger and slat leading edge devices, and tWO-SEgment and single-slot flap configurations. Optimization of these components, as well as effects of nucelles and pylons, landing gear, aileron, spoilers, and horizontal tail were studied experiIrentally.
~ Ackno\vledgements for their support and gUldance are given to the NASA technical monitor for tie contract, Mr. D.L. l-1aiden of the Energy Efficient Transport Project Office at the Langley Research Center and Mr. J.~.
Tulinius, tre on-site NASA representative; also to Dr. R.T. Nhitcomr Qf the Langley Research Center for his concept of tre supercritical wing.
Acknowledg2rnents are also given to tre director and staff of the NASA Ames and Langley Research Centers, at which facilities tre test prograrr.s were cooducted. Tre cooperation and discussions concerning high lift development for high aspect ratio supercri tical wings of rlr. R.J. r-iargason and H.L.
liorgan of tre NASA Langley STAn Low-Speed Aerodynamics Branch was greatly
awreciated.
~ Douglas personrel who made significant contributions to this work were: H. Klotzscre FLEE Program l1anager A. B. Taylor EET Project ~lanager J .G. Callaghan Branch Chief, Configuration I:esign and
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I:evelcprent, ~rodynamics Subdivision
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Aero~ics ProJect Engineer - EET J.T. Callaghan Secti~n Chief, Stabil~ty, Control, 3nd D.E. Ellison Flying Qualities, Aerodynamics Subdivision Task t1:lnager, Aerodynamic Des~gn (Re{X)rt Author) \'l.R. Ollver B.K. Bergren Aerodynamic Design
J .vI. Humphreys Aerodynamic Design
D.G. t-Bd'lilkinson Unit Chief, Advanced ~rcial Progrt.lllls, AerodjlXUnic Design Aerodynamic Design P.}!. Minor Aerodynanic Design G.G. Myers Ae rodyn:unlc Design L.A. Sp:icht Aerody~lic Design N. F. \'la~son Branch ~hlef, Aerodyraml.CS l'lind Tur'.rel J.D. Cadwell (I,lCc 1 Gn IjP /I.£::wdynaml ~ 'hnd ':'unrel :1odel GlOUp :'.13. bcw'::J11
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Aerodynamics i-lind Tunrel ~Iodel Group D.H. BLoceroon p.ecodY.:BITUCS \'h~d Tunrel Model Group
L.o. Scl-)e.r~r
Aero.j'jnam~c ~sign (Secretary) A.P. Morgan
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TABLE OF CCNrENI'S Page S'lM-1ARY SYMBOLS IN'l'IDDUCl'IOO PRELIMINARY TRADE SI'UDIES Background • • • • • HJ.gh Speed Wing Pararret.?..rs • Spanwise DistrIbution of Lift, or Span IDading 11 Spanwise Distribution of 'lbickness 11 A1.rfo~l Characteristics • • • • • • 12 Factors Influencing Selection of High Lift Systems • • 12 Trade Studies ••• • • • • • 17 High Lif t Cotl'{Xlrent Studies • Final Configuration Studies • HIGH LIFl' l£mDYNAHIC reSIGN ProCESS 41 General Design Constraints 41 High Lift Design Methods • 42
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Experimental Data Base 43 2-D Analytical Prediction for 3-D Geometry 45 2-D to 3-D Concepts • 46 3-D Lifting Surface Theory 46 SlaFe Definition for Leading and Trailing Edge Elements 18 VlIND TUNNEL mIEL reSCRIPTION ••••••••• Configuration Notation and Dimensioral Data Instrumentation • 76 Hodel InstailatlOn •• EXffiRll-ENTAL reSULTS AND ANALYSIS • 83 Test Plan and Facilltles ••• 83 Test Technique • • • • • • • Wind Tunrel Test Results and Analysis 85 Crmoo Wing. 86 Naceile/Pylon Effects on Cruise Wlng • 107 Reynolds Number and Hach Number Effects on Cruioo Wing • 118 Cruise Wing with Horizontal Tail • • • • • • • • • • • • 126 .....
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TABIE OF CDN'IENTS - contirued .Fage.
va< Configuration • • • • • • • • • • • • • • • • • • • 137 va< landing and 'lakeoff Optimization • • • • • • • 139 '!Wo-Se<Jllent landing and Takeoff Flap Optimization 177 va< Span and Nlcelle/Pylon Effects • • • • • '. • • • 177 Reynolds Number Effects on va< with '!Wo-Se~nt Flap (Tail-{)ff) • • • • • • • • • • • • • • • • • • • • 198 Va< \~ith '!Wo-Segment Flap:> and Horizontal Tail • • 198 Single-Slot Flap Optimization • • • • 198 Inboard Vcr. Effects •••••••••••••••••• 198 Reynolds Number Effects on Va< with Single-Slot Flap (Tail-Qff) • . • • . . • . • • • • • • • • • . • • • • • 221 Va< with Single-Slot Flap and Horizontal Tail • • • • • 221 Slat Configuration 221 Slat landing and Takeoff Optinuzation ••••••
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l-ach Nurr.ber Effects on Slat with '!Wo-Segnent Flap Reynolds Number Effects on Slat with '!Wo-Segment Flap. 255 Slat with '!Wo-Segment Flap and Horizontal Tail 265 Slat with Single-Slot Flap Characteristics • 265 Clean Leading Edge Configuration 281 Clean Trailing Edge Configuration Aileron and Spoiler Studies • • • • • • • • 297 landing Gear Studies • • • • • • • • • • • 306 Data Summary and Comparlsons of Leading and Trailing Edge High Lift Concepts • • • • • • • • • • • • 306 COl1PARIS)N OF EXPERIMENTAL DATA wrm ESTIMATION ME'lIDDS 335 COUCLUSIONS AND RECDr·M:NDATIONS 343 COnclusi01S •• Fec:x::mrendations • REFERENCES •
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IU1JS'rnATIONS Figure Page
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1 Corrparison of ro...-Speed OBracteristics for Current and Advanced Transports • • • • • • • 10 , 2 Supercritical and Conventional Airfoil Comparison .13
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3 Factors Influencing High Lift System Definition • • 14 t 4 Takeoff Defimtions • • • • • 15 Takeoff Flightpath Profile • 15 6 Clirrb Gradient Sunmary • • • • • 16 7 Aircraft Noise Certification Reference Locations • 16 8 FAR Part 36; ICl'O Anrex 16, Qapter 3 m~w Designs) Noise Level Requirements • • • • • • • • • • 18 High Lift System Configuration Comparison • • • 19 10 ConventlOnal and Advanced High Lift System Study Configuratlons • • • • • • • • • • • • • • • • • 20
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11 Effect of Advanced High Llft System on Fuel Burned • • 21 12 Advanced Hlgh Llft System Fuel-Burned Components • • 22 13 Conventional and Advanced High Lift System 'lakeoff Field Length Performance • • • • • • 0 • • • • 22 14 ConventlOnal Versus Advanced High Lift System Comparison • • • • • • • • • • • • • • • 23 15 Variable Carrber Krueger Leading Edge • 24 16 High Lif t System • • • • • • • • 26 17 DC-X-200 General Arrangement • 28 18 Factors \Veig~d in Determining Optimum Cruise Speed • • 29 19 Effect of Wing S\./eep on Direct Operating Cost • • • • • 29 20 Effect of Adjusted Aspect Ratio on Economic Parameters • • • 30 21 Effect of Adjusted Aspect Ratio on Block Fuel • • • • • • 30
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Page F~~= 22 \ling Area Study Results •••• Jl Aircraft Performance Sttrm:!.ry 23 32 CCInparisoo of 'lbi.ckness and Cruise Lift Coefficients - Ccxlventianal Transport and r:c~-200 • • • • • • • • • •• 33 Improvement in Cruise Lift-to-Draq Ratio for Advanced High Aspect Ratio Wi.rq •• • • • • • • • • • • • • • •• 34 InproVanents in I.GI-Speed Perfox:nanoe for Advanced Technology Configuration • • • • • • • • • • • • • 34 27 FAn Part 36; ICAO Annex 16, Cllapter 3 (New Designs) Noise !evel Requirements ani Est:.ir.lated OC-X-200 Values • 35 28 Estimatei 100-EPNdB Awroach and Takeoff Noise Contours (lOOO-Statute-Uile Trip) • • • • • • • 36 29 FAR Part 36 Flycver Noise I.evels •• • • . • • 37 1O Cotparatiw Olaracteristics 38
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31 Relative Fuel Burned • • • j~ 32 Relative Direct Operat.in:r Cost for Conventional an:!
Advanced COnfigurations • • • • • • • • • • • • • • • •• -.19
33 Win:J Gearetry <llOi.ce
J4 Advanced Fea·.:ures Econatu.c AsSeSSi'o..;!llt 4CJ 3S SUrfaces Defb16d by Cruise and High L:t.ft Considerations. 41 36 ~gh Lift Aercdynam:i.c Design Precess •• • • • • • • .• -l2 Fla-l Diagrar.1 ot Carputer Program for Multi -Elanent Airfoil Des~gn :md Analysis Methcrl (WJWIll) •••• .
Cot;>arisor. Between Ex:perimental and Analytical Pressure Distributions for Multi-Elarent Airfoil 4S 39 r-~ Pressure Dist.rihltions for Three Span Stations.. 46 'rheoret~cal Spanwise Variat~on of Slat and WUss MlnlID..!!l1 Pr.:ssures for the Low-Speed Wind 'funnel
r·bdel (Landulg Flaps S. ;: 3.4) • • .• 47
41 UJ--l86A H,xiel Three-View ••• • • • • • • • • •
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./ Page Fi~.a 42 High Lift Cat'q:alents Evaluated in Exper.inental Test Program •• • • • • • • • • • • • • • • • • • • •• 51 43 Wm;r Diagram (W3B) ••••• • • 44 Horizcntal stabilizer HlA Diagrcm • • • • • • • • • • •• 53 45 Vertical Stabilizer V lA Diagrcm • • • • • • • • • •• 54 46 Nacelle/pylcn N P 2A Diagram. • • • • • •• 55 2A 47 Aileron Section (a ) ••••• •• 56 2A 48 ~g. E:cl;e Device Gap, OVerhang, and Deflection Ilefl.nl.. tiOIlS • • • • • • • • • • • • • • • • • • • . .
49 Slat Planfotm Diagram
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VCK Pla..fcmn Diagran Flap Gap, CNe.rllaD:J, and Deflection DefinitialS 52 'l'jpical Inboard and OUtboard Flap Installations •
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53 Spa:wise Position of Pressure Orifice Inst.rtrcentation 76 54 Wing COOrCt-lise Pressure Orifice IDeations • • • • • 77 55 Slat and WUss Qlordwise Pressure Orifice Locations 78 56 Vt:K Chord.olise Pressure Orifice Locations (Typical for All Stations Except as Noted) ••••••• 78 57 'l\vo-Sagment Flap COOrarise Pressure Orifice Locations
(Typical for PDvIs at Tl = 20, 50, and 72.5 Perc::en t) ••• 79
58 Flaperon O1ordwise Pressure Orifice Locations
(At Tl = 30. 5 Percent) • • • • • • • • • • • •
59 M:ldel Installation in the NASA !ires 12-Foot Pressure Wind Tunnel ••••••••••• lbdel Installation in the NASA Iangley V/srOL llirld 'l'UrUle1 • • • • • • • • • • • • • • • • • • 61 lines l2-Foot Test Results • • • • • • • • • • • 83 cruise Win:] Configuration in the .Pines l2-Foot Pressure Wind Tunnel (Nacelles ard Pylcn Off) • • 88
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Page Figure 63 ~odynanic Cllaracteristics of Basic Cruise \'lin:J-Body and Effect of Uini-Tufts A. Lift and Pitcl'1in1 l-bnent • • • • •• 89 B. Drag ••••••• • • • • • • • • • • • 90 C. Lift/Draq Ratio • • • • • • • • • • • • • • •• 91 canpari5al of laoI- and High-Spee:i Characteristics for the 1CA Cc:nfiguration • • • • • • • • • • • • 92 65 Carrpariscn of High-Speed and I.cM-Speed Lift and Drag for Cruise Wing o:nfiguration • • • • • • • • • • • • •• 93
Mini -Tuft Photos for 9)Uise Wing-Body (Run 22)
A. ~ a 10.51 ••••••••••• B. Q = 12.57 (ClC~) ••••••• FRP
C. ~ = 13.55~ ••••••••••
D. ~ DO 15.44 •••••••••• 95 Experimental Cllordwise Pressure Distribltion for Cruise Wing (Nacelles and Pylons Off)
A. 'TRP = 12.57~ • • • • • • • • •• 96
B. ~ = 13.55 • • • • • • • • • 97
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C. ~ = 14.50 •• • • • • • • • • • • • •• 98 68 Variation of Section Lift Coefficient for the Cruise Win:J (Nacelles and Pylons Off, High Reynolds NUmber COndition) • • • • • • • • • • • • • • . • • • • • 100 Spanwise Variation of s...~on Lift Coefficient for Cruise Wing (Nacelles and Pylons Off, High Reyrolds Nunber COndi tioo) •••••••••••• Effect of Reyl'XJlds Nunber en Aerodynamic Characteristics of the Basic Cruise Wing-Body Configuration A. Lift and Pitchinq Itxrent • 102 B. IlJ:'ag •••• • • • • • • • • • • • • • • •
c. LiftjDrag Ratio • • • • • • • • • • • •
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71 Variation of Section Lift Coefficient for the Cruise Wing (Nacelles and Pylons Off, AtIoospheric Reynolds N\.lTi:)er CCXldi tioo) • • • • • • • • • • • • • • • • • • • • 105 Effect of Reyoo1ds NI.ItIber en the Spanwise Variation of C • • • • • • • • • • • • • • • • • • • • • • 106
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Effect of Nacelles, Py100s, ard Strakes en Aerodynamic Characteristics of the Cruise Win:J A. Lift and Pitc::hirx.J M:ment •••• B • Dl::'a::J •••••• • • • • • • • • • • • ,..
C. Lift,/Drag Ratio. • • • • D. ~ll.i.rg ~t •• • • • • • • • • • • • • • •
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Page Figcre Mini-Tuft Poot05 for the Cruise Win] with Nacelles, Pylons, and Strakes Attached • • 112
A. 'i?RP ... 12.5~ • • • • • • • •
B. ~ sa 13.58 • • • • • • • • • • • • •
c. ~ sa 14.55 • • • • • • • • • • • • • 113
D. Q D 16.56 • • • • • • • • •
FRP
75 ~tal Chord.dse Pressure Distribution for Cruise Win; (Nacelles, Pylons, atx1 Strakes On) • • • 114
A. ~ ... 12.580 (OCr'}.wi •••••••••
B. ~::& 13.58~ • • • • • • • • • • 115
c. ~ = 14.55 • • • • • • • • • • • • 116
D. ~ ... 15.55 • • • • • •• 117 76 Variation of Sect...i.on Lift Coefficient for the Cruise W~ with Nacelles, Pylons, and Strakes Attached Effect of Reynolds Nurber on Cruise ~'liD] with
Nacelle;;, Pylons, am Strakes Attached
A. Lift arx1 Pi t:chi.n:J M:::Irent
• • • • 120 B. Drag. • •• • •• 121 c. Lift;Drag Ratio.
• •• 122
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Experinental Cllordwise Pressure Distributioo for Cruise Wing with Nacelles~ Pylcns, and Strakes Attached ~iAC ... 1.14 x 10 ) o
A. '1:'RP == 8. , • • • • • • • • • • • • 123
B. ~ = 10.42 • • • • • • • • • • • • • • • 124
C. ~ = 12.44 • • • • • • • • • • • • • 125
79 Effect of Mach Nutber on Cruise Wir.g with Nacelles,
Pylons, am Strakes Attached
A. Lift mxi Pitc:h.i.n::J l-t:Jrent B. Drag • • 128 C. Lift,lDrag Ratio. • • •• 129 80 Variation of Section Lift Coefficient for the Cruise
Wing with Nacelles, PylalS, am Strakes Attached
~ = 0.32) •••••••••••••••••.• 81 Effects of Mach arxl Reynolds runber en Cruise Wing with Nacelles and Pylon Attached •••••••••••• 131 82 Tail-cn Cllaracteristics for the Cruise WiDJ with Nacelles, Pylons, and Strakes Attached A. Lift and Pitc::hirx:r Hanent •• • • • 132 B. Drag Cbefficient • • • • • • • • • • 133
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Page Figure Effect of ReyI'X)lds lbrDer en Cruise Wing with Nacelles, PylalS, Strakes and Horizontal Tail Attached Lift and Pit.chinq M:Irent •• A.
Dra.g •• • • • • • • • • • B.
Lift,lbraq Ratio • • • ••• C.
VCK and ~segment Flap Configuration :in the lines 12-Foot Pressure Wind Tunnel • • • • • • • • • • • • • • 138 ~ surface Midspan Ra:Jion of the va< with TWo-Segment Flap Conf~guration • • • • • • • • • • • • • 138
VCK Position Study for Landin:J Flaps
(Ove< = 5sO)
A. Lift and Pitch:i.n;r M::Jnent 140 B. lJl:'ag •• • • • • • • • • Variatial of Section Lift Coefficient for the va<
wi th ~ec}tlSnt Land.i.n:J Flaps
A. 142
· . . . . . . ~ = 55C/55C
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6\0( = 55D/550 • • • • • • • 143
B. · . . . . .
C. 144 6va< = 55A/55A · . . . . . .
88 VCK Position Stuiy for Lan:3.irg Flaps
(~ = 45°)
~. Lift and Pitc:hin:j M:rrent •• B. Drag . . . . . . . . . . . . . . . .
89 Variation of Section Lift Coefficient for the va< with ~-Segrrent Landi..n; Flaps A. ~:: 45F/45F . . . . .
· . . .
B. ~::: 45G/45G . .
· . . . . 148
C. 6VCK = 45E/45E . .
· . .
D. ~ = 45H/45H · . . . . . . . . . . .
E. Ova< = 45E/45E . . .
· . · . .
90 Mini-JI'uft Pootcs for va< with '1W:I-Se:]rrent I.andi.n; Flaps (Run 67) 0 A.
~ = 0.6° , • • • • • • • • •• B. ~ = 17.18 •... . . . . • .•.
C.
c\'IU> D 19.19: (~) • • • • ••• 154
D. '7RP = 21.14 •
E.
'7RP = 23.1°0 • • • • • • • • • • • • • 155
F.
'7RP = 24.06 • • • • • • • • • •
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Par;e Figure ExperiIrental crordwise Pressure Distribution for VCK with ~Sa:;mmt ~ Flaps
A. ~ = 0.60 , ••••••••••
B. ~ = 17.18 ••••••••••••••• 157
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C. ~ = 19.19~ ••• 158
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D. Q = 21.14 • • • . .
FRP 0
B. ~ = 23.10 • •• • • 16C
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F. ~ = 24.06 •••••
92 Variation of Section Lift Coefficient for the va<
with ~Segmmt Land.i.n:J Flaps (ova< = 45E/45G) • • • • • 162
VCK Position Study for Takeoff Flaps
(1Svtl< = 55 )
A. Lift and Pitchir~ ~t •• B. Dr'a.g • • • • • • • • • • • • • • • • • 94 Variation of Secticn Lift Coefficient ior the va< with ~Se:jIrent Takeoff Flaps A. Ova< = 550/550 • • • • • B. 6va< = 55C/SSC • • • • • • • • • VCK Posit.J.on Study for Takeoff Flaps
(6 -= 45 )
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'Va<A. Lift and Pitching l-tlrent B. Drag • 96 Variation oi Section Lift Coefficient for the \CK with 'IW:>-&:!gn'ent Takeoff Flaps A. 0\0< = 4SH/4SH • • • • • • • • • • • • 169
B. 6va< = 4SF/4SF • • • • • • 170
C. 6va< = 4SE/4SE • • • • • • • • • 171
D. 6va< = 4SG/45G 172
97 tti.ni~t 1?ootos for va< with 'J:\..1o-Segrrent Takeoff Flaps 0
A. ~ = 21.05 (ac~)
• • • 173
B. ~ = 23.00 ••••••
• • • • . 173 98 Exper:iIrental Cllom.n.se Pressure Distribution for va< with ~segrnent Takeoff Flaps A. ~ = 21.0SO • • • • • • • • • • • • • • 174
B. ~ = 23.00 • • • • • • • • • • • • • • 175
99 Variation of section Lift Coefficient for the \CK
with 'IW:>-SegIrent Takeoff Flaps (ova< = 45E/45G) • • • • • 176
Main Flap Position St:\]d\1 for 'l\.o-Segroont
Land.in:.J Flaps with VCK
A. Lift and Pitchin:J M::Jnent •••• B. Drag • • ••• • • • 179
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Figure Page 101 Aft Flap Deflection Study £or 'lWo-Segnent Landing Flaps with VCK
A. Lift am Pitd'lin:J M:rrent ••
B. Dr'a:.; •• • • • • • • • • • • • • • • Flapercn Deflection Stuiy for'lWo-Segnent
Iand:inJ Flaps with VCK
A. Lift aIXl Pitching l-tment ••• B. Ilr'ClI3' •• • • • • • • • • • • •
•
103 Main Flap Position and Aft Flap Deflection Studies for 'lW:r~t Flaps with VCK A. Lift and Pi tchi."'lg M:Jnent •• • • B. Drag Main Flap Posit:icn Study for ~t Takeof f Flaps Wl.th VCK A. Lift and Pitc:hing ~t •••• B. Ilra.g • • • • • • • • • • • • • 105 Aft Flap Deflection SttXiy for Two-Segr.ent
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Takeoff Flaps with \'0< ~/ A. Lift and Pitching ~t ••• B. Drag' • • • • • • • • • • • • • 106 Main Flap Position StOOy for Two-Se;ment Takooff Flaps with VCK A. Lift and Pitdri.ng M::ment B • IlI:'ag' •• • • • • • • • • • • • 107 Aft Flap Deflection 5tOOy for ~egrent Takeoff Flaps with va< A. Lift and PitdUng lbrent B. Dra::J • • • • • • • • • • • • • 108 tffect of va< Span and Nacelles, Pylons, and Strakes on Aerod-,tn.ami.c Cllaracteristics for 'l\..o-5egment I.andinJ Flaps A. Lift and Pi~ It:r.ent • • 194 B. IlJ:~ •• • • • • • • • • • • • • • • • • • • • 195 109 Etfect of va< Span and Nacelles, Pylons, and strakes on Aerodynanic Cllaracteristic:s for 'l\..o-5egment Takeoff Flaps
A. Lift aM Pi t:chin:J M::ment • • •
B. Dr'ag • • • • • • • • • • • • • • • • • • •
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Figure Page liO Effect of Peynolds Nurrber 00 1erodynamic Olaracteristic:s of the va< with Two-Segnent Landing Flaps Configuratic:n A. Lift and Pitching M:lrent • • • • • • • • 199 B. Dr'ag ••• • • • • • • • • • • • • • • • • • • 200 C. LiftjDrag Ratio • • • • • • • • • • • • • • • • 201 ill Effect of leynolds NlJtber 00 Aerodynamic Cllaracteristi c:s of the va< with Two-Segrrent Takeoff Flaps Calfiguration A. Lift and Pitching M:Ire."lt • • • • • • • • 202 B. Drag • • • • • • • • • • • • • • • • • • • • • 203 C. Lift/Dra:3 Ratio • • • • • • • • • • • • • . • • 204 li2 Effect of lbrizontal Tail. Detlectioo 00 Aerodyr.arnic Cllaracteristic:s of the VCK with. Two-Sepent Landing Flaps Configuration A. Lift and Pitc:::hi.n:J ~t:ment • • • • • • • • • 205 B. Drag • • • • • • • • • • • • • • • • 206 113 Effect of Reynolds Nmber for the va< with T\r,o-SaJrcent Iand.i.n;J Flaps (Horizontal Tail On) A. Lift and pi tdrl.ng lb':lent B. Dr'at] •••• • • • • • • • • • • • •
) 208
114 Effect of Horizental Tail Deflection on Aerodynamic Cllaracterist.l.c:s of the va< with Two-Segment Takeoff Flaps Configuration A. Lift and Pitdring M::ment • • • • • • • • 209 B. lll::'a:J • • • • • • • • • • • ~ • • •• • 210 SID:lle-Slot Flap Positioo Study for Landing Flaps with va< A. Lift and Pitc::hing M::ment B. Drag Sin:Jle-Slot Flap Position St1.1dy for
-I..andiIxJ Flaps with va<
A. Lift and pi tdrl.ng M::ment . .
B. lk'a.:J ••• • • • • • • • Sin;le-Slot Flap Position Study for Takeoff Flaps with VCK A. Lift and Pitc::hing M::ment • B. Dl:'~ •••• • • • • • • • Single-Slot Flap Position Study for Takeoff Flaps with VCK A. Lift and Pitchlng z..cnent • B. ~ •.•.•••••.•
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Figure Page ll9 Effect of InboaId va< Ratoval al Aeroc1ynan.ic Olaracteristics of the \0( with Single-Slot ~ Flaps CCXlfiguration A. Lift and Pitching It::Irent ••• B. Ill:'aq • • • • • • • • • • • • • • • 120 Effect of Beyrolds NLrrber al Aerodynanic Cllaracteristics of the vu< with SinJle-Slot Landing Flaps A. Lift and Pit.dti.n:J Ibrent • • • • 222 B. Dl:~ • • • • • • • • • • • • • • • • • • • • • 223 C. LiftjDraq Ratio • • • • • • • • • • • • • • 224 121 Effect of Beyro1ds NLrrber al Aerodynanic Cllaracteristics of the va< with Single-S10t Takeoff Flaps Configuration
A. Lift and Pi t:.chin:J lbtent •••••• • • 225
B. Drag • • • • • • • • • • • • • • • • • • • • • 226 ;" -~ -- ~ C. LiftjDrag Ratio • • • • • • • • • • • • • • • • 227 122 Effect of Horizontal Tail Deflectial on Aeroc1ynan.ic O1aracteristics of the VU< with Single-Slot Landi.rq Flaps Configuration
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A. Lift and Pitching l-h:"ent • • • • • 228 B. Drag • • • • • • • • • • • • • • 229 Effect of Horizontal Tail Deflection on Aerodynaml.c Olaracteristics of the \U( with Single-Slot Takeoff Flaps Configuration A. Lift and Pitdung Ucment B. Drag 124 Slat with 'l\.io-Segrrent LaOOing Flap Configuration 125 Slat Position Study for Landing Flaps 0 0
(.5 = 25 /35 )
.sr...ra' A. Lift and Pitching M:rrent • B. Dr~ ••••••••••• .
126 Variation of Section Lift Coefficient for the Slat with 'l\.1o-Segnent Landing Flaps
A. 6SLAT = 25A,135A • • • • • • •
B. 6SLhT = 25B/35B • • • • • • • • • • • • • •
127 Slat Position S~ for LaMi.n;J Flaps o (6 .. l5 /2S ) SLAT A. Lift and Pitc:::hin:J M:Irent • B. Drag Variation of Section Lift Coefficient for the Slat
j28
with ~-·Segrrent Landing Flaps (oSUd' = lSE/2SE) •••• 239
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-OJ Page ~i~E!: 129 Experimental Cx>rdwi.se Pressure DistributJ.on for Slat with 'I\>,o-Segment Unding Flaps
A. ~ = 16.14~ • • • • • • 240
B. ~ = 19.19 •• • • • • • 241
C. ~RP = 21.19 • • • • • • • • • • • 242
D. ~ = 23.09 • • • • • • • • • • • • • 243
130 Variation of Section Lift Coefficient for the Slat
with 'l\o.o-segment Landi.nJ Flaps (oSIAT = 15D/250) •••• 245
131 Slat Position S~ for Takeoff Flaps
(o~ = 15°/25 )
A. Lift ani Pitdting l-brent • B. Dl:'ag' • • • • • • • • • • • , Variation of Section Lift Coefficient for the Slat with 'l.\~-Segrrent Takeoff Flaps A. 0SLAT = lSF/2SF • • • • . • • • • • • • • . 248
B. 0SIAT = 15E/25E • • • • • • . • . • • • • • 249
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133 Experimental Cx>rd-lise Pressure Distribution for Slat with 'l\.o-Segrrent Takeoff Flaps
--
A. a = 15.90~ • • • • • •
FRP
B. ~ = 20.00 • • • • • • • • • •
• • • • 251 C. ~ = 24.03 • • • • • .
D. "'FRP = 27.91 ••••••
134 Variation of Section Lift Coefficient for the Slat
with 'l\.o-Segment Takeoff Flaps (.5 = 15D/25O)
• • • • 254 SlAT ] 35 Effect of Mach Nlt:ber on Aerodyr..:mric Characteristics of the Slat with Th'o-Se:;rrent Landing Flaps Config1..::ration A. Lift arxi Pitching rt:r.ent • • • • 256 B. Drc;tt;J • • • . • • • • • • • • • • • • • • • • • 257 C. Lift/Drdg Ratio • • • • • • • • • • • • • • • • 258 136 Ei::r~ct of Mach NuniJer on Aerodyn.aM..i.c Characteristics of the Slat with ~5egI:'eil.t Takeoff Flaps Configuration .l'\.. Lift and Pitching Narent • • • • • • 259 13. Dl:'a.g • • • • • • • • • • • • • • • • • • • 260 C. Lift/Drag RatJ.o • • • • • . • • . • • • • • • • 261 137 Effec: of Reyoolds Nunber en Aerodynanic Olaracteristics
of tht~ Slat WJ.th 'Tho-Seqrrent Land.i.D1 Flaps Configuration
A. Lift an:l Pitdting l1:Irent • • 26:.!
B. Dl:'ag' • • • • • • • • • • • • • • • • 263 C. LiftjDrag Ratio • • • • • • • • • • • • • • • • 2~4
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Page r'~gure 138 Effect of Reynolds Nurber on Aerodynanic Cllaracteristics of the Slat with 'l\.,o-5e;Irent Takeoff Flaps Configuration A. Lift and Pitching lDnent ••••• B. Drag C. LiftjDrag Ratio. • 139 Effect of Horizaltal Tail Def1ecticn al Aerodyncrnic Characteristics of the Slat with ~Segment
Lanc:lin3' Flaps Configuration
A. Lift ani Pitc::hin; M:lnent ••• B. Dx:'ag- •••••••••••• • • • • • • Effect of Reynolds Nunber on Aerodynanic Characteristics of the Slat with ~Segrrent Lmdl.n;J Flaps Configuration (Horizontal Tail On) A. Lift and Pitching Marent ••••••••••• B. Drag •• Effect of Horizontal Tail on Aerodyn.:lnic Cllaracteristics of the Slat with '.l\o.o-Segrrent Takooff Flaps Configuration A. Lift and pitdling M::xnent • • • • • • • • • • •
J
B. Drag • 142 Effect of Reynolds Nu.-ber on Aerodynamic Cllaracteristics of the Slat with Si.n:Jle-Slot Landing Flaps Configuration
A. Lift ani Pi t:c:hin:J a:rcent •• • • • • • • •
B. Drag C. LiftjDrag Ratio.
Effect of Reynolds Nurnt<.:!: 0:1 Aerod~'l-'dl1l.ic Cllaractenstics for the Slat with Single-Slot Tak ... .'off Flaps Configuratl.On A. Lift and pitching Hcr.ent •••••• B. Ilr:ag •••••• • • • • • • 279
c. Lift/Drag Fatl.:> • • • • • • • • • • • • • • • •
l-t4 Effect of Reynolds Nurber on Aerodynanic Characteristics of the Slat with Single-Slot Takeoff Flaps COnfiguration A. Lift and Pit:dling tbrent • • • • • 282 B. Drag ••••••• • • • • • • • • • • • • C. Lift.;'Drag' Ra.tio • • • • • • • • . • • • . • • • 284 145 Effect of Horizontal Tail Deflection CXl Aerodynamic Characteristics of the Slat with Single-Slot
I..and:in:J Flaps Configuration
A. Lift and Pitc:hi.ng M:xrent ••••• • • • • 285 B. Drag ••• 146 Effect or ~-Seg:rent Flap Deflection on Aerodynanic Characteristics with the Leading F.d:.Je Devices P.sooIled
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A. Lif t and Pi tc:::hin:j Marent • • • • • • • • • B. Drag xvi
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Pag'e Figur~ 147 Effect of Single-Slot Flap Deflection on Aerodynamic Characteristics with the Leadi..Il] B:1ge Devices Ratoved A. Lift and Pit:.chin;f Manent • • • • • • B. Drag ••• • • • • • • • • • • • • • • • •
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148 Effect of Reynolds Nutber on Aerodynanic Characteristics of the Flaps Retracted,l\'O< EXtended COnfiguration A. Lift an:i Pitclti.n.J Manent •••• B. Dl:'ag • • • • • • • • • • • • •
c. Lift/Drag Ratio • • • • • • • •
149 Slat with Retracted Flap Calfig'.lration 150 Effect of Horizontal Tail en Aerodyn.:mic <l1aracteristics of the Flaps Retracted/Slat EKterxied Configuration A. Lift and Pitching M:xrent • • • . • • • • • • • 295 B. Drag • • • • • • • • • • • • • • • • • • • • • 296 151 Effect of Inboard Slat Tr:izr\ Position en Aerodynamic Characteristics of the Flaps Retracted/Slat Ex:t:eroed Configuration (Horizontal Tail On)
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A. Lift and Pitch:in:J !t:ment • • • • 298 B. DJ:ag • • • • • • • • • • • • • • • • • • •• 299 152 Roll..in;J z.t:r:1ent Coefficient due to Aileron Deflection for Cruise Hing • • • • • • • • • • • • • • • • • • • • • 3011 153 RDllin3' Ibnent Coefficient due to Aileron Deflectioo for VCK with 'l\.o-5egment Takeoff Flaps •••• • • • 301 154 Rollin3' M::r"ent Coefficient due to Aileron Deflection for \0< with 'l\.O-SegIl'eIlt Landing Flaps • • • • • • • 302 155 Rolling M:Jrent Coefficient due to Spoiler Deflection for c:rtlise Wing • • • • • • • • • • • • • • • • • • • • • 103 156 RDllin3' ~,ent Coefficient due to Spoiler Deflection for Slat with 'l\.o-5egtrent 1'akeoff Flaps • • • • . • • 304 Rollin:] M::ment Coefficient due to Sp:>i1er Deflecticn for Slat with 'l\.o-Segm;mt Land:in;J Flaps • • • • • • • • • 305 Effect of Symretrical Spoiler Deflecticn on AerodynaMic Olaracteristics of the Slat with ~Se;;ment Takeoff Flaps Configuration A. Lift and Pitching Manent B. Drag
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I Page Figure 159 Effect of Synnetri.cal S[:oiler Deflection on Aerodynar.ti.c
O1aracteristics of the Slat with 'l\o.o-5egment I.and.in:J
Flaps Configuration A. Lift and Pitching It:ment • • • • • • • • • 309 B. Ilraq •••••••••••••••••• • • • 310 Lift Increment ale to Syrrrretrical S[:oiler Deflection for Slat with 'l.\o.o-Se:JIlellt Flap Configurations • • •• 311 Dra:J :Irlcl:'Emmt Due to Syrm'etrical Sp:liler Deflectio.'"l for Slat with ~-Segrrent Flap Configurations • • • • 312 162 Effect of Landing Gear on Drag and C4mx for Vt:K with 'l\.O-Se;rner.t Takeoff and Landing Flaps'""U5nfigurations 313 Tail-off Lift and Pitching M:Irent Cllaracteristics for va< an:! 'l\o.o-Segrtelt Flap COnfiguration • • • • • • • 314 164 Tail-off Drag O1aracterlstics for va< and 'lWo-Segrrent Flap Configuration • • • • ••
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165 Tail-Off Lift and Pitching Uc:ment Cllaracteristics for Slat an:! 'l\o.o-Segrcent Flap Configuration • • • • • • • • • 316 166 Tail-off Drag O1aracteristics for Slat an:!
~Segnent Flap Configuration •••••• 67 Tail-off Lift and Pitchin:J ~t Cllaracterlstics for va< and Single-Slot Flap Configuration • • • • • • • • • 318 168 Tail-off Drag Cllaracteristics for va< and Single-Slot Flap Configuration • • • • • • 169 'l'ail-off Lift and Pitc:hinq M:Irent Cllaracteristics for ~lat and Sin::Jle-Slot Flap Configuration •••••••• 320 170 Tail-off Drag Cllaracteristics for Slat and Sin::Jle-Slot Flap Configuration •••••• .
171 Tail-off C4mv CcJttlarison Be1:wea1 lldvanced Carmercial Aircraft arfl"OC-lO ••• • • • • • • • • • • • • • • • • 322
172 Effect of Ieadin:J and Trailing Ed;e High Lift
Configuration an c,. and S:. ••••• • • • • . . . 325
"M\x a=O 173 'l'r:imned Lift Curves for VCK and 'l.\o.o-Segtrent Flap Configuration •• • • • • • • • • • • • • • • • • • 326
'l'r:imred LID CUrves for va< and ~t
Flap Configuration • • • • • • • • • • • •
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Page Figure T.rilmai Lift curves for Slat ani 'l\-.'o-Segment Flap Configuration • • • • • • • • • • • • • 328
Trimred LID Curves for Slat and 'IWo-Segment
Flap COnfiguration • • • • • • • • • • • • • . .
177 Tri.mred Lift (''u.rVes for VCK ani Single-~not Flap Configuration • • • • • • • • • • • . • 178 Trimred L/D curves for VCK and Sin::]le-Slot Flap Configuration .• • • • • • • • • • • • • • • • •• 329 179 'l'rilrmed Lift Curves for Slat and Single-Slot Flap COnfiguration • • • • • • • • • • • • 180 '1'rimted L/D curves for Sldt arxl Su.gle-Slot Flap C'onfiguraticn • • • • • . • • • • • • • 181 TrJ.m't'ed L/D Cclrparison for the Takeoff Canf~guration • . 331
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182 TrirtuTed L/D CalFarl.son Beb-leen 'lW:)-Se:jn'ent and Sin::Jle-Slot Flaps for the Landing Configuration 183 Effect of Reynolds NurOer en \Q( and SIal with Tw:rSe:JlreI1t Flap Systan (Tal.l Off) •••••• 184 Effect of Hact. uurber on Slat with Two-Se:Jm.ent Flap and Clean-Hllt-J C!.t2t;X •••••••••• 185 Influence of NacelleJPylon and VCK Spanwise Extent an Tail-off C~ • • • • • • • • . . • • • • • • 186 Lift.in:] NSJrnanl'l Gearetry and Lift Calparison for Crtlise t ~iIl:J • • • • • • • • • • • • • • • . • • , 187 Ca:lparisan of Experimental Janeson Calculated Pressures. 336 188 CalFarison of Experirrenta1- and Calculated Spanwise Lift Distributions • • • • • • • . • • • . • • • • 337 189 Catparison of Exper:ilrental and Estimated. Maxl.i'lun-Lift Increrents for VCK and Flaps, and rlap-Lift Increrrent at Zero An:j1e of Attack • • • • • • • • • • •• 338 190 Cotp:lrison of Estir.ated and Experilrental C!.t1t\X ••••• 339 Cc:it'parison of Exper:imental and Estimated L/D
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Cllaracteristi<:s • . . . . . • • • . • . • . . .
xix
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Figure Page 192 Carparison of Experinental and Calculated Spanwise
Lift Distributions (Clean Leadi.nJ Edge with ~t
Flaps) •••••••••••.•••••••••••.• 341
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LIsr OF TABLES Table.
~ 1. SUI11l'ary at Hovable Surfaces 62 2.
Configuration Notation 63 3.
Dimenslonal rata 67 4. Slat Grid Notation 69 5. VCK Grid Notation • • 71 6. ~~in Flap Grid Notation • • • • 73 7. Aft Flap Grid Notation 8. Test Condltlons • • • •
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-..../ 9. FIgure Index for Experimental Data Presented
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SUNllARY This report presents tre results of tre design, evaluation, and \lind-tunnel testing of a low-speed high-lift model of a fuel-efficient advanced technology aircraft. This aircraft, derived from detalled system studies for a medium-range wide-body transport, incoqX)rated an advanced technology high-lift system. The results presented include: design trade studies, ,
design and analysis techniques, and results and analysis of tre wind tunnel
data. TOO experimental results included tre first low-speed l-u.gh Reynold5 numbe.: \o/ind tunnel test for such an advanced transport. Experirrental data 1ndude tre effects on thf: low-speed aerodynamic characteristics of slat and vdrlable camber l<rue'Jet (VCKj leading-edge devices, two-segment and slngle-slot tralling-edge flaps, nacelles, pylons, alleron, spollers, hor! zontal tail, and land1ng gear. !loth 1i.1ch and Reynolds number effects 'dere also studied for selected configurations •
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·rt-.e trade studies U1dicated .31gmficant .unprovenent in takeoff hEld lc:ngth, takeoff .:md landlng lift-to-drag r3tlO, dnd maXlmum 11ft coefficie.lt fo: a l:onfiquration 1ncorporatlng the awarll':d high Ilt:: sl'stems cexnp:.rcd \11 th a conventlonal high lift sysr.e:;I. I. reduct Ion 1n fuel burred "'Ja~ also obta i I'lCIl for thE: advanced configurat.l..l Cor.q:an sons of the ~st l.I11C1ted performance of tee selected configuration with eXlstlnq dlrcraft, lndlcated reduced fuel burned (per seut-mile), improved paylcud cata.;lty for hot/high operations, reduced takeoff and landing noise levels, iml reduced (per-seat) operatlng costs.
'rhe expenmental program .-;onflrm-;cl most of the estimated lo\·/-speecl aEro~'namic performance paramet~rs. 'l'he CrtllSe \ling achiev(::d a trimmed Illaximum 11ft coeffld.::::nt of 1.5 and a Hft-to-drag rut~o of 15.0. For the hlgh 11ft conflguratlons, the values of maximum lift coefficient were s iqni"ficantly improved \o/OOn compared to current alrcraft values. Typ~cal tnrmed naximum lilt coefflcients for takeoff and landing configuraticns Wf're 2.5& and 3.00 (for the VCK \o/ith two-segment flap configuration).
CGrr~sponding lift-to-d~ag ratios for tre takeoff and landing configurations
Here 10.2 and 8.7.
The land~ng configuration lift-to-drag ratio is a
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)' slgnificant improvement over tre previous gemration aircraft values. Tre slat leading edge device achieved maximum lift and lift-to-drag ratios superior to the VCK. Pitching moment trends for the va< were superior to tho~ obtaimd with tle slat.
From an optimization standpoint, results of the experimental program indicated the leading edge compomnts were more performance sensitive than the trailing edge devices. EXferinental data cbtaimd for the leading edge device retracted configuration indicated the leading edge device was of crucial importance to the attairarent of large maximum lift values. Without tle leading edge device, only srrall gains in maxUnum lift coeffici~nt could be obtaimd with the trailing edge flaps. As eXfected, tle two-~gment flap was superior in rraximum lift ccefficient and flap lift increment. Trimmed polar comparisons indicated equivalent lift-to-drag envelopes for the takeoff flap settings. For equivalent values of approach speed, the Ilft-to-drag ratio for the two-segment flap was superior to the
,)
corresponding single-slot flap value.
Aileron studies indicated that, for all flap settings, mgative deflections (trailing-edge-up) were more effective than positive (trailing-edge-down) deflections. The effect of spoiler deflection on roll characteristics indicated improved effectiveness as the flap deflection was increased.
Symmetrical spoiler deflections, for takeoff and landing flap deflection, showed tle spoilers to be very effective in reducing lift and increasing drag. The drag for the landing gear was essentially the same for the takeoff and landing configurations. The lanr.ing gear caused a slight reduction in maxllnum lift ccefficient for the landing configuration.
Comparisons of experUnental data with estimated values were generally in good agreement, lending confidence to the results of the trade studies.
AnalysiS of too data has highlighted areas where continued efforts could result in further improverrents. Trese areas include lift-to-drag ratio for takeoff configurations, pitching moment for the high lift configurations, and ground-effect characteristics. Specific test items are suggested for this continued c1eveloprent.
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SYIiOOIS '1'he longi.tudina 1 aer"X1yna:'ll.c characteristics presented in thJ s pap::r tirf referred to the stabill ty-axis system. Force data were reduced to cceffjci.:::nr. form barel1 un t~~ t[a~zoidal \'ling area. iUJ dilrensional values are given In bo ... h International System of Units (51) and U.S. Custor.,ary Uni ts, tl'e princlpal rr.E:asurcr.lents and calculations using the latter.
Creffl<.:ients dnd synbols usecl !1erein :lre defired as follows: AI' w~ng u~ct ratlu b ... 1:ng span drag ccefflclent Cn C zero-litt protlle drag (coefflcient) Dp co;rbired flcap <1!ld ie..llhnq f:clge de:Vtce fa (uS! t.e drat] CDpARASI'IE F'I.AP + L.E • (ca:fficlent)
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11ft ccefflciel1t l:L CLr,II\."{ r .axirnum 11ft coofficlent CLm tail-uri -l: It- cn:!t i leIer.1.
CUl'RTIl trll!il1ed llfr c~tficll"nt CLa::O 11ft coefficient at zero degree£> dngle '''If attack rolling marent coefflclent ('1 two-dilrensional £ectlOral lift coefficwnt Cl2-D C1 _ three-dimenS10.a 1 rectlonaj lift coeffinent D
(m pitchin~.j r.10mf::nt c02ffic:i~nt
pressuce coeiflClcnt Cp CPCRTI pressure coefficient corresFOnding to 1ncal fl Ckl llar.h number equal to 1.0 h',lnimum pressure coefficient CPruN
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roc direct operating cost e induced drag efficiency factor FCK fixed camber Krueger (flap) fuselage reference plane ERP smpe factor for boundary layer profile H rrean aerodyramic chord of too horizontal tail HMAC incidence angle between the hozizontal tail and the iH fuselage reference plane, positive trailing edge d~ (de9) lift-to-drag ratio L/D maximlUll lift-to-drag ratio (LID) f;12\X trimmed lift-to-drag ratio (L/D)TRll1 !-lAC rrean aerodynamic chord W\DAAH mul tielerrent aufoil design and ana.lysis method O.H.
overmng R ;r.1AC Reynolds number tased on r'lAC
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RSS relaxed static stability S wing area urorcach speed VAPPIDACH \"CI~ variable camber Krueger (flap) Lift-off Speed - tl-e sp<;ed at i-Iffich the airplane fl.rst IfW becorres airborne rrean aerodynamic chord of the vertical tau JiJi'l.C Air HinimlUll Control Sr:eed - the rninimlIDl flight speed at Vn: which, when tre critical engine is suddenly mde inoperativE', it is possible to recover control of the airplane and maintain straight flight either with zero ya\., or with an angle of l:ank of not: more than five degrees MinllnUffi Unstick Speed - too speed at and above which the VI·l[] airplane can safely lift off the ground and continue the takeoff Rotation Sr:eed - the speed at ';lhich the pilot begins to VI< rotate the airplane to the lift-off altittXle Stalling Sp2ed - the rninirnlUll steady flight speed at which
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tre airplane is controllable
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I !~nqll1E' fLlilure Sfeen - the S-p?E:c1 at \Jhil h thE: cntlczJ Vef engine i~ as~umed to fail Takeoff ~cisi(m Sp:ed - the ~ed \OlhlCh the pilnt llseS "1 as 3. reference in deciding whether to continue the tal~eoff or to abort 11- 'takeoff Safety S~ed - the s~ed at 35-foot height "lith one engine ir.o~rati·:e ppp \"lnq referer.ex: plane t;u!:s \vinq W1rer slat surface ,. :;111;1'.11::"£ loJlr:q ·t.:ltlCr.
v r:hor'}"lise \11I1q staticn . 'I Ct
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C!nq1e oi c:.ttad" at CLrlll-J\ ~-!A.X anCJle of altar!. of t.h~ !'IIf'oelaqe rctt:r~i1ce plane, positIve uPP.P nose up (r3eq) a ar~ql~ c)j :; t t.3t I t "1 ~\. 14(1 j 1 ft I ( I
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r rhherlra1 anqlt: J Cn :UG:r: ~r.du~d rir-:'] 17'r:rr::rent ('\ r. to fl.;,' ( .. ::::.effH~lenti
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1 i:rrua: i 11r3 I '(;:!l'lent" "',e t(1 ffi(:l;lle :(" :E:frlCHmt:l
. IIJI'DIIO~D f!:'\CEUR nisrel13neo1l5 '1ra'l If'CrelT'.ent (("~fficlentl l '7n~C .1 __ 1.1 Un drag increr.ent (CoefL"'I:teI't) . ''mm .1- i"axi;-,Ui"\ lift mcren=~.'" .1ue 1.0 flan (coeffICIent) ( ! ''PY.
.:l ra:<i.rlirl 11 f~ ir.C"rr.n2n+- Pile ~ c interfere!'lO? (coefficif'ont) CLl'AX UrIFRFEHEtXE .1 • incre~e.I~. In 11ft dUe to :-rir. (c::'k r,-iClr:nt) ··r"nm j,~ tvo-dinenslOnal r.a::dnulo, 11ft InCrei~l2nt rille to flap lrra~u'p (coeff icient) .1 L' b.Q-ch::enmcrul ii'a;,;irnum 11ft Increrrent duE': to leRni!:n k 11'a.'"r~. E.
edqe (:ie'rice (coefficient) _1 (" (:.:=r.
olo-dirensICI'.al , i ft increment at zero angle of c: ttack (coeffIcient)
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15'" bOlmdary layer displacement thickness
--
o f/l.F'J' aft flap deflection angle, positive for trailing edge down
(deg) effective flap reflection (OPrfAIN + 0.5 0PAFr) angle °FEFP
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flat:eron reflection angle, positive for trailing edge dovm FF1.AFEIDN (reg) 1 main flap deflection angle, positive for trailing edge °Pr1All dovm (reg) aSp spo~ler deflect-.ion angle (symretrical), negabve for trailing edge up (deg) c5 sPr.lJ left-hand spoiler deflection angle, negative for trailing edge up (deg) left-hand aileron deflection angle, positive for trdiling
°au!
edge dovm (deg)
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('Ire of the Intest technological advances t.o be used in current aircraft de~iqn stunies 13 the supercritical airfoil conceived by Dr. R. T.
flhitcomb of the· Pational Aeronaut1cs and Space Administration. The research on f>lIpercritical \vings over the Plst fe\-! years by Ill\SA and the industry has sha.m that definite perfornance advantages are to be gained by applYlnc; ~his technology to future transport aircraft. Accordingly, tre J10uglas Aircraft emPlny, under the Energy-Efficient Transport (EET) component of the rlA...~]\. J''.ircraft Enerqy Efficiency (ACEE) proqram, has been / studyinq the supercntical winq in connection \lith the DC'-X-200, a ~()O-passenIJer, \nde-body, medium-range transport. The results of Cl
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systerlatic Iving development study for the DC-X-200 are presented in Referenc~ 1. ~'lhile r:lUcr ;'leveloprrot'nt \'lOrk still rE'mains to be done to refine the high-speed craracterir;ticE', the cOJrlparisons of the available expermental data \lith al'':llytic:.:,l predIctions establlsh confirl?nre in the nlqh-~ed cEsign concepts. On the other hanel, no e'lu1valent data hase
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e:<ists for high 1 ift configura tions. Horeover, establishinq the t:rue potent.ial of the high lift system for an ai rcraft like thE: DC-X-200 requires slClnificant ('!;.q:-.erll"ental optimization. 'Ihu:; requires lea,hnq" and trailing-edge POS1tl('i: I''V] c1eflect1on f>tudies lnvol',ing a rather ';mplex \lind tunnel rr.orel and all extE'J'1siw: test progr~m.
'!'raE: purpose of this report is to present the lesults of an initial study, made under the BET Program, to develcp the hiqh-lift system for an .1Ovanred transport aircraft with a I:l,,]h-aspect-ratio supercritical wing.
It presents the results of n..)uqlas-sponsoreri trade studies made to defire the an.lantages tl:at an adv:mreci hiqh-lift system could bring to aircraft :: ike rhe OC-X-200. Ir- then r1escrims the desl<jn and \olind tunnel tests of a model to develop th~ advanced high-lift system. The wind tunnel tests
\lere made uS1nC] 3 11.7 percent scale monel of the DC-X-200 ha~Tinq a In.:-
aspect-ratio sur:crcritical wing, representing an advana=c1 design developed in thf- hi<)h-spe€:c1 sludies of Reference ]. Tl e model \vas tested with a 'lariety of leading- and t ralling-edge devices 4ncluding a leading-edqe :>lat, ':arlable-car:;ber Krlleqers, a trCliling-edge two-segment flap, a conventIonal vane flap, and a plain flap. Spoiler and aileron
)
~ffectiveress and the effects of enqire nacelles, pylons, and landing gear
i;\
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\yere also investigated. The tests were made in the Ames l2-foot tunnel at ~ach mn'lbers fran 0.20 to 0.32 and Reynolds numbers, based on \'ling mean aerodynamic chord, from 2.89 million to 5.12 million. Six-canponent forces and manents and static pressure distrihutions were obtained. Flow-visualization photographs using flourescent mini-tufts were also obtained.
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preLnUNARY TRAIE sruoIES BackgrOtUld At the time the NASA Aircraft Erergy Efficient Transport program (ACEE) and lts Energy-Efficient Transport (EET) component was initiated, the Douglas t Aircraft Company \-las studying the DC-X-200, a 200-passenger, wide-body, medium-range transport. Tru environment in which these studies \'lere being made included rapid inflation, concern over escalating fuel prices and decreasing availability, increasingly stringent noise regulations, and deregulation of the airlire inoostry by government. The influences of these factors on the re\'J design were: 1. Due to increared costs of prorucing re\-l aircraft, advanced technolo- gies would be required to obtain, from the standpoint of economics, performance superior to older, less expensive designs.
2. The rapidly esc.."llatwg fuel pnces and uncertaln availability
J
required that fuel-efficient vehicles would be required rather than thore exhibiting increased speed.
3. Greater aerodynamic p:L E :!1Tlance in terms of la-l-speed lift/drag \-Iould be required to supplement engire technology in meeting rew noise requirements.
4. Improved route flexibility througil better takeoff and landing perforrrance would be required to weet the potentlal government deregulation of the airlire inrustry.
AS a result of these conslderations the DC-X-200 fam~ly \-/as configured to employ a high-aspect-ratio supercritical ,·ling (Reference 1), and incorporated an advanced high lift system. The configuration also employed relaxed static stabllity (RSS) to further improve Lhe fuel efficiency and economlCS (Reference 2). In this report, a particular version of this family of aircraft (DC-X-200 nodel D969N-21) was selected for la-l-speed
exper:unental studies and is referred to as tre Advanced Commercial Aircraft
) (ACA) 1n other portions of this report. Other technology development programs related to advanced transports are currently being studied at Douglas under the ACEE program and are reported in Reference 3.
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Fundamentally, the supercritical airfoil can be used to generate qreater lift for a given thickness and drag than a conventional airfoil. Due to the emphasis on fuel-efficiency and direct operating cost (DOC), the manner in t/hich the added benefits of supercritical airfoil technology have been utilized are, basically, to increase ~ing-thickness-to-chord ratio and ast:ect ratio. '!be increase in thickness-to-chord ratio results in a more efficient wing structuraJ system, which, in turn, offsets the weiqht of the higher aspect ratio. Within these ground rules, however, many other design variables must be considered, which, almost without exception, have both favorable and unfavorable aspects when considering the total aircraft design. To introduce the differences in low-speed characteristics resulting fran this awlication of supercritical airfOl.l technology, Figure 1 presents a comparison of the Im'l-speed performance of a current \>lide body trans{X)rt and one of the advanced configurations of the design studies. The former aircraft incor{X)rated a conventional wing with an aspect ratio of 6.8, and a high lift system canposed of a slat and vane-flap canponents. The advanced configuration lIsed supercritical airfoils in the ;nanner previously
~
rliscussed, with an aspect ratio of 12, and an advanced high lift system HIGH LIFT SYSTEM CONFIGURATION ASPECT RATIO AIRFOIL SECTIONS CONVENTIONAL CURRENT 68 SLAT + VANE FLAP ADVANCED 120 SUPERCRITICAL VCK + TWO·SEGMENT FLAP 46% 44% --r-- --r-- 35% 0 0 w W w .... ....
...
U U U Z Z Z Z w Z z w w <t <t I.: <t a:: a:: a:: > a:: > a: > 0 0 0 ::l ::l ::l c( c( <t U U U LANDING .. LANDING LID TAKEOFF LID L MAX 'fP611' FIGURE 1. COMPARISON OF LOW·SPEED CHARACTERISTICS FOR CURRENT AND ADVANCED
)
TRANSPORTS
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composed of a variable camber Krueqer (VCR) and large-chord, high-extension, t\vo-segrrent, flap system. This combination represents the upper bound in lm'l-speed performance for the conf igurations studies and shO\o/s that significant improverrents are predicted for the advanced configuration.
,I High S}:eed \,1ing Parameters Naturally, the p3rameters related to the hiqh-~ed design problem are very siqnificant for this class of aircraft and closely affect the low-speed characterisitcs. The high sreen design and wind tunnel development of the high aspect ratio supercritical ".,in'1 develop!1ent is reported in Reference 1.
'i'he most important Oesigrl considerations fran this reference are qiven belO\v.
Spanwise Distribution of Lift r or SWn Loading ]I~thouCJh an elliptj cal span loading offers the 1m/est induced drag, the optlmum loadlnll, considering the comb1.ned aerodynamic and structural characteristl.cs, is u~ually one "hich has s~ degree of \,rashout. Since, at
J
cruise 11ach number, the initial sep3ration \'lhich t.:letf'rntines buffet unset Ilsual]" occurs on the outboard \Jing JX:Inel, it 1S undesirable tc a110\.1 the local velocitles in thlS region tC"o tecome too high. T11l.nJllng the outboard \ling can alleviate this _ J. :.I3tlon, but at the expense of a penalty 1.n ~lelght. In the final analysis the choice of span leading is a function of not only induced drag but also \·linCj "'leight, and lou-speed and hl.gh-speed clean wing separation characteristics (Le., stall progression and buffet boundary) • SpanW1se Distributlon of Tbickness For the same upper surface pressure distribution, modifications in the thickress result In cbanges to the sect1.on lift values. The decisions on the distribution of lift must be made in conjunction with the decisions on thickress. On the inbcard wing the thickress near the root is affected by such considerations as the depth required for the landing gear, and the volume needed for fue 1. In additlon, the cholce of the spanwi se dlstrlbution of thlckness considers not only the combination of lift and thickress required to meet the cruise performance, but also the impact of the distribution on lo\'/-speed performance in both the clean and high lift
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/ modes. Since, for a glven planform, tte weight of tte wing box is largely a functlon of the lift and the thickress-to-chord ratio, tre aerodynamic and weight characteristics must be considered in lUlison before a £inal decislon can re reacred.
Airfoll Characteristics The choice of chordwire and spanwire airfoil characteristics introdlces many other variables. For the basic outboard airfoil, decisions must be made regarding the leading edge radius and the amount of aft camber. Blunt leading edges are desirable from tte standfX)int of supercritical development at cruise, and maximum lift at low sveeds. '!bey are unoosirable from the standpoint of drag creep and, in some cases, lift loss ~t stall for low speed conditlons. The latter trend can result from the rapid change in curvature rear tre leading edge associated with a large nose radius and a lack of nose chamber. Large adverre pressure gradients aft of the suctlOn peak can induce a rapid change fran a trailing edge reparation to a leading edge separation. Highly aft-cambered airfoils are desirable from the
)
standpoint of achieving good cruise characteristics at high 11ft coefficients, but they have high regative pitclllng marents which, for some configurations, can result in excessively high trim drag. LO\'l 11lt coefficient (dive) characteristics at very high Nach nu:mt::ers can also be urecceptable with too much aft camber, particularly where outboard lateral control devices are used. The spanwise distribution of aft camber also presents a desiCJT1 challenge as it is difficult to carry a large amount (f lift aft on the chord near the root and, at the same time, counteract the Loot effect to maintain satisfactory inboard isobar characteristics.
Fact~rs Influencing Selection of High Lift Systems If the aircraft's wing area is sized by a low speed requirement, for example, approach speed, the wing area may vary by hundreds of square feet eependinq on the utilization of simple or advanced high lift systems. If an advanced high lift system is utilized, the resulting wing area can he relatively smaller and a resulting weight advantag':! should be achieved.
1'he 10\'1 speed characteristics are also influenced by the planform effects of
)
the su~rcritical \-:inq. A small, high aspect ratio wing has a relatively
~
•• ;.; ~) small root chord \/hich requires siqnifh:ant exteru;io~ 11. orrer ro hOllSP. the landing qear. '1l1e use of relaxed static stability IP.SS) I which moves the required qear position further aft relative to the \Ylr.g, further agqravates the problem of available inboard flap chord. 'Ihis re::;ults in a need for an advanced inboa rd high lift flap system. 'Ihe large trailing edge ey.tension unE\oJeeps a significant portion of the inboard \-/ing. This makes the job of the high-speed designer :iore difficult in terms of eliminating root effect, and maintaimnq S'lJeep effectbteress. One method of achieving proper root characteristics ilt hiqh speen IInner these conditions is to increase the leadinq edge sucep in this area. '11n5 impacts the ION speed characteristics by effectively renucin<] inhoard Elap-chord/\-ling-chord ratie and tends to reduce local lift curve slore makin<1 higher inboard ~;talling angles at 10\-1 sp=ed.
The basic airfoil sectlon <:1180 i!!l~cts the> high lift system. A coml=6[lson of the su;:crcritiC3l and ·:-onlJE:ntiona1 a i rfolls, shown in Fiqure ~, indicates slgnifi'car.t differences ir. cllrT.Oll shape. f':.:'(;ordir.gly, the high lift
)
systems cor.tpatible wi tl- these blo ai rtoil shapes have cHfferent de£ign impllcations. For the same ratir of wing chorn to flap chorn, an exanination of: the tral Iln'] edge reCJlon shO\<ls that a flap [.':stem for the supercritl cal airEai 1 "v' 1 lx, thinner and fX'S~Sfi " slgnificant amount of "huilt-w" c:lmber. At the leading edC}E, the increased nose radius of the supercubcal alrfoil \·rill yield larger ma;amum hft coffic~ents aL low
speeds than those of tre conventiorel EEct-ion. Consequently, larger rna;dmum
11ft values c.an re obtalrl:d ',ath the high lift systeln deflected.
In addlt:'..cn to \-l1ng gCC'!loE::tnc pal."; .. ~teC!. there are a numoer of other factors that mfluence the !:e 1 "''(..uon d;: .1 high hft system for an advanced canrercial I ransport. ::!Y:!:3e ilr~ S!"!OVl:1 in FiCJure 1. TO=! most fundamental requirements .:.Ire tnose 0f rne op8cLltor. o')f primary concern to the a1rline i5 tt.e GEOMETRY ~U?EMr.R.llr."'l ~ CONVErHIONAI ~ ____ ~
---- ~
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FIGURE 2 SUPERCRITICAl AND CONVENTIONAL AIRFOIL COMPARISON
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....; -; cap:1bility of an economically profl.table opeo:ation over its route structure with realistic passenger and cargo cap:lcities. In tOO low speed area, the resulting requirements include takeoff and landing field lengths, and in some cases a specified approach speed for the design mission. Naturally, for a given class of transport aircraft, not all airlines have similar requl.rements concerning low speed performance. 'Ibis aeXls another dimension to the selection of the high lift system in that acceptable low speed performance must be achieved for as many potential airline custaners as possl.ble.
OPERATOR REOUIREMENTS o TAKEOFF FIELD LENGTH o LANDING FIELD LENGTH o CRUISE CONFIGURATION ECONOMICS o DEVELOPMENT COSTS o OPERATING COSTS MECHANICAL COMPLEXITY o RELIABILITY o MAINTENANCE REOUIREMENTS GOVERNMENT REGULATIONS o FAR PART 25
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o FAR PART 36 FIGURE 3. FACTORS INFLUENCING HIGH LIFT SYSTEM DEFINITION
'rhe economic factors influence not only the airlines in terms of roc, but
also the rranufacturer in tenus of th: developrent costs of new hardware and technolo~. Improved high lift systems can also influence fuel efficiency
.-
wmch, in trese days of escalating 011 prices, has a significant impact on dlrline profltability.
Reliabillty and reintenance requirements also influence the high lift system deflmtlOn. Tre deslred high lift mechanical system must be lightweight, have a lllgh degree of rel1abill.ty and a minimum of in-service maintenance.
rnother signifiClnt factor in high lift system selection is the government regulations concerning transport aircraft. Two important documents \'lhich speclfy requirerrents for the low speed operation of transport aircraft are Federal Au Pegulations (FAR) Part 25 anc1 36. FAR Part 25 is concerned with
.--
the operatlOml factors (i.e., critical speeds, distances, flying qualities ,-te.) and FAR Part 36 specifies rules concerning acceptable noise characteristics dlring takeoff and landing.
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~ OP.1GINAL PAGE IS
')
OF POOR QUALlTY Figure 4 presertR the 3peed detl mtions pertuent to tre takeoff proftle.
Also shown in F:qure 5 15 urc d~ftmclOn of takeoff field length. T!us 1S the cn t1cal t::ngire-cU1: takeoff '1istance and is the l.hstance frc.:t the start of takeoff to a point 35-feet dbov(> tn~ nJI'lolay at tlc V2 speed, al:>suming an l!ngine failure to be recognized dt: the decision speed, VI. The other deiinltion of ta~eoft distance 1S 115 percent of the horizontal d1stance fror.l the start of t~keo::-: to the [Allnt at ~'hich the auplane is 35-feet above the takeoff smface Wl tn a 11 enc:JiI'es opi.:!"ating. The final takeoff distance 15 specified to be the greater of the two tk:flred tart:off distances.
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~ ... (;..~ FIGURE 4. TAKEOFF DEFINITIONS fr>.v~sn,.,'· UlO SElJrUiJf
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) FIGURE 5. TAKEOFF FLIGHTPATH PROFILE
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ORIGINAL FA:';: OF. POOR QUAL!".!'
The values required for the various takeoff climb gradients are sununarized 1 n Figure 5. Also sha.m in Figure 6 is the mt path. 'This corresponds to a reduction in gradient of 0.8, 0.9, and 1.0 percent for two, three, and four engine aircraft, respectively. Figure 7 presents not only the takeoff, but also, tre awrcach and landing climb requirements. Also shown in Figure 6, are the corresponding configurations for trailing edge flaps, leading edge device, landing gear, awropriate speed, thrust setting, and the number of engims inoperative.
-
NUMBER OF THRUST ENGIN!:S Sl'EEO I
I REOUIRED GRADIENTS FLAPS SLATS GEAR SETTING INOPERATI"~
I __ ENGINE 2 ENGINE 3-ENGINE I AIRCRAFT AIRCRAFT AIRCRAFT
r- W."O~'''~ .. ,.~n .. ,~".n "'.«'n
I , fAKEQFF 10 I.\K~Off FIR~r SE~MENr POSITIVE 03 as tAKEOFf DOWN ~LO TAKEOfF S£CQNOSEGMENT 24 27 30 tAKEOFF TAKEOFF UP V MAx 1 rAKEO~f FINAL SEGME"T 1 2 1 !:I I 1 AE: TRACIEO AETRACTEO ....,2'3 "'s CONTINUOUS ".
APPROA.CH TO ,
~"'HOACH CI 'M8 1 2 1 l' 11 ~PfRO"CH u>
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L.ANDING .... UMB 3.! 31 .J2 lANDtf',G LANOING .... 3V 10 ~J DO"''' s
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FIGURE 6. CLIMB GRADI ENT SUMMARY APPROACH 3 DEG GLIDESLOPE ~'lc;,~""'\
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FAR PART 36 (STAGE 2)
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--~-- -- ICAO ANNEX 16. CHAPTER 3 ANO
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FAR PART 36. STAGE 3 (NEW TYPE
, ~ - '
DESIGNS)
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* 035 N MI F.:lR 4 ENGINE DESIl..:'IIS
) FIGURE 7. AIRCRAFT NOISE CERTIFICATION REFERENCE LOCATIONS
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The landing distance re<luUehlents at the intended destimtion or alternate airport are also deflned in the fo'AR, Parr. 25. Other more general requirements are also contaired in FAR 25 which lIDpact the 10\>1 speed requuements. These are related to controllability, rnareuverability, trim, stability, and stalling characteristics •
•
'I're Federal Aviation Fcgw.:ltions Part 36 establish noise liraits for thr~e different reference locations as shown in Figure 7. 'lb::! requirement for FAA Stage 2 limits (for derivative aircraft) and the more st~ingent limits specified by ICfD Annex 16, Chapter 3 and FAR Part 36, Stage 3 for ne\-/ transport designs are sham in Figure 8.
The takeoff and land~ng performance has a strong sensitiv~ty to the pertinent low spc-ed lift and drag characteristics. In add~tion to reducing takeoff and landing field lengths, reductions in takeoff, approach, and landing speeds are dHect.l'l related to increased aircraft 11ft ccefficient (CL) at a given attitude anu to ~ncreased aircraft ma~draum 1 itt coefficient (C~). Tbe clirrb gralhent 13 strongly dep;:ndent or. the lift-to-drag rctw
)
(LiD) for t~ clirnbout configuration. Aircraft noir:.e level s are directly related to the thrust required, which is in turn related to the drag level of the aircraft. During l 'eoffr at a giveu climb gradlent, the larger ·'alup.s of LiD resull 1.n reduced n013e levels for the surrounding aiq:x>rt conrnunity. Due to the levIer no~se levels operational flexibility for the ,1irlire coul,1 be ~ncreased, since the aircraft could operate profitably out of more noise-critical airports.
Trade Studi.::s A~. a baS1S for the ('ontract '(lork, Douglas funded trade studies \;ere perfGrmed and 1.nl!luded variatloos of: a1.rfo1.1 sectlon (conventional versus supercritical), as},X:ct ratio, high lift system capabihty and type, wing sweep, twist, camber, thlckress, planforra, and various degrees of p,sS.
P-e13ted to the precedlng studles, an EET funded study was conducted to evaluate the effect of an advanced high lift systera on an aircraft typical of this cldSS of transport aircraft. Tre study was cGncerred \-/Hh 10\" speed
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w > I- 90 U w u..
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w
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s" 40 60 80 100 400 600 800 1000 TAKEOFF GROSS V.E1GHT (1000 LSI fiGURE 8. FAR PART 36, ICAO ANNEX 16, CHAPTER 3 (NEW DESIGNS) NOISE LEVE!. REQUIREMENTS
)-
....
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OF rG~:;' l-.:;~"":"'; • ~rformance and fuel efficiency. For this study the a~ct ratlO (AR) .... 'as 12. Figure 9 presents the configurations assessed in,this trade study.
r'lissio:1 requirements for thls study were: a ran~ of 5,556 kIn (3,000 n.mi), a payload of 20,085 kg (44,250 Ib), initial cruise altitude of 10,363 m (34,000 ft), d cruire nach numt:er of 0.78, a VAPPIOACH of not nore than 64.8 rrv's (126 kt) and a takeoff field length of not more than 2438 m (8, 000 ft).
For the advGlnced and con':entional high lift systems the location of the real spar was reld constant resulting In similar chord ratios for the trailing edge flap system.
1f"BO~RD liCK·
[ADVANCED HIGH LIFT SYSTEM I
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FIGURE 9 HIGH LIFT :;YSTEM CONFIGUR~TION COMPARISON The cO:1'lcntlonal I'>~h llft syster:; lncorporates .J. clrcular arc motion vare flap configtlotion. Ths lnboard and outboard flaps are reparated by a hl.gh speed aileron Hhl.ch 1S und:flected 1n the high lift n:od=. The leading edge cleVl.ce COns1StS at an 1nl;mrd and outboard slat, with appropriate t.:ll~eoff .rod 1;1Od1ng posit1cns.
'i'he aU"anc.eCl h1gh lift system ~us ':10 lnboard and outboord t...!o-regnent flap system \'1ith large: e;;t.::ns:;..;)n c6oab111ty. A flaperon, \lhl.ch is deflected
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during takeoff and landing, is located between the inboard and outboard flaps. Tre flaperon translates in tre saIre rranrer as the main flap element of the inboard and outboard two-segment flap system. The leading edge device consists of a Va< for the inboard and outboard FOrtion of the wing.
The results of the sizing study are presented in Figure 10. The conventional high lift configuration was sized by the approach speed requirement. The resulting initial cruise altitude is Slightly higher than the mission requueroont of 34, 000 feet.
ADVANCED CONVENTIONAL 2 2 20532 m 20299 m 12.185 sa FTI 12.210 sa FTI WING AREA 137.393 kg 1302.900 LSI 138.073 kg 1304.400 LSI TOGW 82.135 kg 1181.077 LSI OEW 81.627 kg 1179.956 LS I 134.000 FTI 10.424 m 134,200 FTI INITIAL CRUISE At.T1TUDE 10363 m V APPROACH 619 m/S 11204KEASI 6-t 8 m:S 1126KEASI 2179m TOFL 1.168 m 15.800 FTI 17.150 FTI -- -- -- FIGURE 10. CONVENTIONAL AND ADVANCED HIGH LIFT SYSTEM STUDY CONFIGURATIONS
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'l'he wing for the configuration with tIle advanced high lift system was sized b~l the inltial crU1SC altl.tude requl.rement. A smaller wing area was deterr:ared for the advanced configuration. The configuration incor!X)rating the advan~d high l1ft s~'stem has a signiilcantly reduced takeoff field iength. Moreover, the improved LID cnaracteristics would reduce noise levels fer takeoff and landing operations.
'rhe effect of the adva::ced configuration on fuel burred over various ran~ missions l.5 shewn in Flgure 11. For the shorter range missions, the performance calculations included determination of the most favorable l.nitial cruise altl.tude for the corresponding reduced initial cruise weight.
At:. the design range, fuel savings of 0.6 percent were obtaired. However, average stage lengths for aircraft of this class would be expected to be of the order of 750 to 1000 nautical miles. At these redJced ran~ values, fuel savl.ngs of the order of 1.6 to 1.3 percent of fuel burned were ) mdicated.
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PAVLOAD 20,085 ~g 144,::80 I Sf PERCHH SAVING') BLOCK FUel /
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U~ (I lUOO ZOlli! 3(100 4000 500n Ikn,l o (01 , i'lO(1) (200m (30001 Hi :111 H!A( d~~ FIGURE 11 EFFECT Of ADVANI:Ell IlIGH LIFf SYSTEM ON FUEL BURNED
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."reser~ti:0 In !'lqure 12. ,'\5 thE lun,;e .! l:l.·rc..a!:.:d, fi:ore fuel I'~~; bllrned 1UrlnQ cruise. SInce the ~lzec1 COnh'lllrCltl:.ns ha,-1 nlf:lliar "'1r.'1 arE'3S, the ~ue1 currecl .1urin<1 cruiSE: IS not '3H]nifl' ·l.~·l" ciifferent. ':'I:~ lemainjnq fuel burnell '·Tar. il small r::ercenta(j(' ,~f' I'r·e total, sucp that t:he hlr:,act of the oercentage .:hfferences ,'iu€: to ti'!! hiqr. 11ft 3~lster. \In;- re'-:nced. l\t the fT.: l1er ran'l€s, tt:e rue 1 InA'" ,'llr!:1q taf,p-f: anrl approac1, \7aS c larf1f:l rortkn ,.f t:l:e Fuel bur:-.e,~, £loci rLt" dlff~rerces due t:c. the hi~h , 1ft syr:tCI!'
'''P.r~ll'"~ ',n, ::n:1nif i<.:' It, t.; tl'(? ~rp"71~11:~ ,'iqure', 1 n, lic..l teel.
ri'he takeoff tl':U lenqth perf(Jr::lance comparis.)n, :;hmTn In riqurc 11, at ma:<ir.mrr: tu~p.off '1rn~s "eight 1 In,-l!cates a 19-percE:nt reductlCm In fi('ld lenqth jue to tr.f! 21,!'lanceri .~onfiguratlon. ,(ncreased clPerational fley.lbl] 1 ty coole'l there=ore be ohtained. For a qiven field length, longer range r-l13SlI'lnS Ot Increa:O:E'd pa~11(".ad f('tr a aiven range caul"; he achieveo Ni th the ddvan~d .'nnflCTllratlon.
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CLIMB TAKEOFF TAXI OUT , I I I I , 00 1000 2000 3000 4000 5000 I kml I I t I fO) 11000) f2000) (3000) (N Mil RANGE FIGURE 12. ADVANCED HIGH LIFT SYSTEM FUEL·BURNED COMPONENTS /~
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lOCO 1000 2000 3000 4000 5000 (km) ,01 110001 (20001 (3000) INMII RANGE FIGURE 13. CONVENTIONAL AND ADVANCED HIGH LIFT SYSTEM TAKEOFF FIELD LENGTH
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ORJGirJi~~ r·:'.~~;! ,~~ OF POOR Qu.:.\.!...rrt' A summary of the 10\'1 speed performance results of the trade st.udy for the mission require'iJ2nt~. sp::dfied, is ShCA-nl in Fiqure 14. The improved CL r-tAh is due to the airfoil shaFe, VCK, oro-segment flap, and flaFeron. The lA-percent improvement in landing LID would result in a reduction in a!Jproach nOlse. Po. more uniform sranload distribution due to the flaFeron, ann the improved nrag characteristics of the trailing edge flap system ~lere the sources of the increased LID. The takeoff field length ~JaS redu~d clue to increased CLr-lAX a t takeoff flap settings and the improved LID characteristics.
ATiO -12 ASPEC~~ICAL WING SUPERC 23% '20°0 .s"'"
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Dllrinq the OC-y'-20n stwiles, full span slat and va:: confiqurations as tolel]
~s an inboard fixed carrber krueger (FCK) with "CK outboard, and slat inhoarn ann VCI~ outboard pere e',aluated. '!he largest estiMated CLMAXtolaS achieved \"lth the full span 'lCI(. Barly Height evaluations inc1icated a ueiqht ) advantage for the VCR ca:tr:ared to the slat, although as the desicm studies
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Ci= POOH Q'JA~rrl progress, the magnitude of this difference was reduced. A two-percent improvement in DOC was being shown for the final full span VCR configurations. '!be VCR was chosen as the primary leading edge device based on these results and previous experimental two-dimensional wind tunnel programs for both supercriUcal and conventional airfoil geanetries. "j,llese tests had indicated that the VCR had the potential of superior C~x, and also drag levels approaching that of the slat. Figure 15 shows the VCR configuration of the Douglas-funded mechanical system studies for the retracted and extended positions. This contour-changing surface, was supported by the front spar and stowed in the 10\'ler leading edge of the \1ing, formed the lower surface of the leading edge prior to deployment. It extended fran near the fuselage side to the wing tip, interrupted only in the area near the pylon. When the VCR was deployed, its contour was modified by a Irechanical linkage during the actuation sequence so that the desired shape was attained when the surface was fully extended. Various rrethods of VCR drive systems were evaluated and included: open torque tube, closed loop torque tube, cable drive, and hydraulic actuators. For the
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torque tube systems, both jackscrew and rotary gear box drive systems were evaluated. The final rrechanical system selected \vas a rotary gear box drive for the closea loop torque tube system. The torque tube is driven by two SPACE FOR SYSTEMS RETRACTED SPACE AVAILABLE FOR DE-ICING a= FRONT SPAR SUPPORT STRUCTURE ~...!
EXTENDED ~r::; ROTARY ACTUATOR •• J CoIC.IO 10,1.,.
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FIGURE 15. VARIABLE CAMBER KRUEGER LEADING EDGE
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hydraulic motors each driven ~)y Cl sep:lrate hydraulic system. f3eSlcES t-einq lighter ann more canpact, thi s system requires no asymmetry system, will
operaten under a single failure, and syncronizes the vel( by the action of
the torque tube. The full sp:m slat was chosen for the secondary leading edge configuration for the high lift wind tunnel 1TlOCla1 due to the redu~CI rlraq levels associated with Multi-positioning capability, which already has been nemonstraten on current transport aircraft.
Trailing edge system studies included the effect of circular-arc-motion vane-flap, track-motion two-~grJent flaps \lith very large chord extension, and finally, ~ouble four-har linl<age systems for a tHo-segment flap \'iith reduced chord ratio.
PreVlous e:cp;nmantal data (botn two- and three-dimensional) had shmm lncreascd perfornance benefit:s for large chord vane-flap and b1O-segment fla~) conflqurations \lren compare.- h"1 the cl[culc:r-arc-motion flap. An ;:;.ltemate tllo-segment flap ccnf.lCJucntlcn was a Iso eva luated in the Douglas
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funded system studies. ThlS traill.1g edgE: high 11ft system had redIcetl chord and aft extension, The orig1na1 two-segment flap was a 35-percent chord ratlo two-segme,lt ,;ystel'l \'lith (l chord extension of 25-per~nt (over tre outboard spdn). Tre olt- ,"na~ two-segIrent flap had a 3D-percent chord ratio and a IS-percent chord exto;"!nSlOn for the same reglon of d~ wing.
ThiG flap geonetry and motlon was compatible "/1 th a linkage support system \'1hich offered significant advantages in cnuse drag (due to smaller support facings), in structural weight, and in llaintainability when compared to the large chord lugh-extenslOn track system. Trese studies lndicated the irnp;lct of. the linkage system \'.'as to burn 1.7-percent less fuel and have a reduced duect nperatlng cost of 1.1 pt::rr.ent compared \-lith the track 5'.lstem. '1:'re operating empty uClght \laS also reduced. All of these factors ... ,ere uetermined dt ldentical payload range, takeoff field length, cruise l1ach number, and awroach speed.
Basad on thes,: studles, the redIced flap chord linkage motion was selected "', for tle pnmar'l traihng edge system. Figure 16 shows the trailing edge flap system se lected f rom the design studies. '1:'re flap system consists of double-slotted blo-segment type flaps for the inboard and outboard flaps.
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ORiG:l'~AL r-'.'l~~ IS OF POOR QUALITY ROTARY ACTUATOR VARIABLE CAMBER KRUEGER \VCKI DOUBLE FOUR BAR LINKAGE DRIVE TORQUE TUBE rSPOILER [
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'~u)IIUJl1A FIGURE 16. HIGH LIFT SYSTEM Bet\oleen these \\'as a flap:ron (single-slot flap) which translated in too same manner as the maln flap of the inboard and outboard flap system. The inboard and outboard tralling edge flaps are mounted on double four-bar linkages which are actuated by redundant torque tubes (closed loop) extending from the center of the aircraft. Dual hydrualic motors, powered by separate hydraulic systems, drive tie torque tubes, and one system alone
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can operate the flaps at a recllccd rate. Tie advantages of this system were similar to those of too VO< system. A linkage supported single-slot flap system was selected for too secondary-trail1ng-edge high lift configuration.
TOO experimental evaluation of this system could be accomplished with very little additional expense by retracting the aft flap of the t\'lo-segment configuratlOn.
Final Configuration Studies The Douglas Eunoed system studies, which resulted in the configuration selected for the wind tunnel program, are presented in this section.
t'a turally, the high speed requirements determined the primary design variables for the cruise winq. Since the cruise wing geometry was the starting point for the high lift system design, and directly imp:lcted the low Sf€ed characteristics, the selection of the more basic parameters are
reviewed (see also Reference 1). Also presented is tre wing sizing for the
configuration with the selected high lift system and comparisons of estimated performance of the current configuration and other existing
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transports to illustrate too gains obtained by use of tre application of the
odvanced technologies. P-.s stated previously, the basic configuration was a
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medium range wide-body transport capable of replacing the previous generation of narrow-body aircraft. The design goals for this aircraft relative to today's transports were improved economics, 10\'ler fuel consumption, reduced noise and explnded cargo cap;lbility. ~sign emphasis \las placed on employing advanced technologies that improved aircraft ~ operational economics. The principal advanced technology incorporated in this design was a hiqh a3pect ratio supercritical wing; other advanced technologies inCluded an advanced high lift system consistinq of VCJ< and tHo-segrcent flaps, longitudinal stability augmentation, use of composite structure for selected components, a short core-cowl nacelle for the \"inq mounted engires, and significant advances in digital electronic systems.
The DC-X-20n nodel 0969n-21 was designed for one-stop transcontinental range capability \-lith inherent gro\-lth potential for nonstop transcontinental flight, or a significant increase in car:acit'l. '!he required confiquration perfornance included: Imler fuel consumption per seat-Mile thar, any current turbofan pmlered transport, excellent hotihigh ai rfield performan~, noise
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levels substantially belO\o1 expected rraneatory levels (with adequatp margins for airplane growth), and excellent seat-mile and airplane-mlle oIX'ratinq E'conanics. A three vie\! of the resulting configuratlon, Hhich used DC-In fuselage components, is shat! I ! n figure 17.
Early in the system studies, cruise r\:lch numbers of 0.711 (H = .7P.) and ~, =
0.80 were chosen, the former to favor fu~l burned, and the latter for minimum direct operating cost. Figure 18 illustrates the variation of several of the important operational parameters wit:h cruise r1ach number.
~'lhile the me values are mlnu:lized in the re,)ion of rt=O.79 to 0.81, the
hlock fuel hurneo \'laS miniJ1ll!Ill at sane r'ach nurnher lesE' than 0.7B. These bTO consideratlons were of paramount concern in the selection of cruise ~bch number. Other operatlonal factors \-lere also taken into a-:collnt but to a lesser extent.
'l'he tasle Hinq geometry \'laS selected after studYlng the effects of "linq Clrea, thickress to chord ratio, s\'1eeP, and aspect ratio on direct operating r.ust, fuel hllrned, and other economic indicators. Fiqure 19 illustrate!': results of a varlatinn of winq sweep on DOC. With an average
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thickness-to-chord ratio of n .12" a miniI'lum value of roc \-laS achieved at
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VARIABLE CAMBER KRUEGERS (VCKI TWOSEGMENT LINKAGE FLAPS SPOILERS 15-52 m (50 92 FTI
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TWO TURBOFAN ENGINES GENERAL ELECTRIC MODEL CF6-45 / ~'" RATED 45.000 LB SLST J ./ '-3, ALL-NEW ADVANCED...,./' , ....
TECHNOLOGY WING ' ,\
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4277 m (14033 FTI FIGURE 17_ DC-X-200 GENERAL ARRANGEMENT
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FIGURE 18. FACTORS WEIGHED IN DETERMINING OPTIMUM CRUISE SPEEU tIOlr--------------~------------_;r_------------_r---------- j: z ,~ w CR 080 u a: w OAC BASIC RULES !!; CRUISE CLIMB MISSION :2 WING tic ADJUSTEO FOR z CONSTANT M 0 olv
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FIGURE 19. EFFECT OF WING SWEEP ON DIRECT OPERATING COST
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UP.iG:, : ..... ~ OF FOO.:: C- .• _ .. ': 10-degrees sweep. Aspect ratio effects on economic parameters and fuel burned are presented in Fiqures 20 and 21. '!he higher aspect ratios show significant improvements. Considering the results of these studies, their impact on the nOIse characteristics of the aircraft in low-speed operation, and the risk factors associated with higher as~ct ratio, an aspect ra tio of 10.0 based on adjusted wing area was selected.
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-15 -2(1 -25 100 105 11 0 11 5 8,0 90 95 ADJUSTED WING ASPECT RATIO FIGURE 20. EFFECT OF ASPECT RATIO ON ECONOMIC PARAMETERS +2 MeR - 0.80 ... 0 z w u a:: ~ -2 w > ~ <t ..J -4 IJJ .:c BLOCK FUEL AT 1050 km 1750 N Mil .().
-6 ,--- Iii I • 85 90 95 100 105 110 115 ADJUSTED WING ASPECT RATIO
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FIGURE 21. EFFECT OF ASPECT RATIO ON BLOCK FUEL
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U;~~G:·:\·:'" ~ OF PCU,'\ '''::' -,-,,-, , OJ' t'1ing area study results are shown in Figure 22. Because of terminal compatibility, it was Cesirable that the \-ling SIBn for the high aSfect ratio \ling not exceed that of the DC-10-30. 'Ibis requirerrent, together with fuel usage, DOC, an initial cruise altituCe capability of 10,400 m (34,000 ft),
an approach sp:ed no greater than 182 ktn/h (130 k+) with 30 degrees of flap
.I Ceflection, resulted in a wing area of slightly under 220 M2 (2400 sq ft).
\'lith the S\'JeeP, thickness, aspect ratio, wing area, and high lift system defined, the resulting comparisons of aircraft performance, geometry, ann other aircraft cP3racteristics, are shown in Figures 23 through 33. Figure 23 is a performance summary illustrating payload range, takeoff field lenqth, and altitu& capability. Takeoff field length capabilities of the lX'-X-200 show the effects of this configuration and high lift system. At 0 0 rraximum gross ,'reights, the field length at sea level for a 29 C (84 F) day is only 2173 n (7130 ft).
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The taKeoff field performance at hot, high altitude airports is typified ~J the takeoff \oleight cap:lbility of 127,006 kg (280,000 lb) out of Denver on a 33 r (92 F) day. This would be sufficient to transport a full passenger load with more than 4,536 kg (10,000 lb) of cargo fran Denver to major East ) coast cities.
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Because of the emphasis on econanics and fuel efficiency, the sUJ:ercritical airfoil technology was applied to the DC-X-200 to increase .,ling thickness while still achieving some benefits in buffet boundary. The latter js translated into cruise lift. Figure 24 illustrates these trends. The resulting wing thickness provides a structural weight advantage and the airfoil shape results in an increase in takeoff and landing CLMAX. Part of the weight reduction has reen utilized to increase the wing aspect ratio in order to improve aerodynamic efficiency. Figure 25 shows a nine-perCent improvezrent in the cruise ratio LID \-Ihen ca11Iared to the OC-lO-IO.
For the 10\-1 ~d characteristics, the \-ling and high lift system provide a high landinq CL C.apabIlity \·lith a maximum flap deflection of only 10 MAX degrees. The small required flap deflection at landing results in a
significant improver.ent (70 percent) in the approoch LID over the canIarahle
values for the DC-lO-lO (see Figure 26). The takeoff LID is improved 30 percent and tt.e maximum lift coefficient i'-j int::eased by 17 r-ercent.
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Felated noise cl'.aracteristics are shown b Filj'!res 27 through 29. Figure 27 Indicates that the estimated OC-X-200 leveL., resultinq frCl'l the hiqh aspect ratio 5upercritical \-lin'1, advanced high :!. ~ft systems, high-bypass ratio 02ir---------------------------, () 7 ir----------------------.
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:I: III ~ :; Cl a:: Z (J ~ z w !2 Cl III ct W o a:: w > ct o I II (( ('« (II 03· I "C(II/llI CONVENTIONAL OC·X 200 CONVENTIONAL DC X·200 FIGURE 24. COMPARISON OF THICKNESS AND CRUISE LIFT COEFFICIENTS - CONVENTIONAL TRAf'JSPORT AND DC·X·200
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FIGURE 25. IMPROVEMENT IN CRUISE lIFT-TO·DRAG RATIO FOR ADVANCED HIGH ASPECT RATIO WING
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..I :;; :l ~ X c( :;; DC·l0·1O DC·X·200 DC·l0·l0 DC X·200 TAKEOFF LANDING MAXIMUM TAKEOFF WEIGHT MAXIMUM FLAP DEFLECTION WITH TYPICAL FLAP FIGURE 26. IMPROVEMENTS IN LOW-SPEED PERFORMANCE FOR ADVANCED TECHNOLOGY
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40 60 80 100 200 400 600 800 1000 TAKEOFF GROSS WEIGHT (1000 LSI FIGURE 27_ FAR PART 36; ICAO ANNEX 16, CHAPTER 3 (NEW DESIGNS) NOISE LEVEL REQUIRE- MENTS AND ESTIMATED DC-X-200 VALUES
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. , r _I' engine, contain significant margins with even the most stringent noise requirerrents. Ore attribute of the low noise values for the DC-X-200 is presented in Figure 28. Estimated 100 EPNdB noise contours produced by the nC-X-200 during approach and takeoff are compared in Figure 2R with those of contemporary commercial aircraft for similar operational conditions (1000 mile trip/lOa-percent plssenger loading and the rew takeoff procedure of the Air Transport Association). The land area ~ncompassed by the 100 EPNdB contour for the DC-X-200 is one-fifth or less than the area for the older
----
generation of transport aircraft. 'lhus, the areas nm., expose(1 to annoying aircraft generated noise could be drastically reduced. '!his could permit operation out of noise-critical airports, increasing the utilization and route structure for the airlines. Thi3 can be a Significant factor in today's "deregulation" of airline operations. Figure 29 compares the flyover nOlse levels of existing aircraft \-lith the estinated values for the DC-X-200 confiquration. Very low relative values are obtained for sideline and taKeotf, \'lith 5u~rior values being shown for the approach condition.
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B127·200 AOV (JTSD·151 5
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BRAKE RELEASE FIGURE 28 ESTIMATED 100·EPNdB APPROACH AND TAKEOFF NOISE CONTOURS (1QOO-STATUTE·MILE TRIP)
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DC 10 10 DC 10 10 DC X 200 B131200 lllUI
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~DCX200 AlOOB4~ j 90~ ~ -l 8O~ SIDELINE TAKEOFF APPROACH CUTBACK FIGURE 29. FAR PART 36 FL YOVER NOISE LEVELS Tn order to e'JaluZlte th.- Jm[Klrt ()f \'h2 SUi' .1f the advZlPced technolot1ief; on
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fuel burned and roc, a comparison of the OC-~~-200 ann e;{isting o~rationa)
aircraft. was ~rfonred. The canp:!rahl(~ characteristlcs are shollTn in Fiqure /~
30. The results are sho.offi in Fiqure 1) \·mere a canmrison i::: ram \-lith the
/ DC-IO-IO as a rese, of the relntive fuel burned Fer seat-mile as a function of fuel burned per mile. . lative direct o~ratinq cost per 1050 km (7,0 n.mi) trip ann Fer seat are shChln in Figure 32. The estimaten values for the 727-200 aircraft are the base for this cOMperison.
'!'he OC-X-20n configuration \.;as also cCl'lpared to a confi~uration wi th a larqe \-linq area, an aspect ratio closer to the first generation ;et transport aircraft, and a hiqh lift systen consisting of a slat and sinqle-slot flap.
rri1e performance companGon \Jas r:erfomed at 1050 km (750 n.mi). 'IlIis range is morE' typical of th..? averaqe distance over 'olhlCh th~s class of aircraft Houlel operate. Re.3Ults of this stud" are sh01l.n in Figure 33. Significant improvement in fuel burned and DOC are indicated for the configuration optiMized for the dc,mestic op:!rations.
Since the overall confiquration contains other advancen technoloqy features, Figure 34 presents the breakdown of the imract of these features on 00(' and
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prof~tability. The supercritlcal "ling nncl hiqh lift features are a siqn~ficant r~rtion of the total.
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B727·2oo AOV OC·X 200 IN·211 A300B-4 DC 10·10 ENGINES NUMBER AND TYPE 3,JT80 IS 2, CF6-45 2.CF6 SOC 3 CF660 200 17 145.0001 17837140.1001 THRUST PER ENGINE. KN ILBI 68 95 115.5001 226.86151.0001 NUMBER OF MIXED CLASS SEATS 140 230 257 295 CARGO VOLUME. M3 ICU FTI 323111.1411 94 97 13.3541 105513.7251 130 77 14 6181 WING AREA. M2 ISO FTI 149611.6101 2202 12.3701 245 7 12.6451 343313.6951 MAXIMUM TAKEOFF WEIGHT. KG IL81 86.4091190.5001 132.9021293.0001 ISO 002 1330.7001 1950441430 0001 OPERATOR'S EMPTY WEIGHT. KG IL81 47.368 1104 4301 79.038 1174.2501 90.945 1200 5001 110,236 1243.0301 DESIGN RANGE. KM IN Mil 2.76911.9801 3.63612.6001 2.769 11.980) 4.349.J.ll01 CRUISE MACH NUMBER 081 080 OBO 083 TAKEOFF FIELD LENGTH. MTOGW.
o 2 630 18.6301 2.17117.1301 2103169001 2.82519,2701 SL. 29°C 184 FI. M IFTI APPROACH SPEED WITH FULL PSGR.
BAGGAGE AND RESERVES KM/H EQUIV IKEASI 18211301 1831.311 18J 11311 18211301 PAYLOAD FROM DEN-JFK JJoC 192 FI DAY.
AIRLINE RULES (PSGR/KG ILBI CARGO I 230IJ.946 18.7001 169/0 24610 FUEL 8URNED AT 1389 KM 1750 N Mil. KG ILBI 7.902117.4201 8160 117.9901 11,122 124.5201 123741212801 RELATIVE AIRCRAFT STUDY PRICE IPERCENTI 38 79 85 100' '1971 SlO 4M FIGURE 30. COMPARATIVE CHARACTERISTICS
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STAGE LENGTH 1050 km 1750 N Mil +40 B727·200 ADV ;: z
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.25' PEA SEAl --../ 1977 DOLLARS 16.5¢/kgI50tlGALI FUEL '2801 $,Tlti., FIGURE 32. RELATIVE DIRECT OPERATING COSTS FOR CONVENTIONAL AND ADVANCED CONFIGURATIONS OPTIMIZED FOR COMPROMISE FOR DOMESTIC INTERNATIONAL AND OESIGN 06JECTlVF OPERATIONS DOMESTIC OPERATIONS 1'11 ... 13 GEOMETR',' AilE;. ~r1AlI LARGE: ASPECT RATIO HIGH LOW MEDIUM ~PArl MEDIUM HIGH 11FT SY;TEtI.I VCK PLUS SLAT PLUS TWO·SLOT FLAPS ONE SLOT FLAPS REI.-.lllfE: r-UEI BURNED 0 60 PERCENT RELA TillE DIRECT OPERATING 0 26 PERCENT 9DP81~~ COSTS FIGURE 33 WING GEOMETRY CHOICE
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-. !: INCREMENTAL- OPERATIONAL PROFIT DOC (%, ('J(,' ADVANCED FEATURE .58 -35 SUPERCRITICAL WING (CONSTANT AR' .,0 -2.0 HIGH ASPECT RATIO WING -19 +32 VARIABLE CAMBER KRUEGER +5 -05 LONGITUDINAL STABILITY AUGMENTATION -02 +1 COMPOSITE FLOOR BEAMS AND STRUTS -02 +2 COMPOSITE CONTROL SURFACES/FAIRINGS/WING FIXED TRAILING !:DGE -02 +2 AUTOMATIC REVERSE THRUST -01 +2 ELECTRICALLY SIGNALED SPOILERS -07 ·>10 SHORTENED ENGINE CORE COWL (NO PRIMARY REV' -03 .6 DIGITAL FLIGHT GUIDANCE AND CONTROL -96 +130 TOTALS ·R~l AlIVE lU Al.rERNATIVE INVE~lMENT~ AT 8 ~ •.• ANNUAL lNTEIH.~T FIGURE 34. ADVANCED FEATURES ECONOMIC ASSESSMENT
J
Application of supercritical \·Jing technology for the nedium ranqe transport has shOl·m significant potentlal benefits fran both economic, fuel usage and social aspects (no1se). Increased confidence in these predicted irnproverrents is aired by a substantiation of the performance level for each of the disciplines which are involved in the configuration refinit10n. The current EE'J' studies of the supercritical wing high- and lO\Ol-speerl ~erodyna~ic characteristics, have aided in this substantiation, and also ~ighlighted areas for further development.
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HIGH LIFT AEmDYN.~:C IESIGN PIDCESS General Design Constraints The aerodynamic design pr~ss for the high lift system is quite canplex and involves the interaction of reany disciplines. A primary design constraint is, of course, the requirement of high lift system can};atibility \-lith the cruise wing shaFe and structure. 'Ibis canPltibility is influenced directly by the spanwise variation of the available chord ratio for the high lift system canponents, as well as the shape of specific portions of these high lift devices.
Figure 35 illustrates the regions determined by the cruise configuration and those surfaces which are at the discretion of the high lift designer.
Further, these latter surfaces are also influenced by structural considerations. For example, in Figure 35 the slat high lift device and its associated wing under slat surface (~mSS) are influenced by local chord
"J
constraints and slat trailinq edge thickness and closure angle. l\ W'K concept allows more aerodynamic freenom for the basic shape definition resulting in increased 11ft cap:ibility. Other factors \-1hich can influence the leading edge device <.E:rorl'lnamic definition include front sr:er location, anti-icing ducting requirements, actuation system, fuel, and hydraulic line space requirements.
LEADING EDGE DEVICES TRAILING EDGE FLAPS TWO.SEGMENT FLAP
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I , ---- DEFINED BY CRUISE CONFIGURATION ------- DEFINED BY HIGH LIFT CONFIGURATION
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,opel FIGURE 35. SURFACES DEFINED BY CRUISE AND HIGH LIFT CONSIDERATIONS
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r..;i{ 2 OF F':;C The trailing edge high lift system is influenced by rear spar location, the spoiler actuation system, the spoiler trailing edge chordwise position, and the spoiler closure angle and thickness. Relaxed static stability adversely affects the inboard flap chord, due to the further aft placement of the landing gear and related structure. In addition, due to lateral stability and control requirements, spoiler chord and sFan as well as aileron location, influence the trailing edge flap definition.
High Lift ~sign Hethods t'lith the high lift concepts selected and the guidelines for the related structural constraints defined by the results of in-house studies, the detailed aerodynamic design process indicated in Figure 36 was used to cef in~ the high lift. systems for the wind tunnel model. '!be design functlOn consists of four basic parts: the experimental data base, two-dimensional
,J
(2-D) analytical studies, blo-dimensional to three-dirrensional (2-D to 3-D) concepts and the three-dimensional lifting surface calculations.
3-D LIFTING SURFACE THEORY EXPERIMENTAL DATA BASE 3D HIGH LIFT WIND TUNNEL MODEL DEFINITION ..
2-D ANALYTICAL PREDICTIONS FOR 3-0 GEOMETRY \
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FIGURE 36. HIGH LIFT AERODYNAMIC DESIGN PROCESS ~ ~ ."
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ExP'!rirrental Data Rase An extensive experimental t\'1o-dimensional data base for aerodynamic characteristics of supercritical airfoil high lift systems had been generated by previous Douglas funded studies. This data base included: pffects of conventional and supercritical airfoil sections: thickness ratio: slat and VCR leading edge devices, single-slot, two-segment, and triple-slotted trailing edge devices, and optimization of the position and dr-flection of these systems. Detailed wake and boundary layer measurements for selected configurations han also been generated to prc1vide a more fun~ntal understanding of supercrltical airfoil high lift systems.
This experimental cmta, on conventional and supeccritIcal high 11ft configurations, \-laS used in the design of the high 1 ift system for the current three-dinenr;icnal high lift model. Variuul;; Krueqer leading enge shapes \'1ere evaluated ... /1th resppc:t t(. cl :: and profile n(:'fI mfl characteristICS. 'I\m-dimensllJllal values of slJch parar:let:er~ df, flap IJ.fl C1
) locrement, (Ac =0)' flap maXlmLml lift increnent, (.1 r.lc t 1 j, dnd leadinq
1X a a~ edge cievlce na:dmum lift increment ~Ct max I..I:.. \,rere also utilIzed 1 r. the genp.ration of the est imateel r.haractenstics for the \llOd tunnel r..cdel. In andition to thE'- evalllC'll"ion of 'JnflOUS high lift. ("oncepts, anot.her significant use of the ).-i") , cperimental data In the ~siqn preress \oTaS th~ ~termination of key desiqn crIteria by means of theoretic-al .~al culatl onS for the e~perimentally optimized configurations. The defiiqn criteria lncluded suct1'-~n ~aks, pressure gradients, and calculated separatlon locations and were used for the oo!ngn and element posItion studies on t~ equlvalent three-dimensional model com};Xments.
'T'he t1ul tlelp.f.lent Alrf~ll t Des lqn and AnalysIs Method (rvtADMH) theoretical method (P£:ference 4) waG lIsed to generate the theoretical calculations.
'l'hiG computer program can compute the high lift characteristics of multielement airfoils of arbitrary shape for a V1SCOUS incompressible fluid in the absence of flow separatIon. The program combines a geometry definition routIne, a potential flow solution based on a surface source distribution, and a finite-difference boundary layer routire to accomplish the analysls. The geometry routine can smooth and space the body coordinates for input to the potential flCM program. After the potential ) flCM distubutl.on has been calculated tie boundary layer characteristics are
---
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Crt:C~ ,)F PC'~: -: (. _r,_,. 'i calculated and an equivalent inviscid body is fooned by the addition of ~le boundary layer displacement thickness to the original geoemtry. These coorcinates are used to calculate a new pressure distribution for the ecr~ivalent body. Figure 37 illustrates the computational process.
r---- -I'NPuT-- --_:--, GEOMETRY DOUGLAS DEFINITION NEUMANN PROGRAM POTENTIAL COORDINATE FLOW SPACING AND PROGRAM SMOOTHING POTENTIAL FLOW PRESSURE DISTRIBUTION BOUNDARY LAYER DISPLACEMENT
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THICKNESS POSSIBLE ADDITION RED EF INITION ITERATIVE TO BASIC OF GEOMETRY GEOMETRY LOOP RECOMPUTATION OF VISCOUS SOLUTION r
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L_~-=--=-=- _______ J
FIGURE 37. FLOW OIAGRAM OF COMPUTER PROGRAM FOR MULTI-ELEMENT AIRFOIL DESIGN AND ANAL VSIS METHOD (MADAAM) A comparison of a representative calculated pressure distribution from the t-iADAAH program and the multielement experimental data for a thick supercritical high lift system is presented in Figure 38. Theoretical predictions included both tre potential and viscous flow capabil ties of the
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program.
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,\ ----- EXPERIMENTAL DATA C p - - -- - --THEORETICAL SOLUTION (MADAAM PROGRAM) /' /'
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·10 10 20 30 40 50 60 70 80 90 100 110 120 130 PERCENT UNBROKEN CHORD oP"'"
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FIGURE 38. COMPARISON BETWEEN EXPERIMENTAL AND ANALYTICAL PRESSURE DISTRIBUTIONS FOR MULTI·ELEMENT AIRFOIL 2-D AnalYtical Predicti~for 3-D Georretry For the ~Elected high lift concepts, values of such basic parameters as airfoil tldckre5s and carnb~r, flap chord, and SJ.=X)i1er trailing edge location varied from the existing experimental data base. Moreover, significant variations in these basic parameter3 occurred over the span,of the three-diIrensional wing. In order to account for these effects, theoretical analysis of various sparn'lire stations of the wing were carried out by rreans of the HADAN1 program. Extensive use of this program...-as required to define the flap leading edge shape and positions relative to adjacent elements.
This was due to tre relatively thin trailing edge sta~ of tre outboard supercri tical \'ling, and the basically uncambered trailing edge region and short flap chord near the wing root. Figure 19 presents typical calculations for three spanwise positions. '!he configurations shown include
)
a va< and two-segment flap system for the inboard and outboard spam-tise station, and a VCR with single-slot flap for the midspan wing position.
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c, t:: -:& IN'OARO~ATION
c, E"~·~ ,1 MIDSPAN STATION
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~-r- r-==-- r~ OUTBOARD STATION
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FIGURE 39. MADAAM PRESSURE DISTRIBUTIONS FOR THREE SPAN STATIONS 2-D to 3-D Concepts The calculated 2-D results were then modified by 2-D to 3-D sweep effects.
'l'hese concepts \.,rere based on previous in-house correIa tions of 3-D experimental multi-element pressure distributions and ~moAAM prograc calculations. The correlations acconunodate the basic s\.,reep and taper effects of 3-D \'ling planforms with multi-element airfoil sections. The corresponding theoretical pressure criteria (peak pressures and gradients)
pstablished fran the 2-D experirrental oota base were also transforrred in the
sarre manner, such that canr:arisons of pressure criteria and calculated 3-D
pressures for the wind tunnel model could be maoo. All of these pressure
criteria were evaluated as a function of the three-dimensional sectional lift coefficient (C13-D).
3-D Lifting Surface Theoey For a given airplane lift coefficient (eL) the section lift coefficients
)
\OJere evaluated by rreans of the Giesing Vortex Iattioe program (Peference 5).
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This proqram has been previously used to calculate c1 - for high lift D confiqurations and good agreement between experiment and program calculations ''las obtained (see Reference 6). With the c1 _ evaluated, the 3 D previously rrentioned 2-D to 3-D pressure criteria could be evaluated for a given value of CL.
t'1hen evaluating leading edge device, wing leading edge, or !-russ p;!ak suction values (i.e., to determine prop;!r VCR or slat deflection or position), one additional correction was used. The 2-D experirrental data mse indicated that near cl a significant reduction in flap lift was apparent. If the rrax geometric flap deflection was used in the Giesing program near CLMAX, the c - distribution across the span would not be the prop;!r variation.
f3 D E'urther, for a given 2-D section hft coefficient the r·lI\D.ZW1 flap lift would be overpredicted, and a sur:presslOn e.f too suction peaks at the nose would resul t. Accordingly, the flap deflection input into the Giesing and HADMr-t programs ",as reduced by an amount determined by the 2-D data base when ,) evaluating suction peaks near stall for the leading edge regions preVl.ously discussed. A typical SP:U1"lise variation of calculated slat and i\7tJSS minimum pressures is shown in Figure 40.
SLAT PEAK VALUES ACA TARGET I WUSS PEAK I VALUES I C I ACA TARGET PM1N I I I I I \ I OUTBOARD SLAT S = 35"
OUTBOARD --t- AILERON
I· I· ·1 FLAP I
o 01 03 04 05 06 07 10 0.8 0.9 '!'l/joI"'''IA SEMlsPAN STATION FIGURE 40. THEORETICAL SPANWISE VARIATION OF SLAT AND WUSS MINIMUM PRESSURES FOR THE LOW-SPEED WIND TUNNEL MODEL (LANDING FLAPS C = 3.41 L
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Also shown are the model target values. At each leading-edge device deflection,plots such as trese were used to evaluate tre I::asic position (gap and overhang) of the VCK or slat. Once a basic position was established, the results of the position study as calculated by the method of Figure 37, and the 2-D experimental position trends, were used to determine the alternate grid positions.
Sbape Definition for Leading and Trailing Elements The basic shape for the VCK was defined by a previously tested two-dimensional configuration. Slat and WUSS shapes were defined by aWlication of the MrillAN1 program with systematic variations in WUSS chord, camber, and nose radius at a nominal slat position and deflection. USing ~ .:;:::::- the procedure of Figure 37, a \iUSS shape was chosen for the defining stations.
The MADAru1 program was also used to defire the two-se900nt flap and flaperon . shap:s at the airfoil defining stations on the wing. Within the structural
,~constraints previousl~ discussed, flap nose radius and camber were varied
and the resulting pressure distributions and separation locations were
'---
evaluated at a nominal flap setting based on previous 2-D and 3-D two-segment and single-slot flap experimental studies. The methodology of Figure 37 was again used to determine the flap shape to be tested. '!he ratio of main flap chord to aft flap chord was determined by previous 2-D .?nd 3-D in-house experimental data.
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OF PGC:i ~:"',!._01 : WIND TUNNEL mIEL IESCRIPl'ION 'Iile configuration selected for the wind tunnel test progra."1'\ incorporated the results of the in-house studies. Additionaly, the model canponents allowed the testing of alternative high lift systems. Figure 41 shows the basic configuration shape. It is characterized by a high aspect ratio sUPercritical wing, wide-body fuselage, and wing-mounted engines. The trapezoidal \-ling planform, which has a quarter chord sweep of 28.5 , incorporated leadinq and trailing edge breaks for imploved cruise characteristics. Thickness to chord ratio varied from 15 percent at the root to 11 percent at the ... linq tip. Near the root, the baSIC sllpercritical airfoil section chanqed to a symmetrical shape for improved crUIse perforrrance.
The \-/ind tunnel maCel \-laS a 4.7-percent scale repres;ntation of the OC-X-~OO airplam configuration. The model \,'<1S desiqned and fabricated for testinq ) at high P-eynolds number conditions at the NASA Ames l2-Foot Pressure {'Tind Tunnel. The model designation \-1aS LB-486A for the Ames test and LB-4fl6C for the HASA langley V/S'IOL test.
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DIMENSIONS IN CENTIMETERS (INCHES) r" •. DEL sr ALE
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t -- t __
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36530 (143821 POD LOCATION
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~(27652)
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FIGURE 41. LB-486A MODEL THREE·V1EW
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Figure 42 illustrates the high lift systems for the wind tunnel tests. The primary leading edge configuration was a VCK flap. While the basic confiquratio~ incorporated independent inboard and outboard segments, additional pieces were fabricated such that a continuous full span VCK e~tent was possible (i.e., sealed at the fuselage side and continuous across the engine pylon). Associated with this VCK were cutouts in the lower surface of the wing (VCR wells) representing the absenoe of a continuous lower surface \-lhen the VCR was deployed.
The secondary configuration for the leading edge was a slat. '!his device incorporated a sealed configuration over the engire pylon, but a gap at the fuselage side. Positioning of the inboard and outboard segments was accanplished inde~ndently. Associated with the slat \-las a revised leading edge contour (l'lUSS).
Por the trailing edge high lift system, the primary configuration was an inboard and outboard two-seg:rent flap. Between trese two flaps was a device
)
called a flaperon, which was essentially a single-slotted flap, that was articulated in the same manner as the main flap for the high lift conditions, but incorporated a high-speed short chord aileron in the retracted, or cruise configuration. At too high lift condition, the aileron was locked in an undeflected position. This permitted a contirn.Ious flap span of sore SO percent, resulting in an improved span loading for high lift conditions.
As Figure 42 shows, the secondary trailing edge configuration was a single-slot flap. On the model this configuration was obtained by simply stowing the aft flap into the main flap. 'lbe resulting configuration was an SO-percent span single-slot trailing edge flap.
In order to optimize the high lift geometry, both the leading and trailing edge devices incoq::orated bracketry for changing deflection and position.
The model also incorpJrated an aileron (left wing parel only), spoilers, and remote-drive horizontal stabilizer deflection capability. TOO wing and high lift syste~s were also instrumented with static pressure orifices at the five stall'Nise stations. Other model canporents included nacelles, pylons,
)
landing gear, and a cruise wing trailing edge (i.e., flap:; retracted).
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~
<:::::.:::::::. -~
--- ~,
--
--L- <..:::..:::.~ CLEAN TAKEOFF AND CLEAN LANDING TAKEOFF LANDING "- SECONDARY CONFIGURATION· SLAT SECONDARY CONFIGURATION SINGLE·SEGMENT FLAP ~ ~ ~I~~~,
({~ .. - L
d --
CLIOAN TAKEOFF LANDING TAKEOFF LANDING ~lli'~l'" FIGURE 42. HIGH LIFT COMPONENTS EVALUATED IN EXPERIMENTAL TEST PROGRAM
J
~ fuselage (B2A) consisted of duplicated OC-lO model nose and aft. fuselage shell sections, a new tOl. r- nter section, and a ne\1 wing-fuselage fillet ::ection.
}\.n existing fuselage core \-las adapted for attachment of the fuselage sooll sections, support of two five-module scanivalve systems, support of a bubble pack plate, and attacrment of the \Hng, vertical, and horizontal stabilizer.
An existing fuselage internal pitch system \-1aS installed in the core. ThlS 0 0 system permitted the fuselage to be pitched fromaFRP = 0 to +10 \<1hile
the internal balance remains at amP = 0 • The aFRP (angle of attack) is the
angle which the Fuselage Reference Plane (FRP) nJakes with the equivalent free airstream. '!be other pitch angles are obtained by uSlng the external pitch system.
The ~13B model wing was fabricated from Armco 17-4 steel and was lofted to simulate t~ auplare wing with a 19 lead. The wing geometry and planform dimensions are shown on the wing diagram (Figure 43). h wing-fuselage
)
fillet notation is X2S.
l~
L--
u
J
- \"J1t..1C.OHIr.I~ , 1/1 9ij4148 O~, lliCw. ~ .
• 04696 m' 1~05!t SO f tl AR 10'010 TArER RUIO • o 140J Xw 36 36' /' IUJUI.
DI .... I ... SIQM.IN CENTIMETERS ,'NCHESI MODE L SCALI IC •• ,IN'C 3J J61 • Y '60)80 '63 'OJ, x"". '1 UOl16 3101
Vw lB 2H11!t on" .w
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I '
1418811181"\ I I 201'0182 .. 1\ \
!
\ \ 00 , \ "TI:tJ
~I IJ
-n i:l
o -
U1 O?; N I1It!i :xl j •• 12eU311 .o\J L" ,.
:t- a
r- rfl
~~ :;-
-<w
, - - INBOAAD SPOOl f AS !
'" I' . Vw
I" l ~ ... BOAAD Two SEGMENT flA' I
"-.- 57 73_
Yw" 112 7291 Htll.
I
x", .. 1.3.1011
',- DUT80AnusrOILfAS • j
I _____ 3On11111231----- \ I-- ------ 430n",00,,- "- O\J'IOAROTWOSEGMENT flAP VIEWINM~G REFERENCE 'LAlliE "- '-- REAR S,AR 'LA~f -1-
"
'" fOAIURON "'-LOWS'£: FIGURE 43. WING DIAGRAM (W ' 3ll
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..
~
...
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OF pc;~,~ (J:- .. _.' i '!he morel wing incor{X)rated t.re follo.~ing features: 1) A cruise leading edge, removable at t.re front spar. This leading edge \Vas tested with and without simulated va< stowage wells. Also proviced was a ~-russ leading edge for tl'e slat configuration.
2) A Va< and slat leading edge device with variable {X)sition and deflection capability.
3) A two-oo91lent trailing edge flap supported at five deflection angles by fbed brackets simulating too airplare flap linkage. Variable {X)sition capability was provided for the main flap.
4) A rnanually set aileron, left side only, and s{X)ilers, both sides.
5) Approximately 400 static pressure orifices installed in the Va<, slat, wing, and flaps.
Figure 44 presents the planform of the horizontal stabilizer as \orell as other qeanetric quantities. The horizontal stabilizer was removable for testlng tail-off. The stabilizer was fabricated ir. ore piece, each side, \o/ithout elevators. An existmg remote control system was adapted to vary
,)
the stabilizer incjd::ll~ bet\'leen +5 and -150.
SH " 0 1298 m 11 397 FT2)
--1-- Y " 231 81 '91 2871
It! "3.80 .\ "0.350 SWEEP C " 30" H
I
27384(10781)
r- 100"
I // PIVOT AXIS 58.85 orr.1ENSIONS IN CENTIMETERS (INCHES) - PERCENT C R MOOEL SCALE 1.
4536 (1.786)
,~
ABOVE FRP
"
~~ a I c/4 MAC
/' YH = 16.11 (6344) ~ X " 14 74 (5.804) / ,/ H ,- ~ ! YH " 15.49 (6.097) /' /// ,/ ::t: u
-T -
~ , 9596 (3 178)
1 +Y
H
L 'THEORET!CAL
- TRAILING EOGE 'MODEL ACTUAL TRAILING EDGE (CUT BACK TO ACHIEVE 0.03 (0.0101 THICk. TRAILING EDGE)
)
FIGURE 44. HORIZONTAL STABILIZER H1A DIAGRAM
~
.. -.
)
The vertical stabilizer (VIA) planform is shown in Figure 45. The stabilizer was fabricated as one piece without rudders and was removable to provide a tail-off configuration.
Existing flow-through nacelles (Figure 46) from a DC-IO moael were used for
tre subject model and designated N2A. '!hey were attached to the wing by new
pylons designated P2A. The pylon plane of symmetry had a 1.8 toe-in relative to the airplane plane of symmetry (measured in the FRP) and is
perpendicular to the FRP with the wing in a rigged position \'lith dihedral =
4.05 • Nacelle strakes (ZlA) were attached to the nacelle for most configurations.
The nose gear simulated the DC-lO nose gear in structure and location. '!he main landing gear simulated the airplane gear configuration with oleos extended. Extended main gear wheel well cavities were not simUlated. A
~
Sv = 0.09850 m (1.0603 FT2) M = 1600 A =035 SWEEP ~= 35· DIMENSIONS IN CENTIMETERS ONCHES) MOOELSCALE C 4"MAC Y = 20.86 (8.212) v
@
39.700 Zv = 1666 (6.561) (15.6301 ;'1" MAC = 26.731 (10.524) V +Zv THEORETICAL TRAILING EOGE
-----
ACTUAL MOOEL TRAILING EDGE (CUT BACK TO ACHIEVE 0.03 (0.0101 THICK TRAIW~G EDGE) VERT --..J \_ 36.759 STABILIZER 114.4721" ORIGIN
)
FIGURE 45. VERTICAL STABILIZER VtA DIAGRAM · ....
:!J
'\ l.J.~.'- ... 1 .....
) OF PU::':~ . ~ _ - - . l DIMENSIONS IN CENTIMETERS (INCHES) MODEL SCALE J-4.30 11.69)
"11--3." PERCENT CHORD
4.78 (1.88) MINIMUM NEW DC·X·200 PYLON (P ) 2A _---I L
~--- -
I
1----- ENGINE CENTERLINE
T AT + 1.6° INCIDENCE I TO THE FRP
f:: ---
- - -_ J''-----
EXISTING DC·tO GE NACELLE (N ) 2A FIGURE 46. NACELLE/PYLON N P 2A DIAGRAM 2A
)
nain landinq gear retracted configuration is also provided. The combined nose and rein landinq gear assemblies nomenclature \OlaS G!p,.
The outboard aileron (a2A) \-laS locaten in the outboard extremity of the ] eft hand wing. The cu.leron incorporated a felt-t'{I:e, \..ell-qap seal and had a different upper and lower nose radius because of the vertical location of the hinge line (Figure 47).
The aileron \-Ias tested at deflection angles of 0 , ±So, ±lOo,±lSo, and ~Oo. The angles \'/ere rreasured in a plane normal to the hinge line. The aileron planforrn is sham in Figure 43.
T\% inboard :,poiler segments {f],f2l and four outboard segments 0 , (f3,f4,fS,f6 hoth sides, could be set at deflection angles of 0 _?o, l , 0, -150, -30 -4So, and -600, rreasured in a plane normal to the spoiler h1nq€ line. The t\-IO inboard segments (fl, f2) \-/ere fabricated as sepuate rnrts wnile the four outboard segments (f3,f4,fS,f6) \o/ere not segmented, but were fabricated as one-piece bent plates. An individual plate was provired for
)
each deflection angle.
• (.)
_J
\ ..... ...J
-)
OR\C:~."': - .
OF P0C;-~ (:- -" FELT SEAL
.. _20 MAX'
I
• ,
L--"'---'-'"
·SEE SUMMARY OF MOVABLE SURFACES (TABLE 11 FOR DEFLECTION ANGLES AVAILABLE FOR TEST FIGURE 47. AILERON SECTION (a ' 2A The scaled airplane spoiler trailing edge thickness would have been approximately 0.0762 em (0.003 in.) The model spoilers deviate from these
~~)
dimensions, with the lower surface of the spoiler segments modified to obtain the standard minimum crailing edge thickness of 0.0127 em (0.005 in.)radius, while maintaining the theoretical planform trailing edge locatlon for flap rigging purposes. The spoiler planform diagram is presented as part of Figure 43.
The definitions of gap, overhang (O.H.), and deflection used to position the leading edge high lift devices are illustrated in Figure 48. Figure 48 also shows the method of attachment to the WUSS and clean leading edge (for the slat and VCK devices respectively). As indicated for the VCK, bolts attached the device to an angle bracket and the bracket was positioned relative to too wing by nl~ans of sp:1cers (to adjust the gap) and pin holes through the bracket into the wing for overhang positioning. An additional set of bolts fastered the bracket to the wing. Each deflsction angle had separate brackets.
A full span wing leading edge slat (LIA,li2A) with a seal at the pylon was tested. Positioning capability included three alternate positions for each of two deflection angles. The slat was fabricated in two parts, with the
)
common bomldary at station Xw = 36.367 em (12.123 in). The inboard and
outboard slat deSignations were LlA and L2A, respectively.
...
~
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--
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~t'-~··:o'" •• O I .... • ' ~ - .~!..., ... _~. 'l OF Fu':'.1 SLAT OVERHANG (-)SHOWN MLL VCK GAP MLL
~
BOLTS FIGURE 48. LEADING EDGE DEVICE GAP, OVERHANG, I\ND DEFLECTION DEFINITIONS A WUSS canporent was also provided to replace the cruise leading edge. A planform diagram of the slat is ShCMIl in Figure 49.
Tl'e deflection angles were measured in a streamwise plane oriented normal to the Wing Reference Plare (vJRP). The variable test positions are defired and ldentified in the Configuration Notations Section.
A full span VCK, with an interruption at the pylon, was desigred to be installed at four alternate positions at each of two deflectlon angles.
The VCR was supported from the cruise leading edge component \'lhich was modified for testing with and without simulated VCR stowage wells. A VCK extension to the fuselage was designated LSA, and the VCK extension over
)
tre pylon was designated LGA.
..
•
--
, .
-
')
... ': ..
Xw = 36.3672 (14.3173) DIMENSIONS IN CENTIMETERS UNCHES) MODEL SCALE I Xw = 110.162 (43.3707)
-----~
Y'
FUSELAGE Ii
.~
~
FIGURE 49. SLAT PLANI=ORM DIAGRAM The VCK deflection angle was defired by the angle between the wing maximum length lire and the va< rcaximum length lire measured in a streamwi se plane normal to the \,lRP. The VCK planform diagram is shCMIl in Figure 50. '!he variable test positions are defined and identified in the Configuration Notations Section.
Definitions for main and aft flap gap, O.H., and deflections are ShCMIl in Figure 51. 'lh! flaperon utilized the same definitions as the main flap.
Figure 52 illustrates the method of attacl1nent of the inboard and outboard flaps to each other and to the wing. As indicated for the inboard flap system, the main and aft flap were bolted to the bracket for the flap deflection required. The gap and O.H. for the main flap were varied by spacers and pin holes through the bracket into the wing surface. The flap bracket was attacred to wing by another set of bolts in eccentric holes (to
)
allow for the O.H. variation). Inboard, each canbiration of main and aft flap deflecticn was obtaired by a l:Eparate bracket.
)
LJ 'l. ...
OF PC": .• '- ....
DIMENSIONS IN C~NmflETERS (lNCHESI MODEL SCALE L3A X = 40.6 35 I W (1601. (13 81
I I
.....-.....
~ x = 11080 W (43624)
I
,)
NOTES V- I VCK TRIM NORMAL TO LEADING EDGE AT X -19 9 (7 851.406 (16 01.11080 (43 674) WITH. VCK . 55' w 2 TRIM AT Xw 351 (1381 IS NOMINALLY AT 18~ CANT fltlllRE 50. VCK PLAN FORM DIAGRAM CLEAN WING MAX LENGTH LINE
--
)
FIGURE 51. FLAP GAP, OVERHANG, AND DEFLECTION DEFINITIONS D .-
, ')
'Ji= 'IUt.,.;.~ INBOARD FLAP If1!STALLATlDN OUTBOARD FLAP INSTALLATION
)
~;€. - ® 1f11t£:rn
~N HOLES TO PROVIDE
FORE AND AFT VARIATION FIGURE 52. TYPICAL INBOARD AND OUTBOARD FLAP INSTALLATIONS The outboard flap positioning was accomplished in essentially the same fashion as the inboard flap system. Various aft flap deflections were obtaired by means of a separate aft bracket for each aft flap deflectior, (relative to the main flap).
The inboard fla[:S consisted of an inbcard two-sel}'Oent flap (FlA,F2A) and an
)
adjacer.~ flaperon single-slot flap (F3A).
~
--- --~
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!- - -- ::-. ..: - - _. - --- -- ..... -:..,..:- -.:--' ---- - -,,.'.. -.-
~;=--"."-"- --"~" --:':--;..-:.~. ~-:-'==-
=---
---"
~
'J
'rhe inboard t\-1O-5egment flap \'las supported by bracket~, sImulating the airplare flap hnkaqe, at mun flap deflectlon angles of 0 , 50, 150 , "~50, -.
~nd J~o. Flap supports also provided four fliP positIons [or each deflectlon angle. £lain flap deflection angles were measured 1n a plane ""r lented stream\'Jise and normal to the WRP. The lr:board main flap \-Ias ~signated Fl]\ .• 'i he aft flap of the Inboard two-segment flap ... /as supported at sIngle 0 ----==:..:...
0 0 , 0 posItions for deflec~ion angles of 0 , 7.5 , 100, 12.5 and 15 • The aft tlap anqles \leLt-; measured in a plane oriented streamwlOO and normal to the HRP with the maIn fldP dL 00 deflectIon. The inboard aft. flap \-Ias CIt,slgnated P2A.
"'hf: t L:lperon £i1 r.'1~ .:lot I lap (F3A) was supported by spllce plates to the WbOOLd flap on t:he wlKlCltlt <::1"1,] and I () tie outboarcl fLip on tl~ outboard .:nd 1.01 <lny 'lIven ,.,)mparaLJle yelli L)oSl tl on. .)fiset spl i.~~1d plateh \Tere provl~rl ,(J p!)sltion the flal-'eron dl tl .. - 25 nomin..1.1 L)O~d ti un \lhlle :"hE. 1 nt,odrd :.1nU ll 011tboatd main tldps \verE: dt the 35 nomInal pOSItion. '('he variable Lest
J
I.NSitlons for the Inboard ttlo-£e("Jllent tlap and too flaper,)ll drc. rif-fined ane identliied In the Confir,ur::ttH'11 t'!otatic,ns Sel:tion.
'L'b:: QUtbCX1(rl avo-5egment fldP \JrlS sUf,p"rtr;:d b', t:xtE~lnul fixed brac;);ets at.
0 0 0 nt.:.tin flap deflection angles of 0 , 5°, J5 , 25 , and 35". The flap supportu L,ro'lid=d fcur flap test positionti tor each 11~ I'lection angle. The support brackets awro:<irrated too auplare tla!., 1 inkaqe. 'I'm outi:xErd f:laln fldp was <.hsigmted F4A.
7he l:tfilll flap deflectinn apCJLt:~ Hem measured 111 a plare onented strearrn.Vl~ und nOll,1Ll. l La the Imp.
c ,
'fhe aft flap \-Ju!3 supp)rted at Hinljl€: pos~Lions foc deflectio!l angles ot n
0 0 7.5 , 100, 12.5 , and 15('. Tl~ aft flap angles \~ere r,ie-:u:ured In a plan:!
uLlented stn:dmUlS(; and normal to the \'lRP \Jlth the m3J n Clap .:tt 0 elefle·'tion. Tre outboarl1 aft flap was designated FSA. The ·Jariable test ~losltllJnS for the ollthourd t ... ,o-segment flap system ar£> defIned and lrient lfli.:d in Ue Configucatlcn Polatl0n SectIon.
)
f,l
~
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____ ~.:. :~;..- -:"::::-;~ __ -3:.-":::"---:.:.~~:"-7--=~ --- :..-=
-- ~-- - - . -_ ... - - - -~ - - .-
-~--
--~~-:: - ------ --- ~~----
---:-.:
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)
A Slmlllary of the moveable surface caFSbilities is presented in Table 1.
TABLE 1 SUMMARY OF MOVABLE SURFACES GRID POSITIONS DEFLECTION ANGLE AT EACH TYPE OF
-- .... -...::- --
MOVABLE SURFACE (DEGREES) DEFLECTION CONTROL .....
VCK 55,45 4 MANUAL --..:.
SLAT INaOARD, 15,25 3 MANUAL OUTBOARD- 25,35 3 FLAPS MAIN 0,5, 15,25,35 4 ~.-- ~.~.
MANUAL AFT 0,75,10,125,15 1 ....
AILERON (LEFT-HAND ONLY) O,~5,~10,~15,~20 MANUAL
-
---
(NORMAL TO HINGELINEI LH 0, -5, -15, -30, -45, -60 SPOILERS MANUAL
-
RH 0, -30, -60 (NORMAL TO HINGELlNEI
,)
......
HORIZONTAL STABILIZER +5 TO -15 - REMOTE
........
(NORMAL TO ROTATION AXIS)
--
---
Configuration Notation CL"1d Dimensioml Data The various configuration notation is presented in Table 2. Table 3 presents dimensional quantities related to the wind tunnel model. The various grid position notation and corresponding values for deflection, gap,and overhang are presented in Tables 4, 5, 6, a1"d 7 for the slat, VCK, main flap and flar:eron, and the aft flap, resr:ectively.
-----
)
)'
L.". , -",,- - \.~ I
')
TABLE 2 .... ------- ...
CONFIGURATION NOTATION B2A Simulates the DC-X-200 Model D-969N-2l fuselage. Full scale
dimensions: Length = 42.29 m (138.8 ft); constant section
diameter = 602 em (237 in). The aft fuselage tail cone uses
-- -
tre existing OC-10 model parts. The fuselage is configured for tandem strut support system.
U3B Si..'ilulates tre OC-X-200 Hodel D-969N-2l wing and is lofted to reFresent the alrplane wing with a 19 load. Full scalp.
dimensions: S = 212.597 m (2288.457 ft2); b = 47.252 m
(155.027 ft); aspect ratlo = 10.502; A = 0.1407; HAC = 5.35lm
)
(17 .555 ft). The model wing has a removable leading edge, full span VCK flap, trailing edge two-segment flap, outboard aileron on ore side, and spoilers. The wing is constructed of Armco 17.4 steel and contains five rows of pressure orifices.
\ mng-fuselage fillet for B2AN3B.
X2B
\
HIA Horizontal stabilizer for OC-X-200 (slab surface).
VIA Vertical stabilizer for OC-X-200 (slab surface).
N2A Flow-through, short core cowl nacelle configuration (2).
t2A Uew P'llons for mating ~A to wing W3B (2).
ZlA Nacelle strake configuration (attaches to N2A, 2 each nacelle) • ltai!1 and nose landing gear defired for the OC-X-200 airplane.
GIA
)
l1ain gear \'lheel wells with gear eAtended are not provided.
i)
..... _ .r_
--",_::-.-J_c.:-
------
J'J !' ............
'.
""1
)
TABLE 2 (CONTINUED)
a2A The outboard aileron with inboard trim at Xw = 89.020 cm
(35.047 in) and outboard trim at xw = 109.480 em (43.102 in).
TOO hingelioo is located at 75% C.
Inboard spoiler segments fabricated as individual parts.
fl,f2 Superscript R = right side, L = left side, Nooo = both sides.
fl and f2 inboard 00 spoilers with sreet metal aft extension.
flA,f2A Trailing edge step is filled with wax and faired (LB-486A).
This assembly was refurbished and the T.E. step filled with potting (LB-486C).
f3,f4,f5,f6 Outboard spoiler sagnents fabricated as 000 pieoo.
,-)
Leading edge slat inboard of Xw = 36.367cm (14.318 in) and LIA -'
supported at nominal gap = 2.25% C, O.H. = 2.0% C, and 6 SLAT =
25 •
L2A Leading edge slat outboard of Xw = 36.367 em (14.318 in) and
supported at nominal gap = 2.25% C, O.H. = -2.0% C, and 6 =
SLAT 35 • Leading edge variable camber Krueger flap inboard of wing r'3A
station Xw = 36.367 em (14.318 in) and supported at the nominal
0•
gap = 2.82% C, O.H. = -0.725% C, 6 VCK = 55
, Leading edge variable camber Krueger flap outboard of wing
/
4A
station Xw = 36.367 em (14.318 in) and supported at the nominal
0•
gap = 3.5% C, and O.H. = -1.0% C, 6 V a< = 55
LsA '!be inbc:ard va< extensioo to the fuselage.
)
)
'.
" ,., ..;;",/
')
TABLE 2 (CONTINUED) i6A TOO VCK sactlOn at tre pylon interruption.
PIA Inboard main flap of a twcrsagnent flap with inboard trim at
~I = 13.868 ern (5.460 in) and outboard trim at Xw = 30.793 ern
(12.123 in).
Inboard aft flap of a twcrsagment flap trinmed to match PIA and F2A supported from FlA.
A single-slot flaperon ... /ith inboard trim at Xw = 30.793 cr.l
F3A (l2.l23in) and outboard trim at ~1 = 43.411 ern (17.091 in).
Outboard main flap oi' a t\-Jo-s:gnent flap "lith inboard trim at F4A
~, = 43.411 ern (17.091 in) and ontboard trtm at XII = 89.020 em
(35.047 in).
:)
Outboard aft flap of a blo-segnent flap trimned to nldtch F4A and FSA supported from F4A.
tlJing coordirates (spamnse, chordwisa).
~:I' Y\,l Angle of attack, in oogrees, of the fuselage reference plare °FRP relatlve to the equivalent free airstream. Nose up is positive.
Aileron ooflectlcn, ln degrees. Positive deflection 15
Oa
trailing edge dOom.
Aft flap deflecltion, in degrees (see Figure 51).
°FAFr Main flap deflection, in oogree5 (see Figure 51).
°FMAIN
)
@
~} .."
, -j
TABLE 2 (CONCLUDED) Slat deflection, in degrees (see Figure 48).
bSLAT va< deflection, in degrees (see Figure 48).
bVa< Incidence angle, in degrees, of the horizontal stabilizer HIA iH Positive deflectioo is trailing edge dCMll.
Sunmary Code B2AW3BX2Ba2A. Body + cruise wing.
Sl
B2AW3BX2BtT2AI'2AZlAL3AL4AFlAF2AF3AF4AFSAa2Afl,2,3,4 (S,6. Body +
S2 flapped wing + VCK leading edge device + flaps + nacelles, pylons, and nacelle strakes + va< filler blocks.
~
S2-t'13a+W30. Configuration S2 - va< filler blocks.
S3 f 2A· S2-~13B+W3o-fl,2 + flA Configuration S3 + inboard spoiler S4 trailing edge extensions.
tv S5 B2A 3 BX2BN2AP2AZlN-lAL2AFlAF2AF3AF 4AFSA a2AflAf2Af3f4fSf6. Body + flaPFed wing + slat and WUSS leading edge + flaps + nacelles, pylons, and nacelle strakes.
)
(f)
.1
-'
')
TABLE 3 DIMENSIONAL DATA CDMroWNr WIEL SCAlE WITS.
FUSEIArn (~) en (in) Length 198.77 (78.255) en (in) ~xirnum width 28.293 (1l.D9) en (in) M:lximum reight 28.293 (1l.139) HI.tl.l (W:3B) m (ft2) Area 0.4696 (5.055) m (ft) St;an 2.221 (7.286) m (ft) Hean Aerodynamic Chord 0.251 (0.825) Root chord (tra~zoidal wing) em (in) 37.076 (14.597) Total root chord em (In) 51.895 (20.431) en (in) Tip chord (tra~zoidal wing) 5.217 (2.054) an (in) Total tip chord 9.27 (3.65) Aspect ratio 10.502 'la~r ratio 0.1407
~
St;anwise station of f>lAC en (in) 41.580 (16.370) Fuselage station of 25% f>lAC em (in) 160.28 (63.102)
9.veeptack of 25% qv OOg. 28.57
Dihedral("lg") deg. 4.5 lDRIWN1'AL STABILIZER (HlA) m (ft2) 0.1298 (1.397) Area en (in) 70.234 (27.651) Sf5n r·1\C en (in) 19.91 (7.839) em (in) Root chord 27.384 (10.781) en (10) Tip chord 9.583 (3.773) Aspect ratio 3.800 'laFer ratio 0.35 9.veeptack of 25% chord OOg. 30 Dihedral OOg. 10.0 Fuselage station of 25% HMAC en (in) 247.36 (97.384)
Tail length (25% Wz.1Ac to 25% -en (in)
87.076 (34.282) H!.1I'IC) ,
®
"
-
')
TABLE 3 (CONTINUED) CDMIQNENl' mlEL SCME .IIN.rm VERTICAL STABILIZER (VIA) m (ft2) Area 0.098502 (1.0603) em (in) Spm 39.700 (15.630) em (in) W.C 26.731 (10.524) em (in) Root chord 36.759 (14.472) an (in) Tip chord 12.87 (5.065) As{:ect ratio 1.6 'Iap:!r ratio 0.35 deg. 35 SNeepback of 25% chord an (in) 82.301 (32.402) 'Iail length (25% Wr-1AC to 25% VrlAC) QUTOOARD AIIEroN (a2A) 2 2) an (in 54.4 (8.44) Area aft of hin~lire
.-)
% b/2 Spm 18.4
'-
Chord aft of hin~lire %Cw SIQIIER (f1,f2) 2 2) em (in Area (each) 47.2 (7.32) en (in) Sp:m (each) 13.2 (5.18) SpoILER (f3,f4,f5,f6) 2 2) Area (total, are side) em (in 104.660 (16.222) Span (total, are side) em (in) 43.835 (17.258) NACEIJ£. (N2A) an (in) length 32.00 (12.60) em (in) Maximum cowl height 13.7 (5.38) em (in) Inlet diameter (fan cool) 9.85 (3.88) Exit area (gas generator)
cm2 Un ) 6.86 (1.06)
Incidence of thrust lire to FRP deg. 1.6 Toe in deg. 1.8
)
1>
f.l
•
')
TABLE 4 SLAT GRID NOTATION All gaps and overhangs are ~rcent of local wmg chord 0,SIA!
WING ErATIOR_ IDTATION ~ ~ 2.25
~oJ = 14.140 c;n (5.567 in) -2.0
LIM
:<"'1 = 36.357 em (14 .138 in) 2.25 -2.0
1.50 -1.0
Xw = 14.140 em (5.567 in)
LIAB 1.50
~I = 36.367 em (14.138 in) -1.0
3.25 ~.= 14.140 Q1\ (5.567 in) -2.0 LIAC Xw= 36.367 em (14.138 1n) 3.25 -2.0
~ 2.25
Xw = 14.140 em (5.567 in) -2.0
ISO LIAD
Xw = 36.367 em (14.138 in) 2.25 -2.0
Xw = 14.140 em (5.567 in) 1.50 -1.0
LHE
Xw = 36.367 em (14.138 in) 1.50 -1.0
3.25
~ = 14.140 em (5.567 in) -2.0
I'lAF
Xw = 36.367 em (14.138 in) 3.25 -2.0
~ = 36.367 em (14.138 in) 2.25 -2.0
~M 2.25
:~ = 89.020 em (35.047 in) -2.0
1.50
Xw = 36.367 em (14.138 in) -1.0
L2AB 1.50
~oJ = 89.020 em (35.047 in) -1.0
)
J ___ .• __ _ _ ___ -.. ·ooL--......- ..... _------ .. ___ ___ . _____________ ... ___ _
®
~]
')
TABLE 4 (CONCLUDED)
5.sw
moo STATION tpTATION
~ ~
Xw= 36.367 en (14.138 in) 3.25 -2.0
L2Ac
3.25 -2.0
Xw= 89.020 em (35.047 in)
2.25 -2.0
Xw= 36.367 en (14.138 in)
I-'2AI)
Xw = 89.020 en (35.047 in) 2.25 -2.0
Xw= 36.367 en (14.138 in) 1.50 -1.0
L2AE
Xw = 89.020 em (35.047 in) 1.50 -1.0
-)
--'
Xw = 36.367 em (14.138 in) 3.25 -2.0
~
Xw = 89.020 em (35.047 in) 3.25 -2.0
')
, ......
-~ 1P
')
TABLE 5 VCK GRID NOTATION All gaps and overhangs are percent of local wing choru 5~ NlRLSl'A':'ION IDTATION ~ ~ 51.316 2.82 -0.725
x,,1 = 14.140 r.:m (5.567 in)
L3M
Xw = 36.367 em (14.318 in) 3.5 -1
51.31Bo 2,82 -1. 725
x"oJ = 14.140 em (5.567 in)
L3AB
x.., ; 36.367 em (14.318 in) 55 -2
3.=> 1)1.318 -0.725
~I = 14.140 em )5.567 in) 1.82
L3AC :~I = 36.367 em (14.318 in) 55 2.5 -1 51.318
hw :;: 14.140 an (5.567 1n) 1.112 0.275
,)
L3AO
Xw = 36.367 em (14.318 in) 2.5 0
3.~ -1 [,4M
x.....= 36.367 em (t4.318 in) 55° 3.5 -~
L4AB -1 ;~.I :: 111.274 an (43.809 1n) 2.5
!...lAC
2.:' 0 ~AD y~ 14.140 em (5.567 in) 41.3t13 282 -0.725 L3F£
x..t= 36.367 an (14.:U8 in) 45 3.5 -1
0.82 -0.725
Xw ~ 14.140 em (5.567 to) 41.318°
L3AF 40 -1 ~t= 36.367 cr.\ (14.118 in) .5
)
............ ---~- - .--.'~. --_. ~~---- .... ---- ~------' - "- -- -. .. -.--~ .... - : - - -- -
')
TABLE 5 (CONCLUDED)
--
O~ WIOO_ srATION IDTATION ~ ~ 41.3180 1.82 -0.725
Xw = 14.140 em (5.567 in)
L3AG 45 2.5
Xw = 36.367 em (14.318 in) -1
41.3180 1.82 0.275
Xw = 14.140 em (5.567 in)
L3AH
Xw = 36.367 em (14.318 in) 2.5 0
3.5 -1 L4re
Xw = 36.367 em (14.318 in) 1.5 -1
L.w-
Xw = 111.274 em (43.809 in) 2.5 -1
L4lG
2.5 -1 L4AH
)
)
... - .
)
TABLE 6 MAIN FLAP GRID NOTATION All gaps and overhangs are percent of local \Oling chord Inboard Flap and F1aperon Grid
Xw = 14.140 em (5.567 in) Xw = 43.411 em (17.091 in)
Q....H... Q....H... roTATION ~ °FMAIN ~ 1.3 3.2 2.5 6.0 FIM 0.8 3.2 1.5 6.0 FlAB 0.8 2.2 1.5 4.0 PlAC 1.3 2.2 2.5 4.0 FlAn 1.1 3.0 1.6 2.ll FUE 2.2 2.5 4.0 15 1.3 Fl.l\.F
)
0.8 2.L 4.0 1.5 FlAG 0.8 1.1 1.5 2.0 FlAIl 0.0 1.6 3.0 0.0 FlAJ 0.0 25 1.3 2.5 0.0 FlAK 0.5 2.5 1.0 1.3 FIAL 0.8 0.5 1.5 1.0 FlAM / 1.9 1.1 3.5 -2.0 FlAN 35 1.3 0.0 2.5 0.0 FlAP 1.3 0.5 2.5 1.0 FlAR 0.5 1.1 2.0 1.0 FlllS
)
...-- --~-..- ... -------- ....
..J ..... oil
')
TABLE 6 (CONTINUED) OU'IB)ARD FLAP GRID
Xw = 43.411 em (17.091 in) and 89.020 em (35.047 in)
OFM8IN ~ U&lla. IDTATION 2.5 6.0 F4AA 1.5 6.0
4M3
4.0 1.5 F4AC 2.5 4.0 F4AD 3.0 2.0 F4AE 2.5 4.0 15 F4AF "'- 1.5 4.0 F4AG 1.5 2.0 F4AH
)
3.0 0.0 F4AJ 25 2.5 0.0 F4AK 2.5 1.0 F4AL 1.5 1.0 F4Al1 3.5 -2.0 F4h'l 35 2.5 0.0 F4AP 2.5 1.0 F4AR 2.0 1.0 F4AS fLAPERJN DIFFEEENl'IAL roSI'l'ION
Xw = 43.411 em (17.091 in)
IDTATION ~ !laH...
cSPFLAPEOON 2.5 1.0 F3AR
--
~- :;.:::::-:- -
)
" - /-
74 ~
"'"
~ s .....
~
-- ..... ~-::::---
--:;,
)
TABLE 7 AFT FLAP GRID NOTATION All gap; and overhangs are fercent of local wing chord :~ = 14.140 em (5.S67 in)
Xw = 30.793 em (12.123 in)
b
Fm IDTATION
QA ~ Q....lk ~
7.5 0.3 O.B 0.4 1.1
F2M
10 0.3 O.B 0.4 1.1
F2A8 12.5 0.4 0.5 0.4 0.5 F2AC 0.4 0.4 0.5 0.5 F2AO ~J ; 43.411 em (17.091 In) and 89.020 em (35.047 in) ~ Q...H.&.
J
0.5 7.5 1.5 F5M 0.'; 10 1.5 F5A8 12.50 0.75 0.75 FSAC 0.75 0.75 FSAO
C-l
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InstrlJreIlta tion Aerodynamic forces on tre model were reasured by tre Aires '!ask ~1ark II 10.16 cm (4 in.) diameter internal balance at the Ames 12-Foot Pressure \'lind Tunrel (LB-486A test). For the NASA Langley V/S'IDL Wind Tunnel (LB-486C test), the balance used was the Langley 748 5.08 em (2 in.) diameter internal balance.
Pressures over the macEl wing, out.boord aileron, slat, VCX, and flap systems were reasured by means of a scanivalve system installed within the fuselage nose. The spanwi~ p:lsition of the pressure rows is indicated in Figure 53.
// The wlng and clean leadlng edge orifices were located streamwise at the percent span locations shown in Figure 53. The slat orifice rows were
--
0 0 located streamwise with the slat deflected 25 inboard and 35 outboard.
For the VCK, the orif ice rows were located streaIrMi~ with the inboard and outboard leading edge device deflected 55 • The aileron orifice row was located streamwise with the aileron at 0 deflection. M3in, aft flap, and flaperon oriflce rows were located streamwise with 0 deflection on all flap
J
canporents.
The wlng chorrnnse pressure orifice locations are shown in Figure 54. Slat and \'lUSS pressure orifice locations are detailed in Figure 55. Comparable values for the VCK are presented in Figure 56. ChordNi~ pressure orifice DIMENSIONS IN PERCENT SEMISPAN I I
~
FIGURE 53. SPANWISE POSITION OF PRESSURE ORIFICE INSTRUMENTATION
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. , FIGURE 54. WING CHORDWISE PHESSURE ORIFICE LOCATIONS locations for the two-segnent flap systelTl .:ire shown ln Flgu:::e 57, dnd for the flaperon ln Figure 58.
for the Ames test, the lnglE: uf attack of the fuselage reference plare was measured by electrolytlc alifJn:nent bubbles housed 1 n the fuselage nose.
0 0 From an angle of atta,x of -6 to 0 , the macEl was pitcred by the external 0 0 pltch drive. From 0 to +10 angle of attack, the fuselage was pitched using the fuselage lnternal pItch drive, whlle n1<llntaining the balance at 0 • 0 0 0 A 0 bubble on tre halance housing was monltorell such that 0 balance attitude \O/ati maintalnec1. For angles of attack greater than +10 , the fuselage was ~itched USlng the external pitch drive with a 100 angle maintained uetween the balance axis and the fuselage axis. For the NASA-VjS'lOL test, a rJASA furmshed electronic inclinareter was used for the determinaticn of angle of attack.
)
So _ ...
J
.-
-
/ ..... //
--- ....... ---==--~---
)
21 • 5 6 NO I ,~ WRP :zo WUSS PRESSURE ORIFICE LOCATIONS Ill. CLEAN WRP CHORDI ·' .... ::;1 .lllL\),.\T , .
.... "" .. o,,~
:ZOO 30S SOO 725 900 i x "bl2 .,
\\\ ~2 2~5
33939 55636 80673 100 147 ~.
ENTI~'ElER 187621 1\33621 1219041 13176\1 (394281 '> IINCHES' I 313(LEI 363(LEI 447 (LEI S 00 'LEI 32 tLE' 45 40 2 46 52 60 a:
UPPER ... 3 70 55 60 65 75
4 100 80 80 SURFACE 90 100 '" ::ii :;) 5 140 110 100 120 140 z 6 170 140 130 160 190 I ... .
1 160 :ZOO 22S I ::!
LOWER 17 50 5S 60 I 18 80 80 SURFACE 90 80 90 i<
~
• 0 l' 0
0 19 140 130 ISO I~O ____ ~:O :zo 200 220
-
I
FIGURE 55. SLAT AND WUSS CHORDWISE PRESSURE ORIFICE LOCATIONS ~
,~ "
-INSTALLED AT I} = 50 PERCENT ONLY
.... ~
)
FIGURE 56. VCK CHORDWISE PRESSURE ORIFICE LOCATIONS (TYPICAL FOR ALL STATIONS EXCEPT AS NOTED)
-
..
~ - ... .. ~
•
-
'}
ORIG:N~L : OF. POC~~ ~.~.
7.5 5 '125 40 60 80 90 TE 2.5 LE
/;~To
5 70
,-s,(1 /
GO FIGURE 57. TWO·SEGMENT flAP CHORDWISE PRESSURE ORIFICE LOCATIONS
~) (TYPICAL FOR ROWS AT 11 = 20,50, AND 72.5 PERCENT)
25 7.5 30 45 80 90 IE 5 110 LE CENTER FLAP MLL 2.5 FIGURE 58. FLAPERON CHORDWISE PRESURE ORIFICE LOCATIONS (AT rl = 30.5 PERCENT)
)
~ ~
_.---.1._-- ~ ______ . _
•
r.. .. ~.....-~~J,o..,.~ ~. ..... ... ::e:rs...,... re== JIISW , ...... 7'&___~.... :~ .!':. ...
-; 0;: ~0:.~.
The horizontal stabilizer incorporated a remote drive and dual position potentiometers for tail incidence determination during a run.
MoCel Installatioo TOO model was installed in the NASA hIes 12-Foot Pressure Wind Tunnel on the tandem support system shown in Figure 59. 'Ihe mcxEl was pivoted about the main strut pivot point and was pcMered by the aft pitch strut. The entire strut system was nOD-rretric, (i.e., air loods on the strut arc not sensed by tre balance). 'Ihe struts entered the fuselage as far aft as practical to minimize the aerodynamic interference effects on the mcxEl.
TOO saJre support system "las utilized during the NASA Langley V/S'lOL test program. It was adapted to the existing V/S'lOL tunnel structure and extensions for the main, and pitch struts were added to the basic tandem strut system. The extensions permitted the model to be located rear the vertical position of the tunnel centerlire (see Figure 60).
~
1'.
JOe ..-ft ..
dALA .... Ct , ~rJ ",, rENTE:R -, ,\\.
'" 1%J4 I" 'Jlb
/, \ ~~jt~l;~ ~~
kit ' -~4S2ttr~~2 I
29 ~7 111641 ORJJ ., rllJ. U320'
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'IK II S:'L.\~ E ...... /' / L... DAft. AEFEf'ENCE C[!\IrEA Ot',e 4.;,.lIt,~ IN C~Nrl'AETfAS ,U.O"tt:S, ',HmEl SC.\lE FIGURE 59. MODEL INSTALLATION IN THE NASA AMES 12·FOOT PRESSURE WIND TUNNEL
)
~
...
:..-.... ~ ... . ________ J.
~~~~~;ZWSWa"C7?:CR5!:O__ aa.3'U'3...AI'''' ___________________ _ -; 'jf T5 • 500 431200 17) M5 16028163102) MS - 65301 125 709) \
!
LANGLEY 748 BALANCE PITCH STRUT BALANCE CENTER ___ -.J PITCH STRUT EXTENSION 't MA~T SUPPORT ~
TOP !'URFACE _l
'J -;T.) T:,oJ:I;!OL ~l,).)H -,-' -~
~)
PITCH DRI'IF MECHANIS~,l
/t//
I
DIMENIONS IN CENTIMETERS IINCHES) MODEL SCALE FIGURE 60. MODEL INSTALLATION IN THE NASA LANGLEY V/STOL WIND TUNNEL
)
_ ..... - -- . .-...- ... , .... -
rl'r. .. _ ...
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PRECEDING PAGE BLANK NOT FILMED
')
EY.F£RIMENTAL reSUL'lS AND ANALYSIS Test Plan and Facilities Two wind tunnel tests were conducted between 30 November 1978 and 29 November 1979, to aid in the evaluation of the la.-l speed charactenstlcs of a high aspect ratio supercritical wing transport configuration. In !"loth tests tre conflguration \o1aS tte D-969N-21 wide-body transport with the ~J3B wlng and high lift system.
The first test (LB-486A) was conructed at the NASA limes Re~arch Centcr In the 12-Foot Pressure Wind Tunrel from 30 November 1978 throuqh 26 January 1979. Data for various l-1ach and Reynolds numbers were obtaired cilring this test for selected configuratl.ons. A majority of the configurations tested were evaluated at a high Reynolds number conditloo. Figure 61 presents the configurations evaluated wring tre Ames wind tunrel program.
The second test (LB-486C) was conci1cted at the NASf.. Langley Research Center
,)
in the V/S'lOL Hind Tunrel from 1 November through 29 November 1979, )olntl~' \'Ji th actlvi ties under re lated NASA Contract NASl-15327. Spoller configurations, fabricateo under the current contract, "'ere part of this test.
ESTABLISHED THREE·DIMENSIONAL HIGH LIFT CHARAC I ERISTICS FOR HIGH ASPEC'""T RATIO SUPERCRITICAL WING AT HIGH REYNOLDS NUMBER EVALUATED THE FOLLOWING AERODYNAMIC ('HARACTERISTICS o BASIC CLEAN WING o SLAT AND VCK LEADING EDGE DEVICE OPTIMIZATION " :'INGLE AND TWO SEGMENT FLAP OPTIMIZATION () NACELLE/PYLON AND LANDING GEAR EFFECTS " EFFECT OF INBOARD VCK AND SLAT SPAN o HORI ZONT A L TAl L·ON CHARACTE RIST ICS FOR SELECTE D (.ON FI GU RATIONS o CLI:AN TRAILING EDGE CHARACTERISTICS o CLEAN LEADING EDGE CHARACTERISTICS o '1CK IfJEI.L EFFECTS o All ERON EVALUATION WITH HIGH LIFT SYSTEM ., EFFECT OF MINI· flJFI S ON HIGH LIFE CHARACTERISTICS " ROLL SPOI LER CHARACTERiSTICS " (jROIiND SPOILER CHARACTERISTICS FIGURE 61. AMES 12·FOOT TEST RESULTS
)
~ --~-- ...... __ ., ...... 41---..
,:.».~ - .." So .... ..
~
in
)
Test Technique 'fhe test technique and data obtained were similar for both wind tunnel facilities. Test conditions are sham in Table B. '1l2 1"", values for the tip Reynolds number are noteworthy.
TABLE 8 TEST CONDITIONS
yn x 10- ~x 10-6 %nP x 10-
I\Ift x 10-6
FACILI'lY M1lOI NUMBER AMES 12-ft 0.20* 20.3* 6.2 5.12 1.BB AMES 12-ft 0.20 11.5 3.5 2.B9 1.06 AMES 12-ft 0.20 4.53 1.3B 1.14 0.42 AMES 12-ft 0.26 11.5 3.5 2.B9 1.06 AMES l2-ft 0.32 11.5 3.5 2.B9 1.06 4.53* 0.42 T.ANGLEY 0.20* 1.3B* 1.14 VjS'IOL
J
*norma1 test condition The Ames 12-foot wind tunnel test section has a circular cross-section of 3.65m (12 ft) diarreter. Flat areas on floor, ceiling, and both sides reduce this dimension to 11.3 feet. The Langley V/S'lOL wind tunnel has a test section height and width of 4.42m (14.50 ft) and 6.63m (21.75 ft). The model was mounted on a tandem strut system and tested through a maximum 0 0 angle of attack range of -6 to +30 • For configurations without a leading edge deVIce (i.e., crUIse wing), this angle of attack ran~ was rewced.
The model was always tested with 0 sideslip.
Fluorescent mini-tufts (Reference 7) were used in both wind tUIlllel tests for flow VIsualization of the various high lift configurations. This technique uses a very fire Sfecially treated mooofilament thread which, wren struck by filtered ultraviolet light and photographed with appropriate filters, becomes readily visible on the photographic regative. '1l2 thread, being of extrerrely small di.arreter, produces little disturbance to the flow.
)
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~
~ So ... .- i.
/ ~4N!1fBii'P'''~U':'~~~~~~~''~1~.lI:7*'~~~_:;:;'';~J~:~'{i::'~~~"" .... _. _" _-..
')
Unless otherwise noted, tOO configurations were evaluated w~th tOO nacelles, pylons, and nacelle strakes attached, and the horizontal tail off. The latter permitted the evaluation of high lift cornp:ment inr:rements In 11ft and drag. Tail-on characteristics were obtained for a selected number of configurations representing the better configurat~ons resulting from the leading and trailing edge device optimization studies. Unless landing ljear 1ncrements were being obtained, the cruise and takeoff configurations rad the landing gear removed, and the landing configuration had the gear attached. r'iini-tufts were attached to tOO right wing panel [or most of the configurations tested.
Thc first test program started with the cruise wing configurations. This \lIas follCMed by t.re VCK leading edge device study at nominal landing and takeoff flap defl. ..... llons. 'Ire two-regment and single-slot flap optimization with tIle VCK leading edge device was ~valuated next. Clean lead1ng (·.tge characterist ics al the optlmized flap positions were then evaluated, dnd \'/ere follONed by t.re slat study at the previously deflned flap positions.
Clean trailin9 ~dge char;lctenstics tOL the optim1zed va~ and slclt lXlSl.tions
~
were also obta~ned. Tall-on, U;lch m.unber, Reynolds number, nacE:lle/Pl"lon, and a1leron effects \,Are also obtaired as the awropnate configuration was being evaluated.
\'1ind 'l\mnel Test !<esul ts and AralYS1S Due to thc complex nomenclature assoclated wIth the high lifL
configurat lons, plotted data haVE: I::een identified bJ a smrmary code (see S1
through S5 in Table 2). In addition, a leading edge grid posit1on and deflectlOn for the 1nboard and outboard segments has been identified for the high 11ft configurdtions. For example, 15D/25D notation for the slat, slgmfies that thE: inboard slat deflection is 15 and the gnd position is AD, and for the outboard slat, the deflection ~s 25 and the grid peSl t10n is AD. Table 1\ shows the gap and O.H. to I::e 2.25 percent and -2.0 percent for both the inboard and outboard slat. Similar nOlrenclature has been used for the va< (~.e., 45E/45G).
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.., s. ....
..
~
-- ........ ~ ........ gg .. --~MM __ W.~~MM__ ~.
~~ .. ___ AL" ___ ~":"~" ___ "~ •
')
An example of the two-segment flap deflection, gap, and overhang nanenciature used in tre plotted results is 25K/10B. The slash in this case denotes a main flap/aft flap relationship. The example signifies a 25 deflection for the main flap at position AK and an aft flap deflection of 10 at position B. Tables 6 and 7 indicate an outbrerd gap and O.H. of 2.5 percent and 0 percent for the main flap and 0.5 and 1.5 percent gap and overhang for the aft flap.
The single-slot flap configuration is denoted by a zero follaoling the slash (i.e., 25K/0). By use of the configuration summary code, leading and trailing edge nomenclature, and the tables, the specific configurations in the plotted data can be obtaired.
Test data presented in the following sect:ions are without strut tare corrections. TOO relatj:"e perfornance was determined in this fashion, and selected data and final optimized configuration data were strut-tare corrected. The latter is presented in the Spoiler Deflection Effects and Summary of High Lift Characteristics sections. Strut tare corrections were
'~)
based on an experiIrental evaluation of strut tares for a previous transport configuration, the technique being discussed in Reference 6.
~ possible test variables and high llft configurations were quite numerous for this wind tunnel model. Table 9 presents a figure index for the high lift configurations and test conditions presented in this section.
Cruise Wing The initial configuration tested was the cruire wing-body with the nacelles and pylons removed. Figure 62 illustrates this configuration installed in the NASA Ames 12-Foot Pressure Wind Tunnel. Also ShCMIl in the photograph is the tandem strut support system, and the va< filler blocks (installed in the lower surface of the leading edge). The basic high Reynolds number characteristics (lift, pitching manent, and LID) for the wing-bcdy are shown ~n Figure 63 (see Run 22). Figure 63 indicates that CLMlU{ of 1.513 was obtaired at an angle of attack of 12.57 • The lift coefficient at zero
degrees angle of attack was 0.4. The LlDMAx for this configuration was 20.4
and at 1.2 Vs the LID was 19.45.
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FIGURE 62. CRUISE WING CONFIGURATION IN THE AMES 12-FOOT PRESSURE WIND TUNNEL (NACELLES AND PYLON OFF) The basic wing-body is seen to have an initial outboard stall of significant proportions in terms of the lift loss and the small angle of attack range over which this phenomenon occurs. The result of this trend is a significant reduction in wing-body stability after. stall (Figure 63). Also presented in Figure 63 is a comparison of the aerodynamic characteristics with and without tufts. Good agreement in the lift and pitching moment through CL~mx is shown. Some differences are noted in these quantities after CLM1\X, but they are relatively small. The draq with the tufts on (Run lR) is seen to be lower in Figure 63 by 10 to 20 drag counts depending on the eLI and this results in higher L/D values (Figures 63). Evaluation of the pressure data comparinq Runs lB and 22 did not indicate a significant bnprovement in trailing edge pressure recovery.
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'rhe cruise wing-body configuration, with the N3 wing, had previously been tested in the Ames Research Center II-Foot Transonic Wind Tunnel. This model was a four-percent scale model and was sting-mounted. The current data was corrected for strut tares and compared with the high-speed experimental data. Various lift p;lrarneters are canp;lred in Figure 64. Good agreement with the high speed data is indicated for the lift quantities shown in Figure 64. The lift and drag c!re canp!red in Figure 65 for ~tlch numbers of 0.2 and 0.5. The lift variation shown in Figure 65 shows the expected increase in high-speed lift curve sloJ;e, and also mdicates that the CLr1AX at 0.5 ~tlch number is influenced by compressibility effects not apparent in the Im-1-speed data. The LB-486A Reynolds number based on the HAC was very similar to the high speed test condition at a Mach number of 0.50 and is shown in Figure 65. Good agreement in the drag is obtained for the intermediate CL range. Canpressibility effects in the high-speed data are noted for CL values above 0.6. Pitching manent canp;lrisons were rna~ difficult by the differences in geometry (boat-tailed configuration versus sting-mounted). Direct comparison of pitching manents indicated that for
J
./ the intermediate CL ranqe, the 10\1-~d values were 0.03 more positive than the high-speed data.
~hni -tuft pictures of the \-linq, for Hach number of 0.20, are presented in Fiqure 66 for anqles of attack before and after CL~'IAX' At this point in the 16ri--------------------------~ NACellES AND PYlONS OFF HORIZONTAL TAil OFF HIGH·S?EED DATA 1.2
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FIGURE 64. COMPARISON OF LOW· AND HIGH·SPEED CHARACTERISTICS FOR THE ACA CONFIGURATION .".
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loU 'F Rp ANGLE OF ATTACK (DEGREES) FIGURE 65. COMPARISON OF HIGH·SPEED AND LOW·SPEED LIFT AND DRAG FOR CRUISE WING CONFIGURATION test progralT', the tuft 1 .. 'Iqt:h han not been lengthened to ir.lprove the reaclah111ty of the pictures. tJevertheless, Figure 66 does illustrate the sta 11 phenomena of the high aspect ratio \1in9 at high Re'tnolns nUr.lber conditions. Fiqure 67 presents tbe chor&vise pressure distributions of the 0 0 0 five strear.lwise pressure rQ\vs for 0'C' r-1AX (12.57 ), and 1 and 2 past.
L aCL:-!AX. l,t aeLt_lAX suct10n peaks are evinent f,)r all spamlise locations.
51ightlv neqative trailing edqe pressure coeff1cients are noted for th1;' condition at all sr.amviE".e locations except lJ . 20 percent. Large spanwise flo'", angles are 1mhca ted 1n the corresponding tuft photo for the trailing
enge region. AtapJUl = 13.55 (10 past stall), the 50-percent span station
indicates separation near the leaninq edqe. The 7?.5-percent spm stut10n remains attacher'\ ann tins is confirmed by the tuft photographs. The QO-oercent sran station 1S also separaten at the leaninq edqe. At O'FPJl ~ 14.S0o (2 past stall) the ,0-, 72.5-, and 9D-percent span stations are scpar3ter! at tre leaning ec'lq<.>. On the other ~and, tre inboarr. stations are
) still heavily loaned. Other pressure rlata (not presented) indicated the ?fl-
and 3D-percent stations stalled at 0FRP = IS .44 •
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0" • "w'" -r ... '" .,- .off:' "' C "FRp· 1450DEG FIGURE 67. EXPERIMENTAL CHORDWISE PRESSURE DISTRIBUTION FOR CRUISE WING (NACELLES AND PYLONS OFF) (CONCLUDED) ..
....
~) Figure 68 shows the variation of sectional lift, obtained from an lnteqration of the pressures, for the five pressure rows. 'nle large, rapid lift loss at stall for the 50-, 72.5-, and 90-percent stations is noteworthy. Figure 69 presents the spanwise variation of sectional hft.
0 0 Significant variation in rolling manent for angle3 of attack 1 or 2 above aCL MAX indicated some asymmetry in the stall pattern. Rolling manent incremental values were negative, indicating premature left winq panel stall. Flow syrnrretry \-laS regained at higher anqles of attack.
The resie winq-body was also tested at a Reynolds number of Rtlr!AC = 1.13xl0
and a f.1ach nUMber of 0.20 (atmospheric conditions for the hles facility) • 'T'he results are presented in Fiqures 70 through 72. A. canpuison of Figures 63 and 70 sho\-Is that the CL~1AX is reduced frm 1.51 to ] .13, the reqnitllde of the post ~tall lift loss is rlecreasen, and a posi t1ve Cm shift of anproximately 0.02 is rlpparent for angles of attack prior to stall. The Cl'C1r-lA.,{ has been reduced fror.1 1?S7° to ft.43°, and the confiquratlon stllJ exhibits the same pitch variat10n for the anqles Just after 0 C'LHAh. The
:)
(I';D)~1AX is reduced from 20.4 to 15.3 by the decrease in ne~'nolds number.
Fiqure 71 presents tre sectional lift c:haracterirltlCS for Run ~l. Comparinq F1gures 6R and 71 shows the sect lanaI lift values obtained for the 2(l-~rcent spm station are L. ,. same. Significantly lower values of c£f.~AX \'ere obtained for the remaining stations. The 30-r:ercent spln station for the 1~1 Reynolds number has a c1rw{ of 1.2 and a gradual stall followed hy a moderate lift loss. This is different than the high Reynolds number trend.
The outboard stations (50-, 72.5-, and 90-~rcent) show the rapid lift losn (i.e., leading edqe type separation) and lose almost half of their lift by 2 after their cIMl\X. Clearly the outboard trends are similar to the high P.eynolns number data, but occur at a lor.-ler angle of attack.
It is also notetoTorthy that, at large anqles past CLMl\X, the cl values tend to reach similar values. A chanqe in the type of stall (Le., from a leading edge to a trailing edge type separation) for the flight condition tolould alleVlate these adverse high angle-of-attack outboard \-ling parel lift characteristics. The large increase in the angle of attack for outboard \.,ing pane] stall, due to an increase in Reynolds mnnber, implies that at fliqht conditions the outboard [.anel may stall later than the inboard tolinq.
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MACH NUM8ER • 0.20 RUN 22 R - 512 x 10 N MAC SYM 'I HOR!ZQ..~TAL !~I~OFF 20.0 30.5 50.0 l:l.
12.5 90.0 \l 1.4 1.0 c,
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"c '-MAX 00' . , . 0 4 Ii 16 20 24 Q FRP tDEGREES) FIGURE 68. VARIATION OF SECTION LIFT COEFFICIENT FOR THE CRUISE WING (NACELLES AND PYLONS OFF, HIGH REYNOLDS NUMBER CONDITION)
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ORIGINAL PAGE 19 OF. POOR QUALITY RUN 22 .,v-., I}11 a., MACH - 020 '2 'il A
rue RN -512xl0
L IJ~~ MAX MAC
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08 09 10 01 02 II. ()4 O.!'> 06 07 Ull FIGURE 69. SPANWISE VARIATION OF SECTION UFT COEFFICIENT FOR CRUISE WING (NACELLES AND PV LONS OFF, HIGH REVNOLDS NUMBER CONDITIONI
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A LIFT AND PITCHINIl MOMENT FIGURE 70. EFFECT OF REVNOI.DS NUMBER ONAERODVNAMIC CHARACTERISTICS OF THE BASIC CRUISE WING-BODY CONFIGURATION ,---"
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B DRAG FIGURE 70 EFFECT OF REYNOLDS NUMBER ON AERODYNAMIC CHARACTERISTICS OF THE BASIC CRUISE WING·BODY CONFIGURATION (CONTINUED)
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-z ~IrT COEFFICIENT C LIFT/DRAG RATIO FIGURE 70. EFFECT OF REYNOLDS NUMBER ON AERODYNAMIC CHARACTERISTICS OF THE BASIC CRUISE WING·BODY CONFIGURATION (CONCLUDED)
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00. . t I I '---1 a fJ 12 16 20 24 ~ HP u)EGAEE<.,1 FIGURE 71 VARIATION OF SECTION LIFT COEFFICIENT FOR THE CRUISE WING (NACELLES AND PYLONS OFF, ATMOSPHERIC REYNOLDS NUMBER CONDITION I
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MAX -12 r- 5.11 .~~Eo 11~.5'" j 8.43 1.14 A ~ p r ·F", FOR C , MIN'
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01 02 03 0.4 0.5 0 06 07 09 TJ FIGURE 72. EFFECT OF REYNOLDS NUMBER ON THE SPANWISE VARIATION OF C p MIN 'i'llis would improve the pitching moment characteristics for this wing. It should be noted that the clMAX values at the 20-percent station did not change significantly with R€!ynolds number. '!he 30-{:ercent station O'CLMAX did increase. The evaluation of the low speed supercritical airfoil stalling phe~non at Reynolds nmnt::er awrooching flight conditions is thus a critical area of technology development.
It is pertinent to note that v13B was one of the earlier wing geanetries evaluated in the high speed development. This wing was characterized by significantly greater outboard loading than the finalized designs (see ReferenQ: 5). Future wings to be evaluated in the high speed regime have reduced outboard loadings which should alleviate the imbalance between the ) outboard and inboard angles of attack for stall.
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Significantly lower values of minimum pressure coefficient (Cpt.tIN) \~re indicated for the ICMer Reynolds number test condition. Figure 72 presents a comparison of the CPmN spanwise variation for the high and 10\,1 P.eynolds number condition. The differences are greatest at the 72.S-percent span station.
Nacelle/Pylon Effect on Cruise Wing. - The effects of nacelles, pylons, and nacelle strakes are shown in Figure 73. Addition of the nacelles and pylons (Run 29) resulted in essentially the same CLMAX and lift v~riation after CLl-we as the basic wing-body (Run 22). The pitching moment indicates an improved stability trend after CLr1AX. Addition of the nacelles is destabilizing prior to CLMAX. The drag increment at 1.2 Vs due to the nacelle/pylon is 0.0060 and the L/D is reduced to 17.7. Rolling moment characteristics Indicate the same trend (left wing stall) as the basic \ling-body, but ''lith slightly different p:ak rolling marent values. AdditJ.on of the nacelle strakes (Run 23) results 1n the same CLHAX, a reduced lift loss and SImilar trends of reducerl stability after eLl-lAX as the wing-body
)
configuration. The drag is slightly larger (0.0005) than Run 29 values near 0.69 CL~mx. The rolling moment trends are similar to the basic wing-body values.
rtIni-tuft photos for the nacelles, pylons, ann strakes attached configuratIon are shown in Figure 74. Changes in local flO\-I angle are f:"ident behind the nacelle/pylon group. COt:lparison with Figure 66 (the cruise wing) indicates the sane outboard separation phenomenon. However, the flO\'1 on the wing surface behlOci the nacelle is attached. Canplementary flow-visualization photos for Run 29 (strakes off) indicated the area behind the nacelle is sep:lrated for the angles of attack larger than Cl'CLl-lAX.
Chordwise pressure distribution plots for the configuration '-lith the nacelles, pylons, and strakes attached are presented in Figure 75. The angles of attack selected \-!ere Cl'CLMAX and higher. At an Q of 13.58 , FRP the 50-percent span station shows a collapse of the suction peak. The correspondinq flow-visualization photo, \olhile indicating some tuft activity near the trailing edge, does not show significant tuft motion near the wing ) midchord reqion. '!he 50-percent pressure station is located on the left L07
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ANGLE or ATTACK-DEG -0. !>O A. LIFT AND PITCHING MOMENT \ FIGURE 73. EFFECT OF NACELLES, PYLONS, AND STRAKES ON AERODYNAMIC CHARACTERISTICS \ OF THE CRUISE WING \ ,..
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OF POOR QUALITY A. a - 12.58 FRP
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D. 0FRP - 1656° FIGURE 74. MINI·TUFT PHOTOS FOR THE CRUISE WING WITH NACELLES. PYLONS. AND STRAKES I ATTACHED (CONCLUDEDI
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1'[01:[.1 c..oo t1:lIrt .. r C"lHl A. "'FRP • 1258 DEG (a C L MAX FIGURE 75. EXPERIMENTAL CHORDWISE PRESSURE DISTRIBUTION FOR CRUISE WING (NACELLES, PYLONS, AND STRAKES ON) , I '-..,../
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\l1ng, and the tufts were located on the right wing. It is possible that part of the flow-symretry problem near stall nay be located in this area of
the wing. At aFRP = 13 .58 , the pressure distribution for the 90-r,ercent
span station indicate se'f6rated flow. The tuft photo, however, indicates only moderate trailing edge activity.
Figure 76 prssents the sectional lift values for Run 23. At aFRP = 13.58 , the 50- and 90-r,ercent S'f6nwise positions (left wing) show a lift loss. The lift loss at 90 percent semispan is more than half the c.t :< value. '!he rna 72.5- and 90-r,ercent stations have the sane ~ of lift loss at stall with the nacelles, pylons, and strakes att.'!cheCi. '!he only differences noted are reduced lift at the 30-percent span station (near the nacelle), but no stall, and a reduced lift loss at the 50-r,ercent SIBn station.
Reynolds Nurrber and f:bch Number Effects on Cruise Wing. - This study was performed with the nacelles, pylons, and strakes attached. Figure 77 presents the influence of Reynolds number on the aerodynamic characteristics. The angle of attack for CLMAX and the CLMAX attained
J
decreased with the reduction in Reynolds number. The CLMAX for RN~1AC =
6 6 5.12x10 \-JaS 1.54, and for RNMAC = 1.14xl0 , a CLr1AX of 1.15 was obtained.
lI. positi'!e pitch increrrent after C~ is aPJ;arent. The low Reynolds number data show a pitching moment level shift for most of the angle of attack lange. These data also indi~~te a less rapid buildup of the positive pitching manent increment \-lith angle of attack p:lst acLMAX.
Figure 78 illustrates the low Reynolds number chordwise pressure distributions for angles of attack near, and above CLM1\X. '!be 72-r,ercent
span station suction feak has collapsed at a FRP = 8.420. At 10.42 and
12.44° angle of attack, the 50-, 72-, and 90-r,ercent stations are sep:lrated at the leading edge. Although the Reynolds number has changed the angle of angle of attack for CLMAX and the C1MAx level, the basic trend of outboard stall is still evioont.
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WITH NACELLES, PYLONS, AND STRAKES ATTACHED
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A. LIFT AND PITCHING MOMENT FIGURE 77, EFFECT OF REYNOLDS NUMBER ON CRUISE WING WITH NACELLES, PYLONS, AND STRAKES ATTACHE:D
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Figure 79 presents the influence of !tlch mmi)er on the same configuration.
These data were obtained at %MAC = 2.89xl0 • '!be effect of Mach number is
to decrease CLMAX (CLMAX = 1.43, 1.39, and 1.32 at !tlch = 0.20, 0.26, and
0.32, respectively). Also, the effect of Mach number is to decrease the angle of attack for the outboard stall. At a Mach number of 0.32, the reduced stability shift has already been initiated at 9.5 • The overall shift is more gradual with angle of attack, but the basic pitch trend still remains. !tlximum L/D values are indicated for a !tlch nlBtlber of 0.26.
Figure 80 presents the section lift coefficients for the Mach number of 0.32 (Run 26) Examination of the chordwise pressure distribution plots, indicated leading edge se{Bration at the 50- and 72-~rcent SP:ln stations at an FRP of 10.44 • At 11.46 the three most outboard stations were separated. The outboard lift loss is apparent, as is a reduction in inboard sectional lift curve slope at an FRP of 11.46 • Figure 81 shows the spanwise variation of CPMIN (minimum pressure coefficient) for the l13ch
number tested. Also shCMIl for reference, is the RNftAC = 5.12x10 and Mach
,)
number of 0.20 results (Run 23). CPCRIT for th~ outboard leading edge region is al.,o shown. The corresponding value at a Mach number of 0.20 Clearly, the data obtained at 0.26 and 0.32 ro1ach would be over -16.0.
numbers indicate that the suction peaks are approaching sonic conditions.
Extrapolation of the CPMIN to the next angle of attack, for the outboard sections, would indicate supercritica1 flow. This was not achieved, and raises the possibility of shock-induced separation for these locations.
Cruise Wing with Horizontal Tail. - The horizontal tail-on characteristics for the cruise wing with the nacelles, pylons, strakes, and vertical tail attached is shown in Figure 82. Horizontal tail deflections of 00 , -50, and -100 were evaluated. '!he pitching manent curves indicate pitch-up prior to, or at, CLMAX for the various deflections.
Figure 83 presents the effect of Reynolds number for the tail-on configuration. CLMAX and aCLMAX are reduced at the lower ~ynolds number.
The lift loss after stall at the low Reynolds number is snall. HCM!ver, pitch-up is indicated for both ~ynolds nllnber conditions. '!he nagnitude of the positive pitch increment is larger for the hi~h Reynolds number
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A. LIFT AND PITCHING MOMENi FIGURE 79. EFFECT OF MACH NUMBER ON CRUISE WING WITH NACELLES, PYLONS, AND STRAKES ATTACHED
/
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u
'--./ ,------------- ------ - -- - -- - -- -- ----------.
MODEL LB-Y8b R
RUN MACH SYM () 020 24 026 0 25 0.32 26
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DICAU COErrICl£NT
_________________________ J
B. DRAG FIGURE 79, EFFECT OF MACH NUMBER ON CRUISE WING WITH NACELLES, PYLONS, AND STRAKES ATTACHED (CONTINUED) /' \ '-~
u
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,------------- -
MODEL LP.-Y8b ;::; MACH SYM RUN
0.20 24 0.26 0 25 032 !.. 26 oll- Q,. 9_
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:: b .lr:;-~b -?- ~1'~ ,OcF;I:I[NT L--_ _ _____________ _
------------- ------------------------
C. LIFT/DRAG RATIO FIGURE 79_ EFFECT OF MACH NUMBER ON CRUISE WING WITH NACELLES, PYLONS, AND STRAKES ATTACHED (CONCLUDED)
)
OmGINAL PAGE IS OF POOR QUALITY I\~AUt 0 J2 R 289.1~ NM~ hORIZONTAL TAil OFf- I ~ OB
-)
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If I
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I 7"J. r, <> V ' ~IJ 0 () !
_______ L- ____ ~_
--L- _ ---.J
II .~ s 1 : 10 2-1
"
'FRP oPE .RE=~o fiGURE 80 'JARIATION Of SECTION LIFT COEFFICIENT FOR THF CRUISE WING WITH I'IACELl ES, // PYLONS AND STRAKES ATTACHED (MACH; 0.32)
J
".
., ...
-
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)
l..
Of PG~ .. ',,-' -
MACI! fi "1O~R~
SYM I N MAC
------
020 289 24 a C ~OR OUT80ARD p 026 2.89 25 <> CRIT !:>.
032 .289 26 LEADINl> I:OGE REGION 0 512 020 23 • MACH 026 Il C P M1N OLI ______ ~ ______ _L ______ _L ______ _L ______ _L ______ _L ______ ~ ______ ~ ______ ~ ______ _l o 10 20 30 40 50 60 70 80 90 100 Tl
-=)
FIGURE 81 EFFECTS OF MACH AND REYNOLDS NUMBER ON CRUISE WING WITH NACELLES AND PYLON ATTACHED
)
i '-....../
u
'--'" A ....---~ SYM I RU'" , 'H
OFf 23 --1
CONFIGURATIDr, S,NU,"2AZIA V lAH\~ ,¢
0° UJ
MACH' 020 I _&0 3~ R ' (; 1;1 • 10° 1_10 36 ...
"MA:
o
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t
t '., ,) .. " ..
A. LIFT AND PITCHING MOMENT FIGURE 82 TAIL·ON CHARACTERISTICS FOR THE CRUISE WING WITH NACELLES, PYLONS, AJ',;D STRAKES ATTACHED
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~ (,-V"· ,- (; 3'" l_ ':1. ,I .;~ :-- ': .;"-,..:1: ~,' B DRAG COEFFICIENT FIGURE 8:~, TAIL,ON CHARACTERISTICS FOR THE CRUISE WING WITH NACELLES, PYLONS, AND STRAKES ATTACHED (CONCLUDED) ...
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-..J SYM 'H RUN -1
I RN • 10'
:, MAC CONFIGURA nON S, N2A? 2A ZIA V IAIlIA 0° 34 5.12 I I MACH - 0.20
- 0° 37
1 14 I
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r~ rrt El ., =\-- -<.(1) o c: .', .., 1"- 0 ~ l·· ~ .c:; v fr" :>c m ..,._ rl • . " A LIFT AND PITCHING MOMENT FIGURE 83 EFFEC, T OF REYNOLDS NUMBER ON CRUISE WING WITH NACELLES, PYLONS, STRAKES, AND HORIZONTAL TAIL ATTACHED ... , \
I
'-- v
~ ". ~ ... , ~ 4 •• SYM RUN 'H R x 10' N MAC 5.12 0° 34 ., , - 1.14 0° - ..
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:!:- o j~ ~:~ 3~ J; 1 [l DRAG FIGURE 83. EFFECT OF REYNOLDS NUMBER ON CRUISE WING WITH NACELLES, PYLONS, STRAKES, AND HORIZONTAL TAIL ATTACHED (CONTINUEDI \
.)
\ \ ~ .,
u
--..-/
"
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SYM RUN IH x 16' R I N I MAC \ \ 0° 5.12 ..
0° 1 14
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d .? ' .. 1r· .':: I - 1"-1. po,.
C LIFT/DRAG RATIO FIGURE 83. EFFECT OF REYNOLDS NUMBER ON CRUISE WING WITH NACELLES, PYLONS, STRAKES, AND HORIZONTAL TAIL ATTACHED (CONCLUDED) , ,-",
~
· ...
)
condition. The LID at 0.69 CLMAX was reduced from 16.8 to 13.6 by the
decrease in Reynolds number.
In summary, the basic cruise wing achieved a high level of CI.f.!Ax (1.513) and
LID at 1.2 Vs (19.45). Addition of the nacelles, pylons, and strakes
resulted in a neglible change in CLf.wc, a reduced L/D (17.7), and improved
pitch characteristics at high angles of attack. Increasing the Reynoldls number from atmospheric to the high Reynolds number test condition increased
-
the CLlWC significantly (~CLMAX = 0.39). A Mach number increase from 0.20
to 0.32 resulted in a decrease in Ctr.wc of 0.11. Test data for the cruise wing configuration with the horizontal tail indicated tbe low-speed pitch characteristics require improvement. It should be noted that the ongoing high-speed wing developnent for high as};ect ratio supercritical wings has, in fact, altered the SP:ln loading which should inprove the leM-speed stalling
---
behavior.
VCR Configuration
,)
Both the VCK and slat were evalt.ated during the test program. 'Ibe VCr< was evaluated first with nominal two-segment landing and takeoff flap deflections. Before the VCK position was optimized an extension to the spoiler trailing edge in the region of the flaperon was required at the landing flap deflection. '!be existing spoiler length, and the resulting gap and overhang (O.H.) in this region, resulted in large flap sepuation. Only the most conservative flap grid position (large positive O.H. and small gap)
\-laS unseparated. '!be basic spoiler was extended at Xw of 36.362 em (14.3178
in) by 0.29 an (0.11 in). 'Ibis trailing edge extension was &creased to 0.0
an at Xw of 43.411 cm (17.091 in) and Xw of 14.800 em (5.828 in). This
resulted in a m~re positive O.H. of 1.29 percent at Xw of 36.362 an (14.3178
in). 'Jlle gap was also changed and, for flap position FlAM a reduction of 0.3 percent resulted at the saI:le wing station. '1be remain&r of the test was conducted wit.'1 this revised inboard spoiler configuration (flA,f2A).
Figure 84 illustrates the VCr< with two-segment flap configuration installed in the liles l2-Foot Pressure Wind Tunnel. Before proceeding with the VCK optimization the effects due to the removal of the VCr< filler blocks \Yere evaluated. '!be results indicated minor changes in lift, drag, and pitching moment. The remainder of the VCK runs \-Jere accanplished with the VCK filler
)
blocks removed (see Figure 85).
..,
.-
~ . .,
--
CRi~;!'~ -,,_ P"'~: t':
)
c r- ;'4~r.,r; Q'Jrl!.JT·{ FIGURE 84. VCK AND TWO·SEGMENT FLAP CONFIGURATION IN THE AMES 12·FOOT PRESSURE WIND TUNNEL
J
)
FIGURE 85. LOWER SURFACE MIDSPAN REGION OF THE VCK WITH TWO·SEGMENT FLAP CONFIGURATION
-
..
, ..
')
I' 7he 1nfluence of mini-tuf~s was also examined for the high 11ft confiquration. As ... ,ith the cruise \"O'1ng configuration, the data indicated very small effects on lift and pitching moment ann a slight rpduction in drag (~CD == 0.002). '!he change in high lift characteristics due to nacelle strakes indicated an increase in CLMAX of 0.04, slightly more positive ~ pitchinq moment, and a negligible change in drag.
VCR Landing and Takeoff Ot.?timization. - A VCR deflect jon and position sl1rvpy \'laS evaluated with a nominal tHo-regrrent landing flap configuration. Flqure Hli illustrate!' the VCK posj hon study for ')5 VCR deflection (horizontal tail-off). SigOlficant renuctlon in CLMA.,{ and positive pitch increment 1 S
shewn for the ,)5A/55A grid position (Gap = 3.5%, O.B. = -1%). Piqure fl7
presents the variation of section lift coefficient for trer,e positions. Par run 49 anc1 ')0, the gO-percent station stalls first. Run '10 indicates f.
rapid lift losfi at the 72 .5-~rcent station, and run 51, adverre effect5 due tn the large gap for both 72.5- and 90-r-ercent st.ations Clre lIldicateci.
~
Figure 88 presents the posltion stuny for a VCR deflection of 45 deqrees
acros~ the span. Run,2 (Gap =- 1.5%), 0.11. = -J!\:) has a Hlqnificant
reduction 1n c.I.r11\X. Runs 61i and 1i7 have differt=mtial V(,K qnns (rlHferent: inboard ann outboarn POSIl • ,ns). 'The larqest CL~lAX \''US ohtained in Hun 67
(CLt1l\X = 3.15). '!he pitching manent characteristics lndicate signif icant
variations due to the VCR position. Figure 89 prerents the rection lift characteristics for the position survey. nun 52 indicates that for small
gap and overhanq (Gap = 1.5%, O.H. = -1%) the 90- and 72.5-~rcent SP:lD
stations achieved a large Cr.t1l\.X and qentle stall. The inboard statjonr;, however, reachen their Cr.r-1AX values at very low anqles of attack indica tinq premature stall inboard, ann cnnrequent loss in CLM1\X. Run,l «('ap = :-!.'%, O.H. = -1%) indicates improved characteristics lnboard, but increased lift loss at thp. c)O-percent span station. The 1 i ft values for r.un ,4 Sho\l a reduction in inhoarn lift loss, and a sharp lift loss at 90 percent. 'l'hp 90-percent span station is shown to have a gentle stall for Run ,6 (r-an"- 2.5%, o.H • .::. 0%). Run 66 sectional lift values indicate increased lift lOfls for the 50-~rcent station.
)
wr_ - ~
u
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f
...
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SVM RUN CONFIGURATION S4G'A ~l "- r.'ACH .. 0.20 (> ~ ~o-: ~) R = S.12x 10' "';~I N MAC :.
'r, ..
~ '"' 2 ....
.J j:~ .... H
.-
"E- l
:;00 (.1 '; " I) : 0 c " .. -1 .... -,
e
-.' , .. ,.,.
------r-- El -- --- T -------,-- ~ l' I~ 20 2S ]0 -; DO-j "';0 "tIffi J 0:0;
1 e
o O~ ;0 I·- U[J ~ o ~ " -(' <'0,· i:)-O .
o CJ ..
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o
a -t t~" •
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I
:1 u I I o 6!
._l b - 25M/12C F
I
Ii: r- -V(oiJf-- ~ ~ :' INBOARD OUTBOARD I RUN LE CONFIG LE GRID GAP O.H GAP O.H I ~l". .1 I t.
r---' -("I t)1) 1 49 VCK 5Se/55C 2.5 -1 25 -I VCK 550/550 25 0 25 0
j
VCK 55A/55A 35 -1 35 -1
- - -
A LIFT AND PITCHING MOMENT FIGURE 86 VCI( POSITION STUDY FOR LANDING FLAPS (., VCK " 55°) .1 (030n10NOO) loSS ~:)I\,}) Sd'l11:1 ~NION'I11 1:10:1 AOnlS NOillSOd >lOA '9831:1nDl:l O\fI:lO a ,.'
:: ~, - l'
\ \ I ,
i
. .., - .. ~ (.
Ii) 6~ " Nn!! WAS
•
•
y
')
ORl~!;~;) \.~ ~ Of POC:1 QiJ .... _ ..• 4.0 CONFIGURATION S4G1A MACH = 020 R = S 12 x 10 NMAC of = 2SMI12C 3.6 b VCK ~ SSC/SSC GAP a 2 S% o H = -10% 3.2 2S 2.4 c ,
J
2!J
1.6. /. 0'
c-
.
1.2 l-
/
/ os l-
:00 sao 72S 04 •
uo '~----~~----~------~-------L ______ -L ______ ~
a 4 a 12 "'FRP tDEGREESI
) A. 0VCK " sse/ssc
FIGURE 87 VARIATION OF SECTION LIFT COEFFICIENT FOR THE VCK WITH TWO-SEGMENT LANDING FLAPS
-
-j
ORlGi~':.-i:' ~:_ •. _~ OF. POOR QU:"L:-I-: CONFIGURATION 5 G 1A MACH ~ 020 R .. 512" 10& NMAC b • 25M/12C F liVCK ~ 550/550 GAP ~ 25% o H s 00% -' c~
~ 20
1 2 RUN 50 SYM 1) 0 200 305· L:::. 500 "V ~ ·VCK PRE~~URE PROBLEM
00 ;:------~------~;-----~:-------JL-------L-------J
..
20 24 "FRP (DEGREES) B. 6 • 550/550 VCK
)
FIGURE 87_ VARIATION OF SECTION LIFT COEFFICIENT FOR THE VCK WITH TWO-SEGMENT LANDING FLAPS (CONTINUED I 't - .:.;
)
~:- ... :--
.Jr ~;"":~.~ f..:_. _.:, CONFIGURATION ~~G1A MACH 020 H~IMAC - 512)(10" ~F • 25M/12C 3b U VCK • 55A/55A GAP 35,<> o H - 10'" :!'l .!.t
)
L , !O
~
// ~-"
/0 ,/ ~J
~
I l
- -fiUr0T-i
--...,.--- ---., SYM' I --~
o 200
IJ'J
o l05-
l::;. SOO
o 12"
~Laoo I).J
-' .. I."" P~t-_':;''1ll1l''" ";~t:t-t~ _fl. .. 1
"I ~I------~------~------~------L--- ____ L- ____ -J
o !l 4 .~.w tl)';t..REESI
)
C b a 55A/55A VCK FIGURE 87 VARIATION OF SECTION LIFT COEFFICIENT FOR THE VCK WITH TWO-SEGMENT LANDING FLAPS (CONCLUDED) • 1-14 " \ ( ", \":
()
"------
" . ---./
, , , I, RUN 'vi t:~ 'i ~~- SV\\
i'
L :.
CONFIGURATION Sol G 52 IA I.
" ~ , , MACH· 020 , I c, !i6 R • 512 x 10' fJ I N MAC ~ 66 :1 1" / V 67 / .: • V\.
J , !ii{§6,!,L.: _ v .1 ... 1 ..... ' "" Ii!' ~ r \ , '" ... 1,1 Cl.-. • ;J I!! • '" '" tfj \ go::' 000 at!
., ':' 0 ~
\-
r, C C ~ t i I· @I ,- ,~ " ; " ?J ,", J~ , , \ , .'.0";- 00 z 'TI:u ~ ~ :' I' "QC5 • F • 25M/12C .'
0-
"
I I I, h' ;o~ " , .to , INBOARD OUTBOARD U1 (,0 ;:. ~ '\ RUN LE CONFIG LE GRID GAP a H GAP O.H.
g~ flr,,:'_t or HJlHL'.>DlC:: i ~ l> (:,
o
52 VCK 45F/45F 1.5 -1 1.5 -1 1 '
, R r- "'1
53 VCK 45G/45G 25 -1 25 -1 "AA8~l\
,~cn
54 VCK 45E/45E 35 I -1 35 -1
Vv
08 g
i 56 I VCK 4.5H/4.5H 2.5 -1 2.5 -1 I
o V vV
, I
t- 66 VCK 45E/45F 3.5 -1 15 -1 J' O ... ~ o~ooO ""O:'t..
o 67 VCK _ _45E/45G _~.!;_ 1_ -1 _25 -1 ... ", 1 Q 0 i'i vL'Zi I'iI CJ ': I
81 e V t' -
~ ~ °0 I , \
e $
V ij,..l - ..
., •
] ;10 t ~ 111 • r R\~L ~l ... v L V _I..
A LIFT AND PITCHING MOMENT FIGURE 88. VCK POSITION STUDY FOR LANDING FLAPS (~ ;: 45°, VCK I ,
/
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i I , , , \ I , 'I ,
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I' 1'1 'I /r ii, :1 ' "
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/930nl:lNO:l1 foSt> '>I:J/\'/l SdV'l~ DNIONV'l !:IO~ AonlS NOl.LlSOd >lOA SS 3!:1nDI~ I ' '" 1 ''1 I i , , : ...
,:
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£9 'I.
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9S II "
I : loS \ 1:9 ".
I Z9 "'1',- 1'-, NnU WAS , <.
\ !
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)
en::::. : ~:~
CONFIGURATION S G OF POCR ~:';Ai..i·;~( 4 lA • MACH 020 R 512 x 10& NMAC hI' ~ 25M/12C b - 45F/45F VCK GAP 15% 36 o H - -1.0% 28-- 2.4
)
"l 2 0 h _ I // /
-o-o-o-o-~-o.-a
-...........--~
/
l
RUN 5,Z SYM " 1)8~ 0 200 .2 o.tL 00' I I o A n 20 24
')
FIGURE 89 VARIATION OF SECTION LIFT COEFFICIENT FOR THE VCK WITH TWO-SEGMENT LANDING FLAPS -,
•
-~
)
CONFIGURATION S4G1A MACH .• 020 RN ~ 512" 10 MAC of - 25M/12C
° VCK = 45G/45G
3.6 GAP - l5% OH -10%
)
\~
1 tj
~)
RlIrj bl
u8 .......,... ~~
1-0 200 6. 500
lU~5
o 725
" 900 \,-:
I I
00 n - 12 16 20 24 • ... RP lDEGREESI B 0VCK - 450/450
)
FIGURE 89. VARIATION OF SECTION LIFT COEFFIr.IENT FOR THE VCK WITH TWO.SEGMENT LANDING FLAPS ICONTINUED) '- -- ' .......
"
)
O~;L:: ~ " . : OF POQ~ ~.~ ...... ~ ~ : CONFIGURATION S4GIA MACH 020 R • 512 x 10D NMAC bf' 25MI12C hVCK . 45E/45E GAP = 35% OH ~ 10%
J
c ,
"~
1 2 HUN !>4 SYM " OIl 0 200 725 : 'V 900 I 00' I I I o 4 R '" _ "'FRP (DEGREES' C 6 • 45E/45E VCK FIGURE 89. VARIATION OF SECTION LIFT COEFFICIENT FOR THE VCK WITH TWO.SEGMENT
)
LANDING FLAPS (CONTINUED)
•
-
.-, \,"11
)
CONFIGURATION S4G1A MACH - 020 R • 512x lOb NMAC b - 25M/12C F I>VCK " 4SH/4SH GAP 25% o H ~ 00% ~8
J
c~
"~/
I d f.!IJN 56 SYM ~ 0 20 a ~ SOO 'V 900 00 I I I I I o 4 8 12 16 ~n - •• ~ RP 'DEGREES) o 0VCK 4SI~/4SH
)
FIGURE 89. VARIAT • .JN OF S ';TlON LIFT COEFFICIENT/FOR THE VCK WITH TWO-SEGMENT LANDING FLAPS ICONTINUED) ,50 "".-.
~ ....
)
... ·":·· >. ", t\.,_ .. '. ..- .
OF PC0;" C.J"_.
CONFIGURATION ~/'IA MACH 020 H 512 x 10 NMAC h 25M/12( f 0VCI<. 45E/45F INBOARD OUTBOARD GAP 3 5% GAP· 1 50/..
OH . -10% Ot!
)
C l RUN 66 SYM I')
o 200
o 305 ~ 500
o 725
'V 500 00 I " I o 4 8 12 '''' ... n "FHP'DEGREESI E 6 a 45E/45F VCK
)
FIGURE 89. VARIATION OF SECTION LIFT COEFFICIENT FOR THE VCK WITH TWO"sr:GMENT LANDING FLAPS (CONCLUDED)
•
)
[-'lqure 90 presents mini-tuft photos for Run (,7. Incrcaninq the "lnqk (;i attack past Cl'CLr1AX initiates stall on the inboard \-1inq out t.o t'he area be hind the nace lle/pylon. Saoo sep:lration is also eVldent LOL the trallln'1 c>dge region of the flaperon. Some trailing edge actIvity 1S noted at
(HI-percent span at Q'FRP = 21.14°. As higher angles uf attack are reached
the sep;lration continues over the inner region of the main wing, and dt. the hIghest angle of attack, Increased tuft activity is noted at the wmg t.ip (near the most outboard VCR support hracket). Figure 91 present!:J the corresponding pressure distributions. ('Proml values of -11.2~ c.re reached on the VCI< for the outboa rd teqlOn of the winq. l:'or the luqt,(::st. angle of attack, the presslIre dintribution at c)O-percent span statlon indicates a si~nf1rant reduction in r.l and ti-.e resultinq positlue pitch incren'ent in shoIJIn 111 I-'iqure RR. F.xamina tion of the section 1 Ht characteristics in FIqure 92 also IndIcates thIS crend. The sectinn lift characterlstir.s presented pre'Jlously in l~lCJIH~ fll)/\ (G,1~' -= 1.')% and ,l.H. : g~', 1nclicatE's th3t improved c!1..lract(~rist1c;s CI~llld I"~_ 0iltc1ineo \-71th th1!1 grl,1 I'Psiticn for the outboard portIon of the \llf,:,,).
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l"lqures 1)3 thL0 VJil '19 PtPflt1~t .) "un1-,r ,~Il;=('tLon dnll nO:"tJI i: -:tlldv for a n rwr.:ir:.'ll hlO--s;::'jm::llt tc I--·,·t t i Lhl' if>ll:..'.;tt:-.J1_ In t. ( C"l' I r.: 4S I!C!'
cleflectlons ,,(t_ supenol ., I,E: 'i'-," :letlCt.:tlOilb inrlllH .. ,lZ1L, It.e liit. h)ss f: r tJ'e nlllfipall ann outboara Sectl·"n. Tt-,£> :-;tal.J 'JlollrE:S~'l('l: ~)ho ... m b~1 /tqllr.= 97 t for the 45E/45G W'K qnoi positi(\l1, LS Slrf11l:ir t,) t.hat S!10\·m f:,)r Lit"! landing flap c1eflection. l\gall\, at tb~ tuqhf'st anqlE: of attar.l~ the )I)-percent span Btation lift 10s5 c~slllLs tn reducecl stal-·iJ Lt'! f0r the r.~nfiquratl(n (Flqure 95).
Due L" npcr'ilIUC;)1 ("..Jrl:~1e):)t:". :: (liT. I'.()slhr,n \'CK 1S required lC'lI1f: deflected r·' .... -ltl'.ni. lompart:'on ot ti·(! result,; f'-lr thE: lundinq and tal~e()ff fIal-
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~pttln'lS lWllC':jt-r! Ihat VCR 'lrld poslt:lon 45F.'45G \J~H; the beGt comnrattH',f from the stan('1pcHnt of CLH1\Xt Lin i'lt 1.3 "s' and pitc~' :~haractenstlcs.
1\ l ~o n ... )ted pre'lI o1]sl y, i SI:,Clll er 'lcIP for tre VI.K 1m t I.e ;'Ilthoa rd regl<'m .)f rto('! \!inq "/(")111tl ~1"0 unproved tl,p hlqh anole of attack latch characterlstics.
<lht~ ~ff~ct ()~ H:duced VCl! tlcf1~Gtl\)n (~:;o 'JerSlIS 4')0) also indIcates a poSSIble perfo~"lnce 1ncrease for .:i fllrther reduction in ITCK r1eflection.
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for 2c)o of maln flap deflection the 15K/12C configuration was selected as the best configuration from the standpoint of CLl.tAX, L/D c.t 1.1 'Ill' and pitching Tiloment charactenstics. In FiC')ure 102, for a flap deflection of 0 0 15R/1~C, a reduction of flaperon deflection from 35 to 25 resulteo 1n a small decrease in CLr.1J\X, similar pitch characteristics, and rlecreased oraq.
Th1S ronEiquration Hould rrechanically be very canplex. 35R/12C was selected as the r.est ccrnpranise for this rein flap deflection.
Analysis of the data for the main antI aft flap survey for the IS-degree main flap (FIgures 104 and 1 05) rl'~ultc:d In the selection of ISH/lOP. ror f11le
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degrees main flap rteflection, thp selected configurat10n was 5(,/IOB (sec Fiqure 106 ann 107).
YCK Sgan and rlacelle/l'~~'~ffur.ts. - '{'I.e effec:t 5 l1f HCK span cnrl nacelle/pylon for landinq anel takeof( l\lO-seC]ll'ent flap deflections ".:[("' shm'fl ln F1gures lOR and 109. The data Indiccltes that d CLrw: of ~. 196 \'as
obtaired fot the full spm va (nacelles and pylons off). ll.dditlon of the
nacelles, pylons, and strakes, for the landinC') flap reflection (Fiqurc 10[1) results in a positive increrrent in pItching marent and a small [eduction in
CLr1AXo The pitch trpn\1s at ar.CJles of attack greater than n degrees art'
s1milar, but at a d1tferent leuel. Removal of the over-the-p'.'lon-·Vrr( ,'!xtens10n (L()!\) re~ulterl in a reduction in Cr. 1l\Y. and lmprOW!O hiqh anqle of r ilttack pltcr ch..'lIacteristicR. Remov1nC') the fuselaqe seal VCl~ extens10n rT.'1AI '1iE'loen a 51 iqht rerluction 1n CLr-1AX and a fdrther improveMent in the bas1'; p1tch chd[actensttcs. Other than the drag increase due to the nacelle/pylon and strakes the effect of VCl{ span \las stni\l1 (maxiT"llrl
rlifference tn Lin at 1.3 Vs bemq 0.14 in 1,/0). r.. S1mllar compar ison for
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Reynolds Number Effects on VCK with Two-Segment Flap (TaU-0Hl. - The influence of Reynolds number variation on the longitudinal aerodynamic characteristics with two-segment landing and takeoff flaps is shown in Fiqure 110 and 111. A sizable shift in pitching moment is indicated for the landing and takeoff flap configurations at the lowest Reynolds number condition. The basic trend of the pitching marent at angles of attack near stall is not significantly changed by the reduction in Reynolds nt.mlber. A reduction in CLr.W{ of 0.16 for the flap deflections is also shown in Figures 110 and 111.
VCR with 'l'wo=Segm:mt Flaps and Horizojltal tail. - Figures 112 to 114 present the characteristics of the VCK wi th two-segment landing and takeoff flaps with the horizontal tail, and the effect of Reynolds number.
Characteristics shown in Figure 112 indicate nose down pitching moment for the -150 and -50 horizontal incidences near CL~1AX. Post stall pitch
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characteristics a re again influenced by the tip seraration, and a reduction in stability is indicated at the highest angles.
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basic character of the pitching marent trends.
Single-Slot Flap Optimization. - Results of the optimiza tion of this trailing e~ge device were presenten in Figures 115 through lIB for flap deflections of 35 0 , 25 0 , 150, and 50. The pitch trends for each flap deflection were similar for the various grid positions. As the flap deflection was reduced, more favorable pitch charasteristics after CLr1AX were obtained. Positions selected from this study were: 25K/O, l5G/O, and 5A/D (see tata SllJ'm'ary Section ••• for comparisons ,lith two-segment flap system) • Inboard VCR 8ffects. - The inboard VCR was removed and the ch~nges in the longitudinal aerodynamic characteristics G;re shown in Figure 119. A CLr1AX reduction of n.46 and an abrupt nose down pitching moment after CLr~ 1S
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Reynolds Number Effects on VCR with Single-Slot Flap (Tail-0ff). - The Reynolds number effect on VCK ~ith single-slo. landing and takeoff deflections is presented in Figures 120 and 121. A CL~1AX reduction of 0.16 tVc1S also obtain:!d for the single slot configuration. The pitching moment shift is similar to the two-segIrent flaps.
VCR "lith Single-Slot Flap and Horizontal Tail. - The effect of horizontal tail deflection for landing and takeoff single-slot flap deflections is sh~v.n in Figures 122 and 123. The landing flaps pitching moment tr.ends indica te some reduction in neqative pitching moment prior to stall for il{ ;0: no. Takeoff flaps pitch characteristics are improved.
Slat Configuration The slat with two-segment landing flaps configuration is shown in Figure 124. l\s nention:!d previously, this configuration was evaluated with the optimized flap system defined in the previous lTCT, studies.
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Slat Landing and Takeoff Optimization. - A slat deflection and poSition survey \-/as evaluated \-/ith the two-segment flap system. Figure 125 presents 0 0 the slat 25 /35 position study. A CLrtAX of 3.00R was obtained. The basic- tail-off pitching moment trends indicate positive pitch increments prior to CLMAX' and no significant ..-fe'oct due to the change il"l slat position. The draq values show a slight reduction in drag for the small gap and overhang
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0 0 The effect of slat position for the 15 /25 slat deflection lG sh~m 1n Figure 127. The CLl1AX increased to 3.2, but the pitch characteristics are
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Results are similar to the 25°/35 slat results. Sane improvement is noted o 10 the sectional lift variation at 90-percent span station lo/ith the 7S outboard slat deflection. Figure 129 presents the chordwise pressure
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ORt~L _vEf;-:-:(N' I "-
----------------
B DRAG FIGURE 121 EFFECT OF REYNOLDS NUMBER ON AERODYNAMIC CHARACTERISTICS OF THE VCK WITH SINGLE SLOT ,'AKEOFF FLAPS CONFIGURATION (CONTINUED) , .,..
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--------- - ---------- ----
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t
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-- --- ----- - - ---- -- C LIFT/DRAG RATIO FIGURE' 121. EFFECT OF REYNOLDS NUMBER ON AERODYNAMIC CHARACTERISTICS OF TilE VCK WITH SINGLE·SLOT TAKEOFF FLAPS CONFIGURATION (CONCLUDED) 'Ill
~ u
'-./ MODEL L B - '18b R ..
G <> 'Z CONFIGURATION S4 IA HIA VIA W SYM RUN 'H t: 0 '1001 MACH·020
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-0 700-" _ _ _____ __ _J A LIFT AND PITCHING MOMENT FIGURE 122. EFFECT OF HORIZONTAL TAIL DEFLECTION ON AERODYNAMIC CHARACTERISTICS OF THE VCK WITH SINGLE-SLOT LANDING FLAPS CONFIGURATION ~ .. , \ -......./
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B DRAG FIGURE 122. EFFECT OF HORIZONTAL TAIL DEFLECTION ON AERODVNAMIC CHARACTERISTICS OF THE VCK WITH SINGLE·SLOT LANDING FLAPS CONFIGURATION (CONCLUDED) ....
~ ..
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MODE L. L B -- Lt8tJ R
SVM RUN IH ...
z BASIC CONFIGURATION S4 H IA VIA :::: 0 'i0:>- w OFF 105 MACH = 020 o ~ I 0° 108 W I R = 512 x 10' N _5° MAC 11 0300';
3 ~Ol
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....
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ORIGINAL PAGE IS OF POOR QUALITY
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--- FIGURE 124. SLAT WITH TWO·SEGMENT LANDING FLAP CONFIGURATION
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SVM RUN I MODEL L B-'-i8b R ....
z ') CONFIGURATION S~ G'A w ,
t: ° '100l
[0 MACH L 020 ..
RN ~ 512)( 10' :' I
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ANGLE Of ATTACK-DEG .... 2 00-1 -n ~J z III "'0 [1 o m -0 1001 c:: €I
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C • iii • I ANGLE Of ATTACK-DEG
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o -------------.----~ A LIFT AND PITCHING MOMENT FlllURE 125. SLAT POSITION STUDY FOR LANDING FLAPS (b = 25°/35°) SLAT .....
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1 2 J-t:l"" U RUN 150 ----,- SYM :>00 0 305
08 rr
6 'lOll '\l 900 04 I OO'~------~ ____ ~ ______ ~ ______ ~ ______ -L ______ ~ o q a 12 16 20 "I- III' (DEGREES.
A /, SLA _ .. a 2SAl3SA
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FIGURE 126. VARIATION OF SECTION lIi=T COEFFICIENT FOR THE SLAT WITH TWO-SEGMENT LANDING FLAPS /, ~ .:: ;y: !
(+"
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G~:2:" ·.~,L ~". ~ : ~ ': OF POOR QU.';Lrr( 40.-- CONFIGURATION Ss G 1A MACH ~ 020 R ~ 512 x lOb NMAC of = 25K/12C
36 I- hSLAT a 258/358
GAP - 1.5% o H. ~ -10'0 .!.4
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8. "SLAT a 258/35B FIGURE 126. VARIATION OF SECTION LIFT COEFFICIENT FOR THE SLAT WITH TWO.SEGMENT LANDING FLAPS (CONCLIJDED) ~ ~,
u
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z CONFIGURATION S5GIA
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LE CONFIG LE GRID
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L ______________ .. _________________ . - ......
____ J
A. LIFT AND PITCHING MOMENT
FIGURE 121. SLAT POSITION STUDY FOR LANDING FLAPS (oSLAT = 15°/25°)
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Or~;~" ~: lL '?: OF FOOIl Q...;. \:...: r'j CONFIGURATION SSG 1A MACH; 020 AN z 512 x 10 MAC b ; 25K/12C F 5 = 15E/25E SLAT GAP = 15% o H = -1.0%
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C~ 0-0-<>-0-0-0 RUN 155 SYM 'I I 0 305 C::.
0 725
'V 900 00 I~----~------~----~-------L---- __ ~ ____ -J o 4 8 "FRP tDEGREESI
J
FIGURE 128. VARIATION OF SECTION LIFT COEFFICIENT FOR THE SLAT WITH lWO.SEGMENT LANDING FLAPS (OSLAT '" 15E/25E)
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'FRI' • 1614 OEG A FIGURE 129. EXPERIMENTAL CHORDWISE PRESSURE DISTRIBUTION FOR SLAT WITH TWO·SEGMENT LANDING FLAPS (03nNUNO:» SdVl:1 ~N'ONVl lN3W~3S"OMl HllM lV1S l:JO:l NOllIl81l:J1SJa 3l:JnSS3l:Jd 3SIMOl:JOH:> lV1N3WIl:J3dX3 "6lL 3l:Jnm:f ~II' J. .. "t.olf1.4 ,
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II "£RC("if Cr<Rl "[lter"l CH~ C "FRP ~ 2119DEG \ FIGURE 129 EXPERIMENTAL CHORDWISE PRESSURE DISTRIBUTION FOR SLAT WITH TWO·SEGMENT LANDING FLAPS (CONTINUED) \ ..
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• ~ • • c ~ 1» I. • » • 10 ~ * 1m ItO ~ M p[RCt.r CIi(J<IJ PCllctNi C~CJ.'D o "FRP • 2309 DEG I' FIGURE 129. EXPERIMENTAL CHOROWISE PRESSURE DISTRIBUTION FOR SLAT WITH ·rwO·SEGMENT LANDING FLAPS (CONCLUDED) . , ..
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A r.PMIN value of -15.4 is shown for the 50-percent span station (Figure 1290). This station rema1ned attached for the angles of attack 1nvestigated. Also shown in Figure 1290 are the attached flolll at the 20-percent span station and the reduced lift at the 72.5-Fercent span station. Figure 130 shm'is the section lift variation with angle of attack for Run 154. The relatively constant lift at the highei: angles of attack for the 20- and 30-percent sran stations is readily app:lrent. Canpuison of the 72.5-percent sp3.n station sectional lift characteristics for 25A/35A and 15D/25D (Figures 126A and 130), 1ndicate a reduction 1n ma:dmum lift and lift loss for the larger slat deflection.
Figure 131 presents the results of the slat position survey for tbe t\lo-seqrcent takeoff flap configuration. The largest CLl1J\X \-JaS obta ined ,-Ii th
posit10n 15D/250 (Gap = 2.25%, O.H. = 2%). Significant reduction in CLrlAX
and increase in drag is shown for the large gap ancl overhang position (Run 174). ~e pitching moment trenns are similar for the various grid positions.
)
The sectional lift values are sha,.ln in Figure 132. Reducen sect.ional ] ift values are indica ten for the large gap ancl overhang poSition (run 174) as expected. The small gap ,ncl ovp.rhang (JOs1tion (Run 175) results in the best sectional 11ft. variation at ')1 -~rrent sran station. Figure 133 illustrates the chordw1~ pressure distribution for Run 173 ((;ap = ~.25%, ().I~. ..: -~%).
Note\-lorthy 1:" the increase in loading for the 20-percent sran station at
--------
angles of attack qreater than QCLMAX' ~xamination of min1-tuft photographs for the lnboarn w1nq indicatecl significantly reduced tuft activity canp;lred to the optimizecl VCK configuration. The large chord, and qreater sparn-lise extent of the inboard slat (see Figures 4Q and 50) resulted in siqnificantly less lift loss at high angles of attack for this region of the w1ng. ~his is confirmed by the trends shown in Figure 134 for the 20-percent Sp:)n station. Sane reduction in the 1 ift values at high anqles of attack fOl the 90-percent span station 1S also noted. This \mulel indicate, in a siMilar fashion to the landing slat and the VCK sturlies, that thp. rp.duced gap and overhang position is best for the most outboard region of the \-linq. Sane reduction 1n lift at the 30-percent sp3.n station is noted at the hiqher angles of attack as well as a substantial reduction lr. 1.ift for the 72.'5-percent span station. Comparison of the cruise \-ling, slat
)
extendecl flaps retracted (clean trailing edge configuration), and the
l6)
) -, loA} ~)
~
ORIGINAL PAG~ ;8 OF POOR QUALITY CONFIGURATION SSG 1A MACH = 0.20 RN = 5.12 x 10 MAC 6 - 25K/12C F 6 = 150/250 SLAT GAP·225% o H. = 200% 2.8
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FIGURE 130. VARIATION OF SECTION LIFT COEFFICIENT FOR THE SLAT WITH lWO-SEGMENT LANDING FLAPS (5 :: 1501250) SLAT
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A LIFT AND PITCHING MOMENT FIGURE 131. SLAT POSITION STUDY FOR TAKEOFF FLAPS (b 15°/25°) SLAT !..
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r, 20 0 08 ~/ #' 72.S u 4 \ ) 'F FlP C DEGREESI A. ~ SLAT ~ lSF/2SF FIGURE 132. VARIATION OF SECTION LIFT COEFFICIENT FOR THE SLAT WITH TWO.SEGMENT TAKEOFF FLAPS ..
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FIGURE 132. VARIATION OF SECTION LIFT COEFFICIENT FOR THE SLAT WITH TWO.SEGMENT TAKEOFF FLAPS (CONCLUDED)
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I;R:G!!!.~t P."t~~ :- OF POC~ QUALti)' CONFIGURATION 55 MACH ~ 020 RN ~ 5.12" 10 MAC SF - 5CI10B 3.6 6 • 150/250 SLAT GAp· 225% o H. - 2 OOo~
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"V 900' 'SLAT PRI:.SSURE PRllB,_I:.M 00 LI ______ ~ ________ ~ ______ ~ ______ ~ ______ ~ ______ ~ ______ ~ o 4 a 12 16 20 24 28 'FRP 10EGREES)
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FIGURE 134 VARIATION OF SECTION LIFT COEFFICIENT FOR THE SLAT WITH TWO-SEGMENT TAKEOFF FLAPS (li - 150/250) 5LAT
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)
takeoff configurations just discussed, indicated significant reduction in LID characteristics with slat deflected at the landing position. Improved LID perforrrence \'lOuld appear possible with a sealed-slat configuration (slat trailing edge sealed at the truss surface) and this configuration should be evaluated in a future te~t program.
Hach NUmbe r Effects on Slat with Two-&!mrent Flap. - Figure 135 presents the influence of increasing l-1ach number on the lift, pitching moment:, d:':lg, and r~/D characteristics. The CL~,1AX reduction of 0.23 is indicatf:d for an increase in ~'lach number from 0.20 to 0.32. The most significant diHer.;:~ces in lift occurred at the angles of attack after CL~mx was achiEved.
Increasing the ~Ech number also resulted in slightly larger lift values prior to CLrtA.x. Figure 13') also indicates that reduced stability occurs ac the 10\'1er angle of attack for CLr-1AX as the l1:lch number is increased. S:nall differences in drag and LID are also shown in Figure 135. ]\nalysis o~ +:he pressure data lnc'licated local regions of supeL"critical flow on the outboard -) slat for 0.26 and 0.32 f1:lch numbers.
]\ similar comparison is shO\"n in Figure 136 for the slot with ulc-sec:;::lent takeoff flLlP configuration. The changes in lift and pitching Moment due to increasing tEch number are siITIilClr to those for the landing config.lration.· A red,--,ction in CLr1AX of 0.11 is shO\-/n for an increase in Mdch number from O.~~ to 0.32. A slight reduction in drag, at a qiven lift coefficient, is also sha.ln in Figure 136 and this res~lts in slightly larger values of LID.
Reynolds Number Bffects on Slat \'lith 'l'.i/o-l)eq::ent Flap. - The influence of a reduction in Heynolds number frCl'l the nminal test condition for the slat I·lith t\olo-segment landing flaps configuration is shown in Figure 137. CLr~.x reductions of 0.09 and 0.29 are shO'tm for the intemediate and 10\" Reynolds 6 6 ,
number test conditions (RNMAC = 2.R9xl0 and RNMAC = l.14xl0
respectively). A positive shift in pltchlng moment for angles of attack preceding CLr'IAX is also shmlQ for the 10\-1 Reynolds ntnnber condition. l\s expected, for a given lift coefficient the drag increased and the LID
decreased wlth the reduction in Reynolds number. The tail-off Lin valres at
1.3 Vs \lere 10.20, 9.85, and 9.60 for the high, intermediate, and 10\,1
)
Peynolds number condition.
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A LIFT AND PITCHING MOMENT FIGURE 135. EFFECT OF MACH NUMBER ON AERODYNAMIC CHARACTERISTICS OF THE SLAT , WITH TWO-SEGMENT LANDING FLAPS CONFIGURATION \
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C LIFT/DRAG RATIO FIGURE 135. EFFECT OF MACH NUMBER ON AERODYNAMIC CHARACTERISTICS OF THE SLAT WITH TWO·SEGMENT LANDING FLAPS CONFIGURATION (CONCLUDED) -.
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-0 bOO ANGLE or ATTA(~-D(G -0 ~o -0.700 A. LIFT AND PITCHING MOMENT FIGURE 136. EFFECT OF MACH NUMBER ON AERODYNAMIC CHARACTERISTICS OF THE SLAT WITH TWO·SEGMENT TAKEOFF flAPS CONFIGURATION • ,.
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B DRAG FIGURE 136 EFFECT OF MACH NUMBER ON AERODYNAMIC CHARACTERISTICS OF THE SLAT WITH TWO SEGMENT TAKEOFF FLAPS CONFIGURATION (CONTINUEDI ,.
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Reynolds number effects on the takeoff configuration longitudinal characteristics are shown in Figure 138. The reductions in CLMAX were 0.10, and 0.30 for the intermediate and low Reynolds number. The effects of Reynolds number on the pitching moment characteristics are similar to the landing configuration. L(D values at 1.2 Vs are 11.55, 11.40, and 11.35 for the high, internediate, and 1<M Reynolds number. Comparison with the VCK configuration indicates a greater influence on CL~~ due to a change in Reynolds number for the configuration with a slat leading edge device.
Slat with T\.,o-SegMent Flap and Horizontal Tail. - Figure 139 presents the high Reynolds number characteristics of the sla t wi th two-segment landing flaps configuration with the horizontal tail attached. Reduced stability is indicated prior to CLMAX for both stabilizer settings. Achievement of improved pitch characteristics will result from future tests with modifications to increase the lift loss inboard. The effect of Reynolds number on horizontal tail-on longitudinal characteristics is shown in Figure 140. A reduction in CLMAX of 0.25 is indicated. At the reduced Reynolds
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Pitching manent trends are similar in character to those obtained with the landing flap deflection.
Slat with Singlp-Slot Flap Charact~ristics. - The influence of Payno1ds number on the slat with single-slot flap configurations is shown in Figures 142 to 144. For the 25K/0 single-slot flap deflection, a CL~~ of 2.95 and 2.77 is shown in Figure 142 for the high and low Reynolds number test condition. The resulting change in CLMAX is O.lB. A reduced pitch stability for the angles of attack prior to CLMAX ~s shown in Figure 143. A significant reduction in the pitching manent stahility after CLr1AX is shown for the high Reynolds number condition. L/D values at 1.3 Vs \Olere 10.5 and 9.9 for the high and 1<M Reynolds number condition.
For the l5G/O single-slot flap deflection, a CL~fAX of 2.71 and 2.50 shown in Figure 143 for the high and 1<M Reynolds number condition ( CLr1AX is 0.21).
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landing flap aata. 'l1le LID values at 1.2 Vs are 11.80 and 11.70 for the high and low Reynolds number test condition.
For 51\/0 takeoff flap deflection, CLMAX values of 2.62 and 2.33 are shown in Figure 144. Figure 144 also shows a Significant reduction in stability after CLMAX is achieved for the high Reynolds number test condition. The corresponding LID values at 1.2 Vs are 11.8 and 11.6 for the high and low Reynolds number condition.
The influence of the horizontal tail for the slat with the landing single-slot flap configuration is presented in Figure 145. Pitch characteristics near stall are similar in character to those obtained with the landing oro-segment flap configuration.
Clean Leading Edge Configuration The t'4o-segyrent and single-slot aerodynamic characteristics with the leading edge device removed (cruise wing leading edge) are shown in Figures 146 and 147. For both flap systems, outboatd leading edge separation occurs at
)
decreasing angles of attack as the flap deflection is increased.
Essentially similar values of wing leading edge CPr.ml values \-lere indicated for the various flap detlections. Due to the expected outboard sepacation, only small gains in CL~~X were obtained with increasing flap deflection.
The flaps are very effective in increasing the lift at small angles of attack. The two-segment flap at 25K/12C achieved at CLa=O of 1.50.
'I\ro-segment and single slot ACLr1AX and CLa=O increments (relative to the cruise wing) are presented in the rata Sumrrary section.
Clean Trailing Edge Configuration The influence of Reynolds number on the VCK with flaps retractd (clean trailing edge) configuration is shown in Figure 148. A CLr-1AX of 2.56 for the high Reynolds number condition \-laS obtained. 'l1le corresponding value at the low Reynolds number condition was 2.14. In comparison to the cruise wlng configuration (Figure 73), the angle of attack for Cr.lW~ has been 0 0 increased from 12.58 to 20.85 , the CLMAX increased by 0.71, and favorable stability characteristics at CLMAX was obtained for the VCK extended configuration at the high Reynolds number test condition. As with the takeoff and landing flap deflection, the low Reynolds number condition
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ORIGINAL PAGE IS OF POOR QUALITY The basic trend of the pitcping moment is sl~i]~r for both Reynolds number • . tension of the VCR also reduced the LID at 1.2 Vs to 11.0 (compared to 17.~ for the cruise wing configuration).
Figure 149 illustrates the slat with flaps retracted configuration. The effect of the horizontal tail on this configuraticn is shown in Figure 150.
Hith the horizontal tail removed a CL~tA..,{ of 2.30 was obtained, with an initial nose down pitching moment at CLrtl\X. The Lin at 1.2 Vs was 12.1 (a larger value than for the VCR extended configuration). Comparison of the cruise wing, flaps retracted, and takeoff and landing flap LID characteristics, indicate improved takeoff performance may be possible by an alternate slat position. The LID perforrrance should be increased by rreans of a sealed slat configuration of reduced slat deflection, and this configuration is reccmrended for future experimental evaluation. Figure 150 also indicates, for the horizontal tail-on configuration, a reduction in stability prior, to stall, followed by nose-down pitching moment after
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CL~~. Figure 151 shows the effect of a modified trim position for the inboard slat. The modified trim position was 2.54 em (1.0 in) farther outboard than the basic slat trim near the fuselage. A small change in
CLr1AX (~Lr1AX = 0.027), an increase in the angle of attack for CLr1AX' and a
small reduction in LID at 1.2 VS' is indicated in Figure 151. A substantial
favorable change in the pitching manent characteristics is sha,m in Figure 151 for the high angles of attack. This revised trim position is also suggested for future testing with the deflected flap configurations.
Ajlercn and ~iler Studies Aileron effectiveness is presented for the cruise, takeoff, and landing configurations in Figures 152, 153, and 154, respectively. At pre-stall angles of attack, the aileron effectiveress is well-behaved for the cruise and landing configurations. The takeoff configuration data exhibit similar trends for lilOst of the angle of attack range tested, but rear the stall the
)
effectiveress of the downgoing aileron is diminished.
The shape of the rolling moment curve with aileron deflection indlcates for all flap settings that the negative deflections (TEU) are more effective than the positive (TED). In some cases, the incremental rolling moment obtained for negative ailercn deflections was twice as large as the corresponding value for positive aileron deflection. These data were obtaired in the NASA Ames 12-Foot Pxess~re Wind Tunnel at high Reynolds n~ber conditions.
Spoller effectiveness for the cruise, takeoff, and landing configurations is presented in Figures 155, 156, and 157, respectively. The spoiler data were
obtained in the tIl-SA Langley V/moL Hind Tunnel at an atmospheric P.eynolds
number. The data indicate well-behaved characteristics for the three configurations with increasing effectiveress being shown for increased flap deflections. The spoiler arrangement consists of large chord panels compatible with space available aft of the rear spar and spoiler span corresponding to the flap span. This powerful spoiler configuration is needed because of the reduced roll rate capability associated with
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high-aspect-ratio wings.
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B DRAG FIGURE 151. EFFECT OF INBOARD SLAT TRIM POSITION ON AERODYNAMIC CHARACTERISTICS OF THE FLAPS RETRACTED/SLAT EXTENDED CONFIGURATION (HORIZONTAL TAIL·ONI (CONCLUDEDI
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MODE L LB-486A CONFIGURATION SI N2AP2AZIAHIA V lA I ",,- .~\ ... ~t, MACH ~ 020 RN 512 x 10 MAC I .. 0° H hVCK ~ 010 ~F ; 0:0 001.2 "FRP (DEGI ROLLING MOMENT -4 COEFFICIENT a C\
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MODEL LS486A . ~ rJ~ "'\1_ CONFIGURA nON S4 HIA VIA MACH = 0 20 RN ~5.12x106 MAC I 3 0° H OVCK • 45E/45G OF = SC/l0S "'FRP (DEGI o ROLLING 0008 MOMENT COEFFICIENT, C,
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TEU ,-I __ ---' -20 -10 /) a LH -0008 o -0012 8.12 -0016 FIGURE 153. ROLLING MOMENT COEFFICIENT DUE TO AILERON DEFLECTION FOR VCK WITH TWO·SEGMENT TAKEOFF FLAPS
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CONFIGURATION S4 G H1A VIA 1A MACH - 020 RN ~512xl06 /,4AC I = 0° H b ~ 45E/45G VCK hF 25L112C 0.012 "FRpIDEGI ROLLING -0008 o MOMENT 8.12.16 COEFFICIENT CI
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OF FC':';~ (?~ MODEL LB-486C CONFIGURATION SlN2AP2AZ1AH1A V 1A MACH·0.20 RN ~1.14xl0o MAC HORIZONTAL TA1L-QN -0.14 -0.12 u'" I- Z w ~.10 U u..
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::iE QFRP (DEGl Cl Z ::i __ 0 ...J I I a: ~06 -004 --{).02 -10 -20 -30 -40 -50 -60 SPOI LER DEFLECTION, ~SP (DEGREESl LH FIGURE 155. ROLLING MOMENT COEFFICIENT DUE TO SPOILER DEFLECTION FOR CRUISE WING
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, ..J c;:\-: ' .' r -" ~ OF F0(; . '~ MODEL LB486C CONFIGURATION SSHIA VIA MACH ~ 020 RN ~114~106 MAC HORIZONTAL TAlLON -0.14 hSLA T 15D/25D/25E U 5elloa F -0.12 IDEGI llF R(I ...
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\" 0F H.: MODEL LB-486C CONFIGURATION S~G1AHIA V 1A MACH· 0.20 RN • 1.14" 10 MAC HORIZONTAL TAlLON 0SLAT • 15D/25D/25E -0.16 QFRP (DEGI of • 25K/12C -0.14 -0.12 ~ (J -0.10 f- Z w U
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The effect of symmetrical spoiler deflection for takeoff and landing flap ooflection is shown in Figures 158 through 161. '1bese results were obtained for out-of-ground-effect conditions. '1he large spoiler chord and sr:anwise extent is very effective in reducing the lift and increasing the draq for both flap deflections for a constant angle of attack (i.e. I aFRP ~ 0 ). A siqnificant positive pitching ~oment shift is also apparent for both the takeotf and landing configuration. tVhile the reduction in lift and mcrease in nraq would result in greater deceleration on the ground, the IJOsitive increment of pitching manent would tend to unload the nosewheel. '!he qround effect on pitchinq moment, lift, and drag, with the spoilers deflecte.I nhould be obtainen in a future test program.
Figurf' 160 indicates a reduction in r.r. of 1.18 due to 30 deqrees of spoiler
deflection, and 1.56 for 60 degrees of spoiler deflection for the landi nq flap setting. Figure 161 shais an increase in drag of 0.115 is attained for the 60 degrees spoiler deflectlon at the landing flap deflection. ~ similar maqni tude 0.143 is obtained \iith 30 oogrees spoiler deflection with takeoff
J
flap deflection (for aborted takeoff).
Landing r~ar Stunies The effect of the landing qedl on rlrag and CLr~~ is shown In Figure l6~. ~e
rlraq increment is reduced slightly with lncreasinq Cr. values. The Increnenl
-~- in draq IS very similar for the takeoff and landing flap deflections. 7he effect of the landing gear on rLMAX for takeoff flap deflection IS neqligible (-0.002) and -0.076 for the landing flap deflections.
Data Sumuary & Comparisons af Leading & Trailing Edge High Lift Concepts Fiqures 16] through 171 present the final tail-off lift, pitchinq mOMent and drag characteristics for the VCR and slat with two-segment flap configuration, and the VCK and slat \·lith single-slot flap configuration.
'llie data are for the naninal high Reynolds number condition ann have been corrected for wall ann strut ef.fects.
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rigure 171 presents a comparison of the .... linn tunnel tail-off CLrlAX l~bJeen the current and DC-lO confiqurations. The current configuration high 11ft
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_60 6 29 I F F L- oir\, IS", =r CD B DRAG FIGURE 159 EFFECT OF SYMMETRICAL SPOILER DEFLECTION ON AERODYNAMIC CHARACTERISTIC') SLAT WITH TWO SEGMENT LANDING FLAPS CONFIGURATION (CONCLUDED) O;-~:2.. . .
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" OIY o 10 -20 30 -40 50 60 SPOI LE R DE FLECTION b SP IDEGI FIGURE 161. DRAG INCREMENT DUE TO SYMMETRICAL SPOILER DEFLECTION FOR SLAT WITH TWO·SEGMENT FLAP CONFIGURATION
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A 45E/4!>G 1!lH/l00 OFI- 0 45E/45G 25K/12C ON 0 45E/45G 35R/17C' ON 'FLAPERON DEFLECTION 25° -06 -08 -10 "FRP (DEGREES) 04 -04 m C '/4 FIGURE 163. TAIL Off LIFT AND PITCHING MOMENT CHARACTERISTICS FOR VCK AND TWO·SEGMENT FLAP CONFIGURATION
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t:::. 45E/45G lSH/l0B OFF 0 45E/45G 25K/12C ON ON a '0 0.04 G06 008 010 012 014 016 018 0.20 022 024 026 Co FIGURE 164. TAIL OFF DRAG CHARACTERISTICS FOR VCK AND TWO·SEGMENT FLAP CONFIGURATION
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CONFIGURATION SYM80L LE GRID FLAP GRID LANDING GEAR OFF 45E/45G 5AI0 6- 45EI45G 15H/0 OFF Q 45E/45G 25H/0 ON ON 0 45E/45G 35R/O -- MACH = 020 RN 512>< 10 MAC NACELLES AND PYLONS ON C L
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CD FIGURE 168. TAIL-OFF DRAG CHARACTERISTICS FOR VCK AND SIrJGLE-SLOT FLAP CONFIGURATION
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J6~ 36 1501250 15G/0 t::. OFF 0 1501250 25K/0 ON 32L 32 C L \JoJ N o C _ on Cf4 - 04 FIGURE 169. TAIL·OFF LIFT AND PITCHING MOMENT CHARACTERISTICS FOR SLAT AND SINGLE SLOT FLAP CONriGURATION
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o ~' ____ L- __ ~ __ ~ __ ~ ____ ~ __ ~ __ -L __ -J ____ !- __ J- __ ~ 0.02 004 006 ooa 0.10 0.12 0.14 016 0.18 o 20 0.22 0.24 Co FIGURE 170. TAIL·OFF DRAG CHARACTERISTICS FOR SLAT AND SINGLE·SLOT FLAP CONFIGURATION
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)0l)1.I1)""" FIGURE 171. TAIL-OFF C COMPARISON BETWEEN ADVANCED COMMERCIAL AIRCRAFT AND L DC-10 MAX and circular arc vane-flap system. Figure 171 indicates a significant improvement in CLHl\X for the high aSI=ect ratio wing and advanced high lift system.
The VCR optimization ,,,as performed at takeoff and landing flap deflections with two-segment flap. The landing flap deflection was 2S11/12C. The correspondinq takeoff flap deflection for the VCR optimization was SC/IOB.
Results of the deflection and position surveys indicated that the smaller VCK deflection of 45 degrees obtained the best performance in terms of L/n and CLMAX. Hor,.rever, analysis of the surface pressure distributions sh~d that the VCR was over-deflected. Analysis also indicated that a larqer yap for the inboard VCR \-las required to suppress a premature stall of the inboard region of the ,-ring which reduced CLI-'.AX. This premature inboard stall is thought to be an effect of the adverse merging of the VCR/Airfoil wake system with the flow over the flap. The flaperon section \'las especially sensitive to the leading edge device deflection and gap.
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Increasing the gap permitted the inner wing sections to increase their rnaxllntml lift car:abil ty.
It \'las also noted that the outboard aileron region of the wing indicated poor trailing edge pressure recovery for most high lift configurations.
This was due in part to the relatively low Reynolds number in this region but also due to the conflicting leading edge deflection and position requi rements for the flapped and non-f1aPJ=ed portions of the outboard wing r:anel. For the flap~d portion of the wing, the pressure recovery \'las only to the spoiler trailing edge velocity. This was a Significantly higher velocity than freestream conditions due to the influence of the flaps. For the outboard aileron region (non-flap~d), the pressure distribution on the wing has a more adverse recovery to freestream conditions from large suction peaks due to the adverse effects of sweep, asp:ct ratio and taper. Future tests \-Till investigate differential leading edge device position for these t\lO outboard regions.
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From the VCI< grid positions investigated, the final configuration selected for the VCK leading edge device was an inboard gap and overhang of 3.5 percent and -1.0 percent (VCK trailing edge one percent ahead of wing leading edge), and corresponding values for the outboard VCK \'lere 2.5 percent and -1.0 percent. The VCK deflection was 45 degrees. This VCK configuration was selected by CLr1AX7 L/D and pitching moment characteristics.
R~sults of the slat deflection and grid survey indica teo that, for CLrlAX and L/D characteristics, the minimum slat deflection of 15D/25D (inboard/outboard slat deflection) with a gap of 2.5 percent and overhang of -2.0 percent \-laS the best overall configuration tested. With the original inboard slat sr:amlise extent, satisfactory pitch dmm at stall could !'lot I:e obtained. Improved stalling characteristics were obtained for the flaps retracted configuration \-Then the inboard slat was trimmed further out from the fuselage. This exposed mo::-e of the inner wing clean leading edge and resulted in more inboard wing stall. This slat modification did not cause a siqnificant C''11!AX penalty. Further experinental optimization is required to improve the pitching moment and takeoff L/D characteristics.
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~Tith the VCK leading edge device, both the twcrsegment and single-slot flap were optimized at each of the flap deflections. For the two-segment flap system, the main flap grid was evaluated first and was foll~ved by an aft flap deflection survey with the best main flap position. Results of the main flap position survey indicated relatively SlTall changes in CLr-2\.'{, L/D, and pitchinq marent characteristics. Fran the optimization of the aft flap deflection at each main flap deflection, values of 5C/10B, 15/10B, anc 25K/12C were the best canpranise in terms of CLw\x, CLa=O, L/D and pi tchinq moment characteristics. At 35 degrees of main flap deflection the optimization indicated the fla~ron ... ,as separated for most main flap grid positions. Reducinq the flaperon deflection to 25 degrees while rraintaining the inboard and outboard two-segment flap at 35R/12C improved the drag characteristics; however, this would not be a practical aircraft configuration. '!he single-slo' flap optimization results indicated tha t for each flap deflection the aerodyramic characteristics "'lere 3imilar to those of the train flap grid study of the two-segment flap (i.e., not extremely
)
sensitive to the grid position) • A comparison of the triITl!red CLr1AX and Cr'a=O characteristics for the various high lift systems is shown in Fiqure 172. As noted previously, it is felt that the VCI~ ... ,as over-deflected and the resulting coopuison ShCMS the slat plus two-segrrent flap iXlS the largest values of maximum lift coefficient.
v ith the single-slot flap the differences between slat and VCK are rerlllceci, \·/ith the slat being su~rior at small flap deflections and the VCr: superior at the larger flap deflections.
'T'he lift coefficient at zero angle of attack for the two-segment flap \litl: the VCK is su~rior to that with the slat irr:talled. On the other hand, very little difference is noted for the single-slot fl~p \;lith either the VCI~ or the slat leading edge device. It can also be seen that the two-segT.'lent flap at landing deflections han a higher naxirtum lift caIBbility than the ../ sinq1e-slot flap regardless of the 1eadinq edge device.
Figures 173 through IPO present tr.e trhw.ed lift and L/D characteristicn :or the VCI{ and slat with olo-seqITEnt and single-slot flaps. A campa rison of the vuriOllS configurations dt takeoff conditions is shown in Fiqure lPl.
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LEADING.
TRAILING EOGE SYMBOL EDGE --0-- VCK TWO SEGMENT TRIMMED --0-- TWO-SEGMENT SLAT STRUT TARE CORRECTED ---A--- SINGLE SLOT VCK SLAT SINGLE SLOT ~ 28 08 C CL",=o LMAX 24 04 o ~' ______ ~ ________ ~ ______ ~ ______ ~ 20 L..' __ -l. __ --1 ___ L-_---I 20 30 40 o 10 20 30 40 o 10 b (DEGREES) hFEFF IC'EGREES) FEFF \I OP8166 FIGURE 172. EFFECT OF LEADING AND TRAILING EDGE HIGH·LIFT CONFIGURATION ON C L
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AND C MAX LQ~O Due to overdeflection of the VCR, the slat ShCMS superior performance a t the lm.,er Cr, values. Only minor differences in the L/D envelopes are ShQ\-ffi between the two-segment and single-slot flaps. The overall L/D values indicate that further optiMization of the leading edge device positions for takeoff flap deflections could result in improved LID characteristics.
A cOMparison of the t\-lo-seqment and single-slot flap systems ,·lith the VCR at lanaing conditions is shown in Fiqure 182. For an equal CL (i.e., equal approach speed) the tyro-segment flap system has a superior approach L/n. In fact, it is of interest to note that the two-segment and single-slot flaps (at deflections of 15/10 and 25/0, respectively) have the same approach CL- For this CL' the two-seqrrent flap has a l6-percent improvement in approach L/D relative to the single-slot confiquration.
Figure 183 presents the Reynolds number trends for the clean wing (i.e., flaps and leading-edge device retracted), and the slat and VCR with the olo-seqrrent flap system. These rest..lts are tail-off, nacelles and pylons
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on, and ,.,ere obtained at a constant Mach number. The basic "ling shOt-]s a
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FIGURE 175. TRIMMED LIFT CURVES FOR SLAT AND TWO·SEGMENT FLAP CONFIGURATION CONFIGURATION SYMBOL LE GRID FLAP GRID LANDING GEAR 0 0'0 OFF ISO 250 0/0 0 OFF 150'250 SCIIOB OFF 6 150/250 15HIIOB OFF 14 150/250 25K/12C ON :; 0:
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R ~ 512 x 10 OF POOR Q~A-. :-'/ N MAC C CONFIGt..RATION LTRIM FLAP GRID LA~'OING GEAR SYMBOL LE GRID 0 0 0/0 OFF 0 45E/45G 0/0 OFF 0 45E/45G 5A/0 OFF ~ 45E/45G 15H/0 OFF Q 45E/45G 25H/0 ON 35R/0 ON QFRP (DEGREES) -04
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FIGURE 177. TRIMMED LIFT CURVES FOR VCK AND SI!'!(;H .. I; .. ~LpT FLAP CO~fIGURATION MACH • 020 CONFIGURATION RN ·5 12 x 10 SYMBOL LE GRID FLAP GRID LANDING GEAR 20 MAC 0 0 0/0 OFF NACELLES AND PYLONS ON 45EI45G 0/0 OFF 45E/45G 5A/0 IS OFF 6- 45E/45G 15H/0 OFF 45E/45G 25K/0 "\l ON 45E/45G 35R/0 ON ,
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(L'Dl TRIM 4' I I I I I , I ! I ! I
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C LTRIM FIGURE 178. TRIMMED UD CURVES FOR VCK AND SINGLE·SLOT FLAP CONFIGURATION r.
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MACH • 020 R • 5,2"0 N MAC NACELLES AND PYLONS ON C 16 L TRIM CONF IGURA TlON FLAP GRID LANDING GEAR SYMBOL LI:GRIO 0,0 OFF 0 a 010 OFF 0 150/250 150250 5A a OFF t:;.
150.250 15G,O OFF
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~F RP 10EGREESI FIGURE 179. TRIMMED LIFT CURVES FOR SLAT AND SINGLE·SLOT FLAP CONFIGURATION CONFIGURA TlON LE GRIO FLAP GRIO LANOING GEAR SY~'80L 00 O~F 0 0 150250 00 OFF SA 0 OFF 150 250 150250 15G 0 OFF A 150'250 25K 0 ON ~ J: , ....... _~ 12V ~ ~Q ...... '-f'~ s """1>~_ .....................
13V ~--::., ...... , s "'..2.CH 020 R 512.10~ N·.,so.C NACELLES ANa P~ LONS ON CJ6 08 16 10 01 04 14 1 a 1 :2 10 24
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c ~, ....... , FIGURE 180. TRIMMED LID CURVES FOR SLAT AND SINGLE·SLOT FLAP CONFIGURATION
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14r.----------------------------------------------------------, .... - ~ - - ~- ..... )<-----.- .
.,;-" .-.,Ir ., 12 Vs LlDTRIM _____ VCK .. TWO SEGMENT FLAP ----- SLAT + TWO-SEGMENT FLAP )( )( VCK + SINGLE-SLOT FLAP
- * - ~- SLAT" SINGLE-SLOT FLAP
4~1 ____ ~ ____ ~ ____ ~~---L----~----~----~----~ 06 08 1 2 20 16 18 10 14 22 CLTRIM i DelO 13.)09
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FIGURE 181. TRIMMED LID COMPARISON FOR THE TAKEOFF CONFIGURATION hF 15/10 1 3 Vs
-
25/0 -.
",.".----------
25/12 L'D TR1M
_____ '3;;--
35/12 ---- VCK .. TWO·SEGMENT FLAPS -- -- VCK .. SINGLE·SLOT FLAPS 4 IL- ______ -L ______ ~~ ______ ~ ______ ~ ________ ~ ______ ~ ________ ~ ______ _J 14 1.8 22 24 08 10 12 16 20 CLTRIM '} DC.J .330'j
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FIGURE 182 TRIMMED LID COMPARISON BETWEEN TWO-SEGMENT AND SINGLE-SLOT FLAPS FOR THE LANDING CONFIGURATION ,.
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J2 t " . "'" ~ · · ,LA T
G-~::: eVCK C 30 __ -
LMAX
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28 t ',' ,1> 5IVCK]
/iF s 5/10 (SLAT) ..
C /~
L MAX 26 ~
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MAC FIGURE 183. EFFECT OF REYNOLDS NUMBER ON VCK AND SLAT WITH TWO-SEGMENT FLAP SYSTEM (TAIL OFF) signficant increase in CLr1AX as the Reynolds number is varied from
atr.lospheric to maximum pressure conditions (.lr.Lrw. = 0.42). Also note\.,.orthy
is the difference in variation beb..een the VCK and slat configurations. The Feynolds nUTllber effect is Much larger for the slat confiquration. With the siqnificantly recluced suction peaks of the VCR configuration, a smaller variation with Peynolds number \-lould be expected. r~ch number can also affect the hiqh lift characteristics due to the large velocities about the leading-edqe elements near stall. This is illustrated in Figure lR4 for the clean \-Jinq and the slat with blO-segment flaps at the takeoff setting.
These results were obtained at a reduced Reynolds number due to tunnel operating limits. The clean "ling sh~'s a n.l reduction in CLr1AX as the 'bch nLmlber is increased from 0.2 to 0.32. This increment increases to 0.14 for the takeoff flap setting. These trends are similar to those of current transport configurations.
The influence of the nacelle/pylon and VCR spanwise extent is shown in
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Fiqure 185. On the left-hand plot are the CLr-1l\.x values obtained at high -..
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OF ~ 5110 SLAT EXTENDED TAIL·OFF Rn/MAC ~ 2.93 x 10 26~~------------ __ C LMAX 2.5 CLEAN WING CLt.I'AX 022 024 026 028 030
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MACH NUMBER 90P81Q7 FIGURE 184. EFFECT OF MACH NUMBER ON SLAT WITH TWO-SEGMENT FLAP AND CLEAN·WING C L MAX 36~i-----------------------------' -03 ~TOTAL
~IVCKCUTOUTAT
-02 <3 I FUSELAGE VCK CUTOUT AT ~CLMAX C SYMBOL VCK EXTENT N ANO P PYLON LMAX -01 FULL SPAN OFF 0 (NACELLES
-<>-- =¢
FULL SPAN ON l:J. AND PYLON ONI d PYLON CUTOUT ON NACELLES AND 0 BASIC ON _L- __ , o i i PYLON EFFECT - o 10 20 30 (FULLS:>ANVCKI 2.0 hFEFF (DEGREESI 1.6 Li __ ---:1..- __ -1... ___ '-- __ -'- __ --' 40 50 20 30 o 10 OF (DEGREESI EFF ,,0(. OJ JI.'
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FIGURE la5. INFLUENCE OF NACELLE/PYLON AND VCK SPANWISE EXTENT ON TAIL.QFF C L MAX
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Reynolds number with the h.:>rizontal tail removed. A full span VCR was tested both with and without the nacelle pylon group. The full span VCR was sealed at the fuselage sire and continues uninterrupted across the pylon ann extends to the wing tips. Also shown is the effect of the removal of the VCR piece in the area of the pylon, and finally, a portion of the VCR near the fuselage to obtain the basic configuration.
Figure 185 indicates a CLrlAX of 3.4 was obtained for the full spm VCR \>lith nacelles and pylon off at the o,lo-segrrent landing flap settinq. Addition of the nacelle pylon group to the full span VCR configuration resulterl in a CLr1AX reduction approaching 0.1 as shown in the right-side of the Figure IRS. Removal of the VCR in the region cf the pylon resulted in a loss of another 0.1 in CL~mx. Finally, removal of the VCR near the side of the fuselage resulted in a 0.02 decrease in CLrlAX. It is to be noted tha~ as the CLr1Jl.X recreased due to these cr.anges, the pitch characteristics a t stall
)
,,,ere improving. This resulted from the nacelle pylon addition and the reduction in the inboard VCI~ extent which were pranoting more inboard stall.
At takeotf and landing flap reflections, the total CL~1AX reduction was 0.25 and 0.20, respectively.
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CDNPARlOON OF EXPERlME~rrAL DATA HI'llI ESTIMA'lEn ME'IHODS Canparisons of theoretical and eAperimental results have been made for the crliise ~ling configuration W3B. '!he theoretical results were calculated by the Giesing vortex lattice nethod (Reference 5) and the Douglas version of '.
the Jameson-Caughey (FL022) three-dimensionial transonic flow program (Reference 3). The latter program includes an approximate fuselage simulation, an accelerated iteration step, and an iterated twcrdirrensional strip boundary layer solution.
Figure 186 presents the comparison of the theoretical and the e~rim:mtal data. '£\.10 sets of experimental data are presented. The basic data have been corrected for wind tunnel wall effects, but not for the influence of the support system (tanrem struts). The flagged symbols have been corrected for this effect. The strut tares were derived from an extensive e~rirrental p~ogram retailed in Reference 6. The struts for the current
)
configuration were placed further aft than for the previous e~rimental data base on which the strut tares were evaluated. A theoretical study utilizing the Douglas Three-Dinensional Lifting Neumann program (P.eference 8) was performed to assess the significance of this further aft strut placement. This analysis, ~lhich included the effect of the strut wake system, indicated that the further aft placement of the systeM was not a significant factor. The theoretically predicted lift increment due to the strut system was in good agreenent wit:h the experirrentally determined value.
Figure 186 also ShO'>'IS the Neumann geometry for the configuration and the theoretical model of the strut and \oTake system.
Both the Giesing vortex lattice and Jameson results are in good agreement with the strut tare corrected experimental values. The test data were obtained at the nominal high Reynolds number condition in the Ames 12-Foot Pressure Wind Tunnel.
Comparisons of experimental and Jameson calculated wing pressures were generated and an example is presented in Figure 187. '!his canr:arison is for the 72.5-percent span location at the angle of attack for stall. Good
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agreerrent is shor,m between experiment and program calculations. Comparisons alii
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Of FO':'r~ I~ .10"._., •
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1.4 1.2 C LoT • • O OB o BASIC DATA
---«- CORRECTED FOR STRUT TARES
_ GIESING VORTEX LATTICE JAMESON O~~--~--~----~--~--~ o 4 B 12 16 -4
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.OPIIU6) "FRP IDEGREES) LIFTING NEUMANN GEOMETRY AND LIFT COMPARISON FOR THE CRUISE WING FIGURE 186.
-16 PERCENT SEMISPAN ~ 72 3 " FOR PRESSURE -14 COMPARISON -12 -10 C 08 LTO C p -8 -6 0 ... '----- QFRP (DEGREES) EXPERIMENT -4 JAMESON -2 20 40 ao 100 PERCENT CHORD 9-OPIUS
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FIGURE 187. COMPARISON OF EXPERIMENTAL AND JAMESON CALCULATED PRESSURES , I '
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t, • .' • OF FC::~ Q:..: ... _ ••• at other anqles of attack and other spanwise positions indicated similar agreement between experimental and calculated pressure distributions.
Comparisons of CPMIN experimental values and the Jameson calculated values indicated the theoretical minimum pressures at the angle of attack for stall ,,'ere within 0.8 of the experimental values. The Jameson program was predicting a more positive pressure. The resulting theoretical boundary layer characteristics indicated the initiation of a rapid forward shift in seIEration location over the outboard wing p:mel at angles of attack greater tpan the experimental stall angle.
A comparison ot the spanwise variation of sectional lift coefficient for the Giesing vortex lattice, Jameson, and the Nonplanar Lifting Surface program (Reference 9) is shown in Figure 188. The theoretical methods are predicting very siMilar sp3rnlise variations of sectional lift values. Good agreement betHeen experiment and theory is shown for the inboard stations,
)
but all theoretical rethods overestiITate the sectional lift values for the outboard span positions.
o 0 INTEGRATED EXPERIMENTAL PRESSURE DATA CRUISE WING NACELLES AND PYLONS OFF ---- GIESING LIFTING SURFACE PROGRAM )( )( JAMESON PROGRAM MACH = 0,20 - - - NONPLANAR LIFTING SURFACE PROGRAM RN s 5,12 x 10 MAC 2.0 C 1.513
--
L 16 MAX
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c~ 0.8 I C = 0455 L H ¢ 0"
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0 01 0.2 0.3 04 0.5 06 0.7 0.8 0.9 10
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FIGURE 188. COMPARISON OF EXPERIMENTAL AND CALCULATED SPANWISE LIFT DISTRIBUTIONS j
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/ ,.... - p,'; ; r .-' L ~. " ...... - -' Figure 189 presents a cCX'lp:lrison of experimental and estinated naxirnum lift increments for the VCK and the two-segment flap system. Also shown in Figure 189 is a comparison of the flap lift increment at zero cegree angle of attack. Giesing calculated flap lift jncrements were essentially identical to the estimated values. The estimated VCK maximum lift increrrents were based on the two-dirrensional and three-di.mensional high lift experimental data base ''lith corrections for local chord ratio. '!he uaxirnum lift increrrents and .:lCla=O for the trailing-edge flap system were estimated by means of the t,.,o-dimensional eX{:erirrental data base and methodology usee successfully for current transport aircraft. The comparisons sho\,Tn in Flqure 189 indicate good agreement for these lift components. The flap lift increroent is the difference between the clean leading ecge with flap ceflected and the basic cruise wing lift at zero degree anqle of attack.
The maximum lift increment for the VCR is the difference in CLrlAX with and vithout the va< device for the various flap ooflections.
)
.6. EXPERIMENT (VCKI ---ESTIMATE (VCK) o EXPERIMENT (TWO SEGMENT FLAPS) -- - - ESTIMATE (TWO SEGMENT f-LAPSI I G VCI{ • TWO SEGI\IENT FLAPS TWO SEGMENT FLAPS "Cf' 121- g. __ 1 2
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A ", / / .lCL"A>( 08 LF .lC u-o 08, ,/r;Y'/
o ... .,., ... cr-
,/ / 04 o ... / 041- 0/' ",'" / ,/
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a ~I __________ ~ ____ ~~ ____ ~ ____ ~ '" oJ.:.:::.
o 10 20 30 40 50 10 20 ::0 40 o 50 0F ,DEGREES) CF IDEGREESI EFF EFF ;c~ ~_-s FIGURE 189. COMPARISON OF EXPERIMENTAL AND ESTIMATEG MAXIMUM·LlFT INCREMENTS FOR VCK AND FLAPS, AND FLAP-LIFT INCREMENT AT ZERO ANGLE OF ATTACK
A trimmed Cr. COMparison is presented in Fiqure 190. The predicten
rlAx
raximum lift values \.'ere estiIrated using an incremental buildup of the basic
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\-ling, trailing edge flaps, leading edge device, interference, and tri!':l ..
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OF P0 .. >: t~ _,' _
---
VCK + TWO-SEGMENT FLAP TRIMMED CORRECTED FOR STRUT TARES C LMAX
o EXPERIMENT (C )
L MAX -- ESTIMATE 2.0 20 40 o 10 30 "F (DEGREES) EFF ~OP8184 FIGURE 190. COMPARISON OF ESTIMATED AND EXPERIMENTAL C L MAX ) p;:!nalties. The basic equation is as follows:
CLrfAX = CLr1AX + CLlll\X + CLr.lAX + CUrnIl-t + CL~IAX
CLEAN FLAP L. E. INTERFERENCE The clean wing rraxirnum lift contribution is obtained from an estimation of the three-dimensional section l~ft distribution which first attains the estimated t\'lO-dirnensional section naxirnum lift at sene point along the span.
The three-dimensional section lift distri~ution is predicted using lifting surface theory (Reference 5), and integration of this lift distribution yields the value of wing maximum lift. The values of maximum lift coefficient across the span are obtained from a correlation of two-dimensional wind tllr'_'1el data, which account for airfoil Reynolds number and ge<::lretry, including leading e'::1ge radius, naximum c&nber, and loca tion of lJ".aximurn camber and thickness. '!he trim lift penalty at CLrlAX is estimated for the appropriate oenter-of-gravity location using the relevant aircraft geometry and tail-off pitching rnanent at each flap setting. l-tiscellaneous interference r:enalties have been obtained from previous aircraft data and applied to the estimates. Such items as nacelles and pylons, leading-edge device cutouts, and effect of landing gear are incluced in this interference
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increment.
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--
-- -- ... ~---~ Pj
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:;t= f~-:, ~ ~_:-~_~-i~1
The comparison shown in Figure 190 indicates good agreement at zero and maximum flap deflections. For the takeoff flap deflections, the estimated CLr1AX values are Imver than the experiment...Jl data. '!he estimates are, in ge!1eral, conservative for the VCK with two-segrrent flap configuration.
Figure 191 presents a ccmplrison of e~rirrental and estimated tail-off Lin
characteristics fer the VCR with two-segrrent flap configuration. The estimated tail-off L/D characteristics were obtained fran drag polars based on an increrrental buildup rrethod similar to that used on curr.ent transport aircraft. Increments for parasite and indu~d drag of the flaps ann vcr, nacelle drag, \Olere applied to the estimated lo\o/-speed cruise configuration polar. The basic equation for the drag buildup is as follows:
~2
q, = Cop + ;AFe + CopARllSITE + ACDrNooCED + f.Comooa:o + I1~SC
FLAP NN'EIJ,ffl ~ FLAP + L.E.
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HORIZONTAL TAIL OFF VCK LEADING EDGE DEVICE ENVELOPE FOR ~ F 0,5/10,15/10 ~---=
-:;:::;- ~
~ h F • 25/12 1 3 Vs _~ __ ---..J LiD • ESTIMATED 6 EXPERIMENTAL 4_'--~~ __ ~ __ ~~ __ -L ____ ~ __ ~ __ ~ 08 10 16 12 14 1.8 20 22 C L !lOP8lS' ~ FIGURE 191. COMPARISON OF EXPERIMENTAL AND ESTIMATED LID CHARACTERISTICS
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The parasite drag increment for the flap plus leading-edge device is based on previous three-dimensional test data for high lift configurations incorporating two-segment flaps, VCR, and slat geometries. Flap-induced drag incr~ments relative to the clean wing were generated by the Giesinq vortex lattice program (Reference 5) for the various flap settings. The nacelle-induced dr~g estimates are also based on the same method. The miscellaneous drag tenn includes the effect of the landing gear.
C,ood agreement bet\'leen experimental and estimated L/D characteristics is shown in Figure 191 for the takeoff flap envelop:! and the 25/12 t\olO-segment flap deflection.
Figure 192 sha·/S a canfE!.rison of the span loading for the clean leading edge (no leading-edge Oevice deployed) with the flaps deflected ISH/lOB. The theoretical results "~re existing calculations at a slightly different flap deflection and without the nacelle and pylons. The shape of the span load
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distribution is predicted well by the Giesing method. The lift values, ha-lever, are sanewhat overestinated.
MACH tJ 20 24 6 ~ 12. 10 Hr~., ol.
fj.
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fj.
1 G c C, oa INTELRAT EO E )fPERIMENT Al PR -S$URE OAT A .6..
",. 15H lOBI "~ACeLleC' AND PYLONS ON GleSING VORTEx LA. TlICE II,. 151251 NACEllE!; .l.NO PYl'~\NS OFF f)' I Q 0 1 02 0 J 04 05 06 01 OB 09 1 a 'I
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FIGURE 192. COMPARISON OF EXPERIMENTAL AND CALCULATED SPANWISE LIFT DISTRIBUTIONS (CLEAN LEADING EDGE WITH TWO·SEGMENT FLAPS) ~
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i~ECEOING PAGE BLANK NOT FILMEO CONCLUSIONS AND RECOMl'ENIll\TIONS Conclusions Results of the Phase I lOW-S{:eed high lift tests have indicated significant aerodynamic performance improvements for the supercritica1 wing with advanced high lift systems. As with any rew configuration, certain items require further development. This configura tion is no different in this respect than other aircraft that have preceded it. t-/here improvements are Cesireo, the neans to achieve these goals seem available, and are contained in the recamrendations. '!his canbined NASA-Douglas research effort has helped to provide confidence in performance levels, established a comprehensive data base for analysis of developing methods, and highlighted future development areas.
The following conclusions can be dra,m fran the analysis of the test data:
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1. The cruise wing achieved, for the high Reynolds number test condition, a trimmed CL~1AX of 1.5 and an LID at 1.2 Vs of 15.0. Pitch characteristics indicated that outboard wing panel stall characteristics are influenced by changes in l1ach and Reynolds number.
The ongoing high-speed wing development has altered the span loading in a manner that \lill improve the stalling behavior.
2. The optimization of the leading-edge devices indicated superior CL~iAX and LID performance for the siat configuration, whereas the pitch characteristics for the VCK ... ,ere superior to the slat. The loading obtained on the VCK indicated that improved performance may be achieved by a reduction in VCK deflection. The superior VCK pitch ~ characteristics resulted from an increased amount of inboard stall relative to the slat configuration and the ability of the VCK to prevent significant lift loss over the outboard ... ling. '!he VQ{ cutout at the mcelle pylon and the fuselage side contributed to the improved inboa rd VCl{ pi tch characteristics. The best compromise between perfornance and stability and control for the VCK was obtaired
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... lith a 45 VCK deflection, and for the slat, the corresponding values were a 150 inboard and a 25 outboard deflection. Improvements in aerodynamic performance and pitch characteristics could result from further leading-edge device optimization studies.
3. Trailing-edge flap studies indicated that the changes in ~rforrnance nue to gap and overhang variations of the flap system were not as significant as the corresponding variations for the leaning-edge devices. Optimization of the aft flap deflection, at each main flap deflection, resulted in the selection of SC/IOB, ISH/lOB, and 25K/12C as the best compromise in terms of Cr.~lAX' CLa=O, L/D, aircraft attitude, and pitching ~onent characteristics. As expected, the two-segment flap was superior in CLI1AX and flap lift increment.
Trimmed polar comparisons indicated equivalent L/D envelopes for takeoff flap settings. For equivalent values of approach speed, the L/D values for the t'llo-segrrent flap were su~rior to the single slotted
~
flap. Improverrent of the pitch characteristics nay also be obtained in future studies by means of different flap deflection for the inboard and outboard sections (differential flap deflection).
4. llach and Reynolds number effects \-lere studied during the test program for selected configurations. The cruise wing C!.r1AX was reduced 0.1 as the Bach number \'las increased from 0.20 to 0.32. The slat ""ith olo-segrrent flap ... ,as also evaluated for llach number effects and the reduction in CL~1AX' for the ~rne increase in l\:\ch number, \VclS 0.14 and 0.24 for the takeoff and landing flaps, res}:'ectively. The cruise wing showed a significant increase in CLrlAX as the Reynolds number WnS varied fren the atmospheric condition to the value at maximum pressure conditions (Rr-Tr1AC = 1.14xl0 to %l1A.C = S.12xI0 ). '!he CL~1AX increa~ \VclS 0.42 for this Reynolds number variation. The angles of attack for CLllAX were changed by the Reynolds number variation, but the character of the stall t-laS similar. For the high lift configurations, the effect of Reynolds number on the slat leading-edge configuration ... las much larger than for the VCR. This '>laS true for both takeoff and landing
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flap deflections. For the range of Reynolds number tested, typical values of the increase in CLMAX were, for takeoff and landing two-segment flaps, 0.32 and 0.34 for the slat and 0.15 and 0.05 for the VCK. The change in Reynolds number did not alter the basic charact-=r of the pitching moment data. Angle of attack for CLMAX and the magnitude of the pitching moment variation for poststall conditions were influanced by the Reynolds number change.
5. The effects of the nacelles, pylons, and VCI< span interruptions at fuselage side and near the pylon were to reduce CLr11\X' However, the pitch characteristics of the configuration were improved. The full-span VCK, without nacelles and pylons, achieved a CLrtAX of 3.4 with landing blO-segrrent flap deflections. This CLMAX was reduced to 3.2 with the normal VCK cutouts. The slat configuration \vas sealed in the area of the pylon and p.ad an inboard trim pasi tion closer to the fuselage side. I~proved pitch trends could result from increased
)
outboard trim position (further fran the fuselage and closer to the VCK sparn-lise position) and a revised trim over the pylon.
6. The clean leading-edqe characteristics showed dramatically that the development of the leading-edge configuration was of equivalent ~portance to the developrent of the trailing-edge high lift system.
t-1i thout the leadinq-edge device (clean leading-edge configuration), the ACLr,lAX \olith flap deflection \vas only 0.6. Achievement of large CLr1AX fer these configurations requires significant leading-edge protection for the outboard wing panel.
7. The sla t extended, flaps retracted pitching moment variation Has improved significantly by a revised trim position for the inboard slat.
This revised ~osition was at an increased distance from the fuselage side. Future evaluation of this configuration for the flaps deflected case at high Reynolds number is expected to show improvements in the pitching moment trends.
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-' R. Aileron effectiveness studies indicated that, for all flap settings, negative deflections (trailing-edge up) were more effective than positive (trailing-edge down) deflections. In some cases, the incremental rolling moment obtained with the negative aileron deflections reached values twice as large as the corresponding value for positive aileron deflections.
9. The effect of spoiler deflection on roll characteristics indicated improvea effectiveness as the flap deflection was increased.
Symmetrical spoiler deflection for takeoff and landing flap deflections showed the spoilers to be very effective in reducing lift and increasing drag.
In. The influence of the landing gear was shOlm to be negligible on CLr1c'\X at takeoff conditions, but resulted in a 0.076 reduction in CLrw.x for the landing flap conflguration. The incremental drag nue to the
~
landing gear \-las essentially the sat:le for takeoff and landing flap reflections.
11. The effects of the mini-tuft flQ\-l visualization technique on tpe low-speed high lift aerodynamic characteristics (including CLrtAX, pi tchinq manent, and drag) \..ere not significant. 'llie data did indicate a slight reduction in drag for the crUlse wing and high lift configurations with the tufts mounted on the model.
Recomnenda tions Analysis of the data obtained has highlighted areas \'lhere continued efforts could result in further improverrents of the technology. Items are suggested to improve the follo\'linq lO\-T-speed aerodynamic characteristics: LID for takeoff flap deflections, pitching marent for the high lift configura tions, and improved lift and drag by replacement of the flaperon. The test iter.s reCOMmended for future studies include:
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/' 1. Sealed Slat Leacling-Edge Configuration. - This configuration is primarily airred at improving the takeoff L/D {:erfornance. 'Ibis ty{:e of configuration has been successfully used in previous aircraft and although nechanically more complex, the existing L/O trends indicate the takeoff {:erformance can be improved.
2. Reduced YO< Deflection. - Evaluation of CPz,lnl for the VCR configura tion ; indica ted that the VCK, at the minimum ceflection tested, tray have been overceflected. A reduced VCR deflection may result in lower profile drag and improved outboard wing pal~l characteristics.
3. Segrrented Outboarc'i LeaQing-Edge DeviceS. - Both the VCR and slat grid studies indicated that the flapped and nonflapped portions of the outboard wing achieved their best performance \-lith different grid posltions. Thus, the ability to position the leading edge device incepmdently in these regions would result in improved lift, pitching
:)
moment, and drag. A cut near AD-percent semispan position for the leading-edge devices is recararended.
Revised Slat Trim. - Analysis of the test data indicated the existing 4.
inboard slat configuration did not result in appreciable lift loss at high angles of attack. A revised slat trim for the inboard slat (at a greater distance from the fuselage) was evaluated for the clean trailing-edge configuration (i.e., no flap deflection). This revised trim resulted in improved pitching moment characteristics with an acceptable penalty in CL~rnx. This configuration is reccnmended for future study with the takeoff and landing configurations. Also, a revised trim in the area of the pylon (absence of pylon/slat sealing) should result in improved pitch characteristics •
. -
5. Inboard Fixed Cani:ler Krueger {constant chordl. - lis an alternate to the existing VCK and slat configurations, this device, of simpler mechanical design, has the potential of improving the stall characteristics for this configuration by means of a more rapid lift ~, loss at the high anqles of attack.
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'" /~ 6. La rge Chord Fixed Carrber Kru~. - t'1hile sane aerodynamic ferforrnance penalties may result, the magnitude of these penalties needs to be assessed with relationship to the reduced complexity of this leading-edge device.
7. Differential Flap Peflection. - Deflecting the inboard flap system to a larger anqle than the outboard flap system is recommended for future testinq to improve the stall characteristics.
R. Ground-Effect Aerodynamic Tl3ta. - At present, one area of aerodynamic data which is still lacking for this class of transport aircraft is the influence of qrouod effect. Test data of this type are required for the basic high lift configurations at high Reynolds number. ~~so required are the effect of large syrrmetrical spoiler deflections on the high lift characteristics Hith ground effect.
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" / REFERENCES 1. Steckel, O.K.; Dahlin, J.A.: and Henne, P.A.: Results of Design Studies and Wind Tunnel Tests of High Aspect Ratio Supercritical Hings for an Energy Efficient Transport. NASA Contractor Report 159332, October 1980.
2. Sizlo, T.R.: Berg, R.A.: and Gilles, D.L.: Development of a Low-Risk Augmentation System for an Energy-Efficient Transport Having Relaxed Static Stability. NASA Contractor Report 159166, I:ecember 1979.
3. Taylor, A.B.: Advanced Perodynamics and Active Controls for a Next Gereration Transport. Douglas Pa~r 6696, larch 1978.
Callaghan, J .G. and Beatty: T.D.: A Theoretical Hethod for the 4.
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Analysis and ~sign of UUlti-Elerrent Airfoils. Journal of Aircraft, Vol. 9, fo.To. 12, ~cernl.::.ar 1972.
5. Giesing, J.P.: Lifting Surface Theory for vling-Fuselage Canbinations. r~cDonnell-Douglas ~port DAC-67212, Vol. I, August 1968.
6. Callaghan, J .G. : Perodynamic Prediction f-ethods for Aircraft at Low S~eds "lith r.echanical High Lift ~vices. AGARD Lecture Series No.
67, Prediction lethods for Aircraft Aerodynamic Characteristics, lby 1974.
7. Crm.;der, J.P.: Fluorescent Mini-Tufts for Non-Intrusive Flow Visualization. l1cDonnell-Douglas Report lIDCh17374, February 1977.
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I R. Hess, J .L: The Problem of Three-Dimmsional r,ifting Potential Flow and Its Solution by Means of Surface Singularity Distribution.
Computer Hethods in Applied ~lechanics and Engineering, Vol. 4, No.
3, November 1974.
9. Goldhammer, H. I.: A Lifting Surface Theory for the Analysis of tlonplanar Lifting Systems. fUM Pap:!r 76-16, January 1976.
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