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ASA CONTRACTOR N A.S-A REPORT
i. f! COMPUTATIONAL SYSTEM $ORAERODYNAMIC DESIGN AND ,I' :ANALYSIS OF SUPERSONICAIRCRAFT
1.'Part 2 - User's Manual ,I :c;a' I j* 1j;W. D. Middleton, J. L. Lu~zd~y, and R. G. i Coletnun
..
..I Prepared by I. . . _. BOEING COMMERCIALAIRPLANE COMPANY .. ye2 0 Seattle, Wash. 98 124 5 % 4' for Langley Research Center '>76.1916
.I NATIONALAERONAUTICS AND SPACE ADMINISTRATION WASHINGTON, D. C. AUGUST 1976 TECH LIBRARY KAFB. NM
~~ ~ ." 1. Report No. 2. Government Accession No. 3. Recipient's Catalog No. NASA CR-2716 " 4. Title and Subtitle 7- 5. ReDOrt Date A Computational System for Aerodynamic Design and Analysis August 1976 of SuDersonic Aircraft. 6. Performing Organization Code Part - User's Manual .- i .~ 7. Author(s) 8. Performing Organization Report No. D6-43798-2 W.D. Middleton, J.L. Lundry, and R.G. Coleman ~. 10. Work Unit No. 9. Performing Organization Name and Address
Boeing Commercial AirplaneCompany 11. Contract or Grant No. P.O. Box 3707 NAS 1-13732 Seattle, Washington 98 124 13. Type of ReDort and Period Covered 12. Sponsoring Agency Name and Address Contractor Report Jan. 1975-Mar. 1976 National Aeronautics and Space Administration 14. Sponsoring Agencv Code Washington, D.C. 20546
E Supplementary Notes
TWOof three final reports.Langley technical monitor: H. W. Carlson
~~ .- .~ IG. Abstract An integrated system of computer programs has been developed for the design and analysis of supersonic configurations. The system uses linearized theory methods for thecalculation of surface pressures and supersonic area rule concepts in combination with linearized theory for calculation of aerodynamic force coefficients. Interactive graphics are optional at theuser's request.
The description of the design and analysis system is broken into three parts,covered in three separate documents:
Part 1 -General Description and Theoretical Development Part 2-User's Manual Part 3-Computer Program Description
This part contains a description of the system, an explanation of its usage, the input definition, and example output.
These three documents supersede NASA contractor reports CR-2520, CR-2521, and CR-2522, which described an earlier version of the system.
-~ "- - 17. Key Words (Suggested by Author(s1 ) 18. DistributionStatement Wing design Unclassified-Unlimited Configuration analysis Wave drag Subject Category 2 Pressure distribution I 19. Security Classif. (of this report) 20. Security Classif. (of this page) 21. No. of Pages 22. Price' Unclassified Unclassified 266 $8.50
~~ ~ " . For sale by the National Technical Information Service, Springfield, Virginia 22161 ,
.
CONTENTS
Page
1.0 SUMMARY ...... 1
2.0 INTRODUCTION2.0 ...... 3
3.0 DISCUSSION ...... 5 3.1 System Communications ...... 5 3.2 Geometry Module ...... 9 3.3 Plot ...... 11
3.4 Skin Friction ...... ~ ...... 11 3.5 Far-FieldWave Drag Program ...... 11 3.6 Near-Field Wave Drage ...... 14 3.7Wing Design and Lift Analysis ...... 17
4.0 INPUTFORMAT ...... 49 4; 1 Executive Control Card Summary ...... 49 4.2Geometry Program ...... 51 4.3 PlotProgram ...... 60 4.4 Skin Friction Program ...... 63 4.5 Far-FieldWave Drag Program ...... 67 4.6 Near-Field Wave DragProgram ...... 69 4.7 WingDesign Program ...... 72 4.8 Lift AnalysisProgram ...... 82
5.0 TYPICALCASE AND PROGRAM OUTPUT ...... 91
6.0 REFERENCES ...... 251
APPENDIXA-Interactive Graphics ...... 253
iii A CO:4FUTATIONAL SYSTEN FOE AZ:FGCYi\lAi$IC DZSIGN AND ANALYSIS OF SUPERSONIC AIkCEiAFT
W. D. Middleton, J. L. Lundry,and R. G. Coleman BoeingCommercial Airnlane Company
1 . 0 SU,rnAf. Y
AL integrated system of computerprograms has hen developedfor the desiqr, arid acalysis of suFersonicconfigurations.
Thesystem ccrcists of anexecutive driver and s+ver: basic conputer programs irxluding a plot module, which are used to build up the forcecoefficients of a selectedconfiguration. Docurrentation of the systerr. has beenbroken into 3 parts:
Part 1 - Gersral Descriptionand Theoretical DeveloFment Part 2 - Usx's ?/lama1 Part 3 - Conputer Prograr Description This par%, the us?rlsrranual, conkains a dzscription of the system, an explanation cf its usage, the inputdefinition; and eXaITF1e OU+pUt,
These threedocumects supersede LASA ccntractcrreports CR-2520, CF-2521,ar,d CR-2522 whichdescribed an earlier versior. of the systom.
Interactive graFhics foruse with the systom are op%ional, em5loyifi.g the NASA-LFC CkT displayand associated software. A description cf the iriteractive graphicsportion -of thesystem is given in AFperdix A.
The comFutfrGrogram is writtenin FORTRAN IV for a SCOPE or XEGNOS operatingsystein arid library file. It is designedfor the C3C 6000 series of conputers aEd is overlay-structured. The system requiresapproximately 1150008 central memory wordsand uses eight. peripheral disc files inaddition to the inputand outFut files.
2.0 INTRODUCTION
A seriFs of ir,dividualcomputer programs for design or analysis of supersonicconfigurations has been linkedtogether into a single system. The system,because of built-incommunication between the programs, is substantiallysimpler to inputand. use than the individualprograms aperatifig in a stand-alonenode. In addition, a ccnmon gecmetryformat, based on the NASA-LFC configuration plotting program,hasbeen adopted to standardize the input requir2ments of the basicprograms.
InteractivegraFhics have been included in the system, to display or editinput and to permitmonitoring and read-out of program results. Thegraphics arrangement is tailcred specifically to the NASA-LtC CDC 250 cathoderay tube and associated software. However, all graphicsapplications nave been subroutined to the main Frcgramsand could be easilyconverted to a different graFhics set-up.
3 .0 DISCUSSION
A schexatic of the designand analysis system is shown in figure 3.0-7. Thesystem consists of anexecutive "driver1' and seven basiccomputer Frograms including a plotprogram and a geometry input module,which are used to build up the face Coefficients of a selacted configuration as shown in figure 3.0-2. The system may be used with or without interactive graphics.
The comgletedesign and analysis system is a single overlaid computerprogram, with the executivedriver as the mair, overlay ar,d the basic programs as primaryoverlays. The basicprograms manipulateinput (geometry module), draw a picture of the configuration.(plot module), or Ferformdesign or analysis calculations.
Aercdynamicforce coefficients for a selectedconfiguration ars built upthrough superposition. The individualmodules of the systen!provide data for the forcecoefficient build-up as follows: Skin friction is computedusing flat plate turbulent theory.
.m Wave drag is calculatedfrom either r.ear-field (surface pressureintegration) or far-field(supersonic area rule)methods, Tho near-fieldmethod is usedprimarily as an analysis tool, where detailcd pressure distrikutions 3re of in'erest. The far-field n?eirhod is usedfor wave dragcoefficient calculations and for fusslageoptimization acc0rdir.g to area rule concepts.
0 Drag-due-to-lift is conputedfrom the lift analysis proqram, whick breaks arbitrarywing/fuselag~/canard/ nacelles/horizontal tail configurationsinto a mosaic of "Nach-f=cxf1 rectilinearslements which are EmFlcyed inlinear theory solutions. A complementarywing desigr. andoptimization program, also using the Mach-box apFroach, solvls forthe wing shaperequired to support anoptimized Frzssure distribution at a specified flight condition.
3.1 SystsmCommucications
Communication between the extcutive and the different kasic modules is performed by disc files andlimited common block stcrag€.
5 I
Executive
graphics
. I ."""-."" I I I-". -I I Geometry Plot Skin I Far-field Near-field module friction I method method I I i I (pressure (supersonic I I Integration) area rule) I I "I I- """"- I """"""" Drapdue-to-lift Zero Lift Wave Drag (Wave and vortex)
FIGURE 3.0- 1.-INTEGRATED SUPERSONIC DESIGN AND ANALYSIS SYSTEM
L SUPERPOSITION METHOD OF DRAG ANALYSIS
Drag
Near Field UPp
IU - SKIN FRICTION
U DRAG DUE TO LIFT AND TRIM DRAG (WAVE &VORTEX)
ZERO LIFT WAVE DRAG
FIGURE 3.0-2.-DRAG BUILDUP
7 Input
All input to the basic modules is handledthrough the conmongeometry module and its associated interfaces. A fundamentalconsideration in the setup of thesystem has been thatinput to the basic rrodules would not be changed by theirincorporation into the overall system. However, to minimize and simFlify systeminput requirements, a specialgeorretry module has been created to read all input,and then sort andstructure the input Reeds of the basic programs.
ProgramSequencing
Prograrr;execution is ordered by means of special identificationcards, read thein executive, which initiate a specificoperation; for instance:
G E0M This card instructs the executive to havethe geometrymodule read configuration geometry.
PL0T This cardorders a plot of theconfiguration to bedrawn, according to size andview requirements which will be suppliod.
SKFE. Compute skinfrictior! for the configuration.
Other similar cards control the otherbasic modules. me configuration that is to be plotted, or analyzed, rieed not be ths completeconfiyuration that has beeninput.. Also, thegeometry dpfirition may be updatedwithout complete replacement of the gecnetryinFut.
A sunmary of theexecutive control cards is given ir: Section 4.
For each basicFrogram, there are some inpu?sthat ars not geometry.(e.g., iwiach number, nurbfr of longit-udinalcuts in analysis, ~CC.) These inputs are giva immediately after the prograrcalling card and are readin the properixterface routine in r-he geometrymodule.
3) Program Answers
A lirrited,amount of con!mon storagebetween the different Froqrars is usza to preserveanswers 2nd transfer data betweennodules, Tne wingdesigr! nlodule is thelargest singleprogram in thesystem. Therefore, some cormon blocks us2din the wing design program are carried also
8 in the executive level withoutincreasing total system size. These data blocks include:
0 Wing camber surfacedefinition Wing thicknesspressures 0 Fuselageupwash kouyancy pressures Nacelle pressurefield
0 Asymmetric fuselagebuoyancy field , (nonmid-wing configurations)
Another data block transfers the optimized fuselage area distribution, based onwave drag considerations, to the gecmetrymodule for updating.
3.2 GeometryModule The function of thegeometry module is to read systemgeometry input,update it if required,and arrange it as needed tor the individualprograms of the system. A schematic of thegecmetry module is shown infigure 3.2-1.
Thegeometry module is accessed by theexecutive corltrol cards GEOM NEW (input new configuration) or GEOM (addition or reFlacement cf comFon2nts) . The geometrymodule is also called to updatethe fuselage or wingcamber surface d2finitions if the executivecards FSUP or WGUP are read, In addition,the geometry module is called by the executive as an interrrediate step in the execution of anyof the basic programs. This requiresthe proper interface routine to be entered,the systerr geometry to be put into the correct form for theprogram to bc, exscuted, and any special(non-qeometric) data required to be rsad. This is all stackedin the proE;erorder, whereupon the executive then calls Che basic program.
Incrdrr to niniriz? core storagerquirements of the input data, both the basicsystem geonetry and the transferrsdinput- (from the geometrymodule to anotherprcgran) are stored on tape (or disk). The kasic system geometry is reserved on a tape when the geometry moduls is not in core, and tht input %tack'! for a givenprogram is written on a tape to be readby %he program when called byth? executive. The inputtape created by ths geometrymodule %hus mersly replaces the usual input taFe written from cards.
The fornat of tne sys%emgeometry input is the same as that of the NASA-LKC Flot program(reference 2). Thsre are some restrictions (relative to thereference 2 input) in the allowable numtGr of inputdefining stations, however.The gecmetry forrat and limits are giver! insectior. 4. Some optionalgeometry has alsobean added. Thisincludes provisioris for fuselag2 perimeters to he inpt {if needed by the skinfriction program), and Frovisions for
9 r""""""- 1 I Executive I L"""" -4 IL """,""""Read control card I GE@M or SKFR. ANLZ. etc. I GEOM NEW card
Interface subroutines
(Set up input forselected program)
or update basic Near- configuration Skin field field . design Lift Plot geometry I friction analysiswave 1 and I ::: drag optimization I
"""" I """""""""I + """"" """ "- 1"""""" 1 """""""_I I r Read additional input as necessary Store basic geometry I I I when not Write input tape for selected program
(Basic geometry + addition?l input, structured as required)
Basic Input file geometry for selected file program Itape 1) (tape 2)
i
FIGURE 3.2-1.-SCHEMATIC OF GEOMETRY MODULE wingcamber surface input at planformspanwise stations other than thcsr sFecifiedfor ths system geometry. This canber surface definition,called WZQIHD, is datain the formnormally ger,erated or ussd ky thewing dPsign and analysis prograrrs. Also, nacelles may he located eitherin the coordinati system of the basic gecrretry, or relative to the local wingsurface, whichever is more conveni2nt .
3.3 Plot
The plotmodule ger.erates the necessaryinstructions for drawinqs of tk? inputconfigurat.ion, either in hard-copy form (Ca1 corn^) or bnthe cathode ray tu'be. Various view options are available. The viewcFtion and drawing size are controlled byprogram inputs.
The lot Frogram was developed at NASA-LX andhasbeen incorGoratedinto the system with minimum change.Docums~t ation of the proqram is presented in reference 2.
A +yFicalconfiguration drawing generate3 by thlplot Frogram is shcwr ir, fiqure 3.3-1.
3.4 Skir,Friction Skinfriction drag for a configuration is computed by separating theairplane into its components,then calculating wetted area aad thfccrrespondinq turbulent skin frictiondrag tor each component, Thewing, tail and/orcanard (corr,ponents which may havelarge variationsinchord length) are strip-integrated to obtainan accurate,averaqe skinfriction coetficilnt. Skin friction coefficients are computed from the method of refererce 1.
Flightconditions for skinfriction calculations may be input 21th2r as iqach Rumber/altitude, or Reynoldsnumber per foot acd total temperature.If the userwishas to inputwetted areas for thedifferent conponents, rather than have the program generate the wetted areas from thesystem geometry, several specialinput oFticns are Frovided.
A schoxcaticcf the skin friction program is shown in figure 3.4- 1.
3.5 Far-Field Wave CragProgram
This Frogram CC~FU~~Sthezero-lift wave dragof an arbitrary configuration by means of thesupersonic area rule.The program was criginallydeveloped at the i3oeing Company, andhas been docurrerked(reference 3) andupdated by NASA-LEC. Theversion of the Frogramused in tne design and analysis system is that of LRC. FIGURE 3.3-1.-TYPICAL PLOT PROGRAM DRAWING r --1 I I
13 The far-field wave dragprogram is extremely versatile,and includEs a fuselagearea optimization feature which is very useful, The fuselageoptimization is accomplished by requiring the program to optimize the overallar2a distribution of wing- nacelles-tail, etc., subject to a few fuselagearea control points or The programl'fills-inttthen the non-restrained fuselagearea distribution in an optimumfashicr. for minimum wave drag.
In the dtsignand analysis system, a fuselagearea distribution may he optimizedby initially defining it in the basicgeometry, optinizing the definition in the far-field wave 6ragprogram, and then transferring the optimizeddefir-ition to the geometrymodule for use in further desigr, or analysiscycles. The actualtransfer of the optimizedfuselage geometry is performed by use of the executive card FSUP, as described in Section 4.
3.6 Rear-Field Wave Drag
The near-field wave dragprogram computes zero-lift thickness pressuredistributions for arbitraryan wing-body-nacelle configuration, The pressuredistributions are integratedover %he .cross-sectional areas of the configuration to obtain the resultant drag force. This force may or may notcorrespond directly to the dragcompted by the far-fleldmethod, depending upon the degree of tttransparency''specified for the near-fieldpressure integraticns. i3y transparency is meantthe assumFtion of the far-field mtkhod that Frsssure fieldsfrom all comFonents "pass through"and interact with all othercomponents, regardless possibleof physicalbarriers imposed by in-hetweencomponeEts.
Typicalpressure data from the near-f ield program is presented in figure 3.6-1. A wave dragcoefficient summaryfrom the program 4s shown infigure 3.6-2. .
The near-fieldprogram has three principaluses:
1) As ananalysis tool for studyingthe zero-lift drag forces associated with theinteracting pressure fields of d,ifferentconfiguration components. In this respect, thenear-field program has anadvantage ovEr the far- field wave drag method in that there need be no assumption of transFarency.
2) As a source of loads data for structuraldesign and analysis.
3) As E sourceof thickness pressure fields for usein the pressurelimiting 0ptior.s of the wing desigr.and lift 0
i I I I
3 3
I 0 I d rn N 0 I 0 I 0 0 I X I 4 0 II I hl cn I N C I '5c:
-aP 0 nI) 0 1 I I 0 I I e I (vrn I 0 I 0 I 4 I 0 0 I1 > 2 51 0 nI) 0
UI d hl' hl 99 9
0" >- I 1 I I I 0 0 0 0 W W d N 0
15 *I Body-on-wing interference
-.0041 I I I I I 0 .2 .4 .6 .a .10 Semispan fraction
Wing-Body Terms
Cgwing= 0.00639 = -0.00013 CDwing-on-body interference = 0.0071 1 = 0.00072 = 0.0001 3 '"body '"body-on-wing interference
Nacelle Terms Inboard Outboard Isolated C 0.000750.00075 "wave Body-on-nacelle interference -0.00002 0.00000 Nacelle-on-body interference 0.00005 0.00010 Nacelleon-nacelle interference Direct 0.00034 0.00023 Image 0.00054 0.00046 Wing-on-nacelle interference -0.00043 -0.00058 Nacelle-on-wing interference -0.001 56 CDnac = 0.00064
Wing-body-nacelle C = 0.00775 "wave
FIGURE 3.6-2.-TYPICAL WA VE DRAG COEFFICIENTSUMMARY NEAR-FIELD PROGRAM (M= 1.1) analysisprograms - (This option. is describedin section 3.7, butbasically requires that the total surface pressurecoefficient on khe wing, i.e., thickness+lift, cannot be less than some specified fracticn of vacuum pressure coefficient.)
If the wingthickness pressures are to beused by the wingdesign or liftanalysis proqrams inpressure linitingoptions, then the near-fieldprogram must first be run,Curing program execution, thicknessthe pressures are loadedinto a system common blockand are then available whPre needed,
"""_Nacelle -pessure """_ """field oJtions""e - The near-field programallows for UF to 3 pairs of nacelles located external to the wing- fuselage (Or 2 pairs plus a singlenacelle at Y=O) - The nacelles may ke eitherabove or bslow the wiq (or both). The mcelle Fressurefield is thepressure field imposed cn the surface of the wing by ther.acelles. 9- feature of the near-field prograr is the choiceofftwrapl* orI1glanceff solutions for the nacellepressure field, as shown in figure 3.6-3, (The far-field wavedrag program uses EssEntially the !'wrapI* solution).
AvailableexFerimenta1 data do not make it clear whether a tfwrap*' or **glanceffsolution is more correct. Sincethe nacelle-on-wing interference term is substantial, both solutions are availablein the Frogram(ccntrolled by aninpu% code) .
3.7 VingCesign and Lift Analysis
The wingd2sign and lift analysisprograms are separate 1ift.ing surfacerethcds which solve the direct or inverse problerr, of:
d Dssign - to defir.2 the wing camksr surfaceshape required to produce a select2d lirtirlgFrczssure distrihut.icn. The wing designproqram includes methods fordefir?ing an optinurnpressure distribution.
Lift analysis - to d2fir.ethe liftingpessur2 distrikutionacting on a given wing cambersurface shape,and calculate tke associated forcecoefficients.
The lift analysisprogram contains solutioRs for the effect of fuselage, nacelles, canardand/or horizoctal tail, and wing trailingedge flaps or ir-crementalwing twist. Using superFosition, the programsolves for drag-due-to-lift,lift curve slcF2, Fitchi2gand mcmnnr cnaracteristics of a given configurationthrough a range of angles of attack at a selected ;4a ch r.umlc;sr.
17 The nacelle pressure field and accompanyin?shock waves "glance" away from the wing when encountering adjhilentnacelles. In application, the nacelle generated Pressure field is terminated on encountering anothernacelle.
PRESSURES "WRAP" AROUND ADJACENT NACELLE
The nacelle pressure fields and accompanyingshock waves "wrap" around adjacent nacelles. In application, the nacelle generated pressurefield is allowed to pass through another nacelle as if it were transparent.
FIGURE 3.6-3.-NACELLE PRESSURE FIELD CONCEPTS c
Thewing design program is mor? limited in scoFe, six.5 it is used tc sclvefor the wing shape required to support a designpressure distrikution at a specifiedflignt cocditior.. The program also contains,howver, a number of optionalfeatures for identifying thedPsign pressure distribution, This is a demandingso.lution, bEcause it Leequires that:
0 Crag-due-to-lift of thewing be minimized at a qiven total lift,svhject to anoptional pitching moment constraint.
0 Constraints be applied to designthe Fressure distribution -to provide physical realism.
0 Effects of fuselagfupwash, nacelle pressure field, etc., be reflectedin the design solution.
wingJesign and Optirt5zation
Given a wingplar.forrn and flight conditicr., the wingdesiqn Frcgram solves for an oFtimum (least drag)Fressure distribution and the corrssponding wing shape,subject to specifiedconstraints on :
0 Total lift
0 Pitching moment at zero lift
Uy.Fer surface pressurecoefficier,t level and/or streamwis?gradient Ordinate at defipndFlanforn locations
Basically,the method of thewing deeigR program is that of refersncos 4 and 5. Foruse in the integrated design and analysis system, however, theprogram has beensubstantially expanded to providethe following capability:
0 Use of anycombination (or all) often basic lifting pressureloadings, in an optimum fashion. .. Optimalimposition of pressure level andpressure gradientconstraints on thewing upper surface, to preventoccurrence unrealisticallyof low pressure coefficients.
a optionalconsideration of threeconf iguration-dependent loadings(fuselage upwashandbuoyancy, andnacelle pressure field) .
19 ...... " ...... _.. . ._.. - - , .,, , , ,,,,
M = 2.7 c = .1 Ldesign
Note: At the design points denoted by circular symbols, C 0.7 C P upper Pvacuum Pupper surface Wing thickness pressures included
Two and three loading combinations are the first two and first three loadings in Table 1
Uniform loading and nacelle buoyancy loading
Flat trrlnn - .0070 Two loadings and nacelle buoyancy loading
Three-loadingoptimum .0060
Optimum combination of three-loadingsand nacelle buoyancy loading
.0040 Nine-loading optimum Optimum combination of nine loadings and nacelle buoyancy loading .0030 I I -.02 -.o 1 0 .o 1 .02 .03 .04
FIGURE 3.7-1:-EFFECT OF NUMBER OF LOADINGS ON WING DESIGN
20 e Optionalconsideration threeof wing camber-induced loadingswhich are proportional to threethe configuration-dependentloadings. This introduces camber-related terms to modulate the configuration relatedloadings (Example: trailingedge reflex for nacellebuoyancy loadi.ng).
0 optionalidentification of a small planform region (e.g., trailingedge flap) for special incremental loading. * Optior,alconstraints on cambersurface ordinate at specifiedplanform locations.
The Fresentat.iop of thewing design results, for selection of an oFtinurr.Fressure distribution, is in theform of drag-due-to-lift versuszero-lift pitching moment (Cmo ) . A typicalpresentation is shown infigure 3.7-1, illustratingthe effect of increasing the nurrber of designloadings and adding the nacelle-buoyancy loading.Selecting a CL and C,, combinationfor the wing defines a COrreSFOndingpressure distribution which may thenbe used to generate the associatedwing canker surface shape. (The bucket slot is notused with ordinateconstraints, however , and only the sclutioncorresponding to the designpoint values of CL and cm0 is printed. )
"""""""""_P.ressure const.raints. - The use ofa large number of basic wing loadingspermits great flexikility in idEntifying a theoretically oFtilruE1iftir.g Fressure distribution. Such anoptimum may be physical'lyunrealistic, however. Linear thsory contains no lirritationsallowableon surface pressures, e.g., "optimumt' pressuredistributions may well involveupper surface pressure coefficients lower than vacuum Cp. To avoid this possibility, a pressurecopstraint formulation has been added to the solution. Thisfunctions by limitingthe total wingupper surface pressure coefficient %o beequal to or greaterthan an input Cp, andby limitingthe longitudinal gradient of thisupper surface pressure to be less than or qual to an inputgradient level. By superposition, the totalupper surface Fressure coefficient is the sum cf wingthickness pressure (from the near-field wave drag program, as noted in Section 3.6), fuselagepressure field, and +-he uFFersurface lifting pressure.
The effect of constraining the allowabledesigr, pressure distribution for a basic wingplanform fuselage){no is illustratedin figure 3.7-2. For a givenplanform and set of loadirgs, the programcycles to find an op-imum pressure distribution (least drag)subject to inputconstraint conditions. First an optimum loadingcombination is found,then the CorrEsFondingFeak pr9ssure level and gradient are located. If either violatethe input limits, a new optimumLoading is found
21 DESIGN POINT nM = 2.7, CL = .10 ARBITRARY REGION (REF) 1. Basic wing solution (no fuselage) 2. Cp limit = .7 vacuum level 3. Gradient limit set by maximum obtained with optimum 3 load solution -
Flagged symbols are .6 for solutions having same maximum wing upper surface pressure gradient
.5
.4 11 loads
.3
.2
.1
-0- I I 1 I I 1 -.004 .004 .008 .012 .016 .020 .024
CInO
FIGURE 3.7-2.-EFFECT OF PRESSURE CONSTRAINTS
22 with a Fressureconstraint applied at the locationof the raximum. pressureviolation. The new oFtiIrum is thsnexamined, etc. Gradient is everywheresatisfied hefore level is constrained, as describedin the theorydocument, volume 1.
Th2 cyclicoperat.ion continues w-til the wingprPssure distributioneverywhere satisfies the pressure constraints. Tn the 2xamFle case shown, the effect of addingpressure constraints shiftsthe drag minimum fromthe bucket plot level to the level indicated by the flagged symbol. It canoccur that the inputpressure gradient constraint cannot be satisficdwithir; theother cortstraint bounds of Cmo, wing thicknesspressures and/or Z constraints.Inthis case, the prog-ram automaticallyincreases the inputacceptable gradient level by 20 percentand tries again.This process will continue until the gradientlevel is satisfied. No similar oFtion is aFFlied to tke Fressurelevel constraint, however; ifpressure level carmotbe satisfied, the progran halts.
A furtherdiscussicn of pressureconstraint application is given on page 29.
"""Loading """""" definitions. - A tabulation of the pressureloadings available within the designprogram is given in Table I on pqe 24. 'The confiquration dependentloadings may be usedboth as a su~erimposed,independent effec% ar.d also as a definitionof a loadingwhich may be varied (bywing camber) in the optimization Frccess.
As an independenteffect, theconfiguration-deFendent loadingacts uFon the wing in the optimizationprocess, but cannot be varied (loadings 15- 17) .
0 As a loadingdefinitior, (12-14), a configuration- dependentloading produced wir.9by camber may be introduced in addition to its independenteffect, The optimizationthencould cancel the lift of the independenteffect with thiscambEr-generated loadir,g, if that were the-optimum solution.
A cocfiguration-dependentloading may not be used as the source of a variableloading without also using it as anindependent loading.
"""Fuselage """_ in wing""_ design """" solution. - The fuselage may be included in the wingdesign solution by inputof fuselage geometry and specifying a side-of-fuselagesemi-span station. The resulting sclution is then split into two parts: the wingpart (outboard of side-of-fuselagestation) with loading definitions as described previously,and the part(inboard of side-of -f uselage station) . Loadingsinboard of the side-of-fuselagestation are of TABLE I DESCRIPTION OF WING LOADING TERMS
Loading Number Definition
1. Uniform 2. Proportional to x, the distance from the leading edge 3. Proportional to y, the distance from the wing centerline 4. Proportional to y 2 5. Proportional to x 2 6. Proportional to x(c - x), where c is local chord 7. Proportional to x' (1.5 c -x) 8. Proportional to 2 (1 + 15 5)-0,5C 9. Proportional to (1.05 c-x)'.~ 10. Elliptical spanwise, proportional to i- 11. Proportional to x, the distance from the leading edge of an arbitrarily defined region 12. A camber-induced loading proportional to the body bouyancy loading 13. A camber-induced loading proportional to the body upwash loading 14. A camber-induced loading proportional to the nacelle buoyancy loading 15. The body bouyancy loading 16. The body upwash loading 17. The nacelle buoyancy loading
24 theltcarry-overl1 type, and are dependentfunctions of theloadings outbcard of the side-of -f uselage.
Drag of the8tfuselage11 part is calculatedby applying the carry- overloadings to thefuselage camberline. The outboard, or wing, part is handled as for the wingalone case, withintegrations begirning at the side-of-fuselage. The wing-fuselage solution ' thusreflects icterdependencethe wingofand fuselage contributions to the wing designoptimization. There are severalconsiderations of importancethein wing/fuselagesolution:
a Wing paneling(internal definition of winggeometry) may require a slightshift in the inputside-of-fuselage station.This is accomplishedautomatically thein Frogram and anexplanatory note is printed. * Thefuselage attitude and wing camberline at %he side- of-fuselagemust approximately align for thedrag integrations to bevalid. Experience has shown that it is necessary for 53 constraints to be applied to the wing canberline at thPside-of-fuselage forathisto occur. (e constraints are discussedin more detail on page 33). * For convenience, the fuselageattitude (relative to +he basicgeometry definition) in both the lift ar.alysis program (to generate upwash loading)and the wing diesigr, programcanchangedbe withcllt revising the basic geometry.(In the lift analysis rndyle, this is done by a specialapplication of thepr2ssure liriting option. Set FLINIT=l .O, anappropriate value of vacuum fraction VACFR, andincrement the fuselageangle of attackin TLALP) . Inclusion of thefuselage theinwing design solution is illustratedin figure 3.7-3. In the example cas2 snown, the wing canberlin? at the side-of-fuselage was constrained at the four lccationsindicated. A bucketplot is notproduced when E constraints are used;however, the eff2ct ofadding pressure constraintsto the solution is illustrated by the syxrbols on the drag-due-to-liftversus Cmo plot.
""""""Use cf confiquration-dsendent "-"" """ """_ loadings. - AD. example of the inclusionof a configuration-dependentloading is illustrated ir figure 3.7-4, showinglvreflexinglt of the wingdue tonacelle influences. The wing trailingedge is bentupward locally, or reflexed, to takeadvantage of positive pressure coefficients from the nacellepressure field.
2 5' Design Conditions: M = 2.7, CL = .10 /
CPlimit = "137 dCp/dX = .0025
17 loadings 8 2 constraint +- 60 100 140 180 140 100 60 220 260 X
.60 -
Drag-due .50 - to-lift, xfi 'D/CL2 +
.40 - I I I .006 .010 .014
.7 5 0 .e5 ~ ~~~
\ \ .50 4
Z
.25 8 Camber sutface for CL, z. cp, cn 12 r
FIGURE 3.7-3.-W/NG DESIGN OPTIMIZATION WITH FUSELAGE ANDZ CONSTRAINTS
26 L Nacelle I
I-X 0.05 0.30 .3
0
.3 .4 .5 .6 .7 .8 .9 1 .o XI L
0
-4 N Camber surface for optimum combination Yl2b of nine loadings plus nace lle pressure nacelle field * 0.30
-1 2 Camber surface for oDtimum combination oi nine loadings
-16 L FIGURE 3.7-4.-EFFECT OF ADDING NACELLESTO WING DESIGN SOLUTION
27 The loadings due to thefuselage include both liftcaused by the fuselage upwash field and also lift due to asymmetricdistribution of fuselagevolume ahve and belcw the wing(nonmid-wing arrangement) . As a special case, theasymmetric fuselage buoyancy loading (number 15), can be usedeven if its net lift is zero;this feature permits theinclusion of fuselagethickness pressures in theFressure limiting case for a mid-wingarrangement. However, if the fuselagebuoyancy lift is zero,the wing camber loading proFcrtiona1 to the fuselagebuoyancy loading (number 12) cannot. be used, since it wouldcause the optimizationsolution to fail.
Optimizationofthe wing design considering influence ofthe fuselage upwash field is performediteratively, using both the wingdesign and liftanalysis modules. A fuselageshape and iricidence is first assumed, the upwash fieldand corresponding lcading is calculated by theanalysis program, and the design solution is performed. aecause theresulting wing shape probably differsfrom the shaFe used ir; the initial upwash solution,the upwashloading is incorrect, It. may be desirable to thenrerun the wing dlsignsolution and/or alter the fuselageangle of attack. A representativeprogram executive card sequence would be: ""_Event """"""_Executive Card Definefuselage G 90M
Calculate upwash ANLE (WHUT=1.0) loading
Wipg desigr,solution WDEE
Kecalculateupwash ANLE (WHUP=l .O, loading with pew TXr'X=3.G) carcber surface
When the wir.g cambersurface is finalized, it may be transferred into the hasic gecrnetry by the executivecontrol card WGUP, (With interactivegraFhics attached, the desicjnwing shape xay also be viewedand edited betweendesign and analysis solutions).
Sxall--Elapfq+m- "" reqion oDtioy.. - Since thers ray be small rlgions of the wing(such as a trailing edge flay:).. that could he relatively highly loaded tc goodadvantage, a programoption allcws the definition of sucn a region md a correspondingloading (nc. 11 ir, Table I).
28 An example of the use of theFlanform region oFtion.is shown in figure 3.7-5. Inclusion of the regionand loading 11 results in a srall improvement in drag-due-to-lift,especially as C,, is increased.
A conditionimposed upon the planform region option is thatthe regicncannot be re-entrant in thespanwise direction, relative to the forward end. The region is inputstarting at the most inboard spanstation (which will be at the wing trailingedge), and successivespa.n stations must increase monotonically.
Loading 11 andthe small planforn:regicn are only used in combination with each other .
”Inptl ””””“””” considerations. - The wing design program principally requiresthe specification of a set ofloadings, a designpoint, and the definition of fourbasic control parameters. The control parameters(on card 7 of the designprogram inpt) govern the type andextent of the solutior..
The designpoint solution may beobtained with constraints on:
0 CL only a CL and Cmo W CL andupper surface pressure 0 CL, Cm0,and uppersurface pressure
If ordinateconstraints are specified,they are includedwit.h each of these fourtyFes of solutions.
The fourtypes of solutions are notcomFletely independent. If the CL andconstrained Fressure solutio;l is requested,then the programmust firstgenerate the CL onlysolution. Similarly, if the CL, CmOs andconstrained Fressure solution is requested,then the Frogrammust first gsnerate the CL and Cmo solution.Thus, if the uppersurface pressure ccrstraint conditlon is requested, the prcgramperforms the correspondingno pressure constraint solution whetter it was requested or not.
It is not necessary to calculatethe cambersurface shape corresponding to a specificdesign point (lift coefficient, pitchir;g moment coefficisnt,constraint condition) in order to obtain +he drag-due-to-liftversus Cmo plot. Also, if thedesign cambersurface is requested, it may be onlyFrinted out, or’ may be also Funched intccards (for later ir,putinto the lift analysis. Frcqrav) . zgessure-constraint aEplicatihq. - The constraintsapplied on the pressuredistribution of the wing uppersurface are of two types. Both pressurelevel and the longitudinal gradient of pressurecan he ccnstraired, This has beec donabecause linear aerodyanmic theory can producepressure distributions (in terms of pressure
29 C = 0.1 Ldesign Linearly varying region (loading 11)
Nacelle pressure field .7
f Flatwing Uniform loading .6 dc f Three-term solution Without small .5 planform region Option 7 d I 1
.4
CD- c2L .3 planform region option solution envelope, ILower bound solutions 17 loadings satisfying constraint .2 2 0.7 C ‘Pupper Pvacuum wing surface 17 loadings .1
I I I -.02 -.o 1 0 .o 1 .02 .03 .04
cmO
FIGURE 3.7-5. -PLANFORM REGION OPTION
30 coefficient)requiring pressure lower than vacuum on thewing uppersurface; the first type ofconstraint can be used. to restrict .pressurelevels to more realistic' values,High, positive values of Fressuregradient tend to producebouxdary layer separation, ar.d thesecond type of constraint offers a meansof control.Constraints on pressuregradient are also quite valuable in ccntrolling wide variations of pressure level.
Pressureconstraints of bothtypes are specified by theuser with tahlesof acceptable values. The tabular format hasbeen chosen so thatthe values can be varied deper,ding on planform location, After a desiqnsolution has been obtained in terms ofthe compcnentloading factors Ai' thepressure level andgradient on the wing upprsurface are compredwith the acceptable values. Ifeither or kothtypes of conditions are violated,.the program retains the Flafiformlocations of the violations in order to apply constraicts, When Fressure level violations are detected,the prograr computPs thelevel of liftingpressure coefficient at the crit.ica1 planform locationwhich in combination with thickness pressure will produce a wingupper surface Fressure equal to the acceptablevalue. A cox-,strainedlifting pressure coefficient set at 95% cf the allowahlovalue is then inposedon subsequent solutions for the optinurrcombir,ation of loadinqs. When Fressure'gradient violations are detected,an acceptable vzlue of pressuregradient due to liftingpressure coefficients is calculated at thecritical planform location. Account is taken of the gradientdue to wingthickness pressures and any configuration dopendentloadings. A constrainedgradient set at 75% of the allowablevalue is thenimposed on subsequentoptimizations.
The two factors (95% and 75%) havebeen applied. to diminish a tecdency of thesolution to varyslightly above the limiting value inthe immediatevicinity of theconstraint. This type of behavior is most noticablefor cases involvingthe configuration- dependentloadinqs, which are specified by tables as a functionof FlaRformlocation and are notnecessarily smoothly varying functions.
The application of constraints is cyclic.If the solution without pressureconstraints does pot satisfythe pressure criteria, a pressurec0nstrair.t is imposed and a secondsolution is obtained andtested. If thesecond solution does not satisfy the criteria, a secondpressure constraint is imposed,and a thirdsolution generatedand tested, and so forth.
Constraintsimposed early inthis cyclic fashion can become unnecessary as later constraints are imposed. Thesolution processincludPs logic to search for unnecessaryconstraints after thesecond and Subsequentcycles (the first constraint would not
31 beaFplied if it were notnecessary). At eachsolution cycle, any unnecessaryconstraints are removed,and the next cyclehas fewer pressure constraints. Theconstraint identified on thecurrent cycle as most critical is also deleted from thenext solution cycle;with fewer pressureconstraints, it might no longer be the most critical Fressurelocation. In sunmary, the results of eachsolution cycle are tested in three ways. First, Lagrangethe multipliers of solutionthe corresponding to thepressure constraints are tested. If any constraints are unnecessary,they are removed,andanother solutioncycle is begun, Ifall of thepressure constraints are necessary,then the most critical pressuregradient is tested.If it exceedsthe allowable value, another gradient constraint is iqosed,and another solution cycle is begun. If.the solution pressuregradient is satisfactory,then the most criticalpressure level is tested. If it exceedsthe allowable value,another level constraint is imposed,and another solution cycle is begun. If the most critical pressurelevel is acceptable,the solution as a whole is acceptable.Solution cycling continues until either an acceFtablesolution is obtained(cycling stops), or until a limit on the numberof constraints is reached. Thereare two types of limits on the maximurr numberof pressure constraintsthat can be imposed. One limit is imposedby the numberloadingsof used, The total number constraints,of includingthcse on lift,pitching moment, configuration-dependent loadings,ordinates, and Fressure, can at most be equal to the number of loadings.This situation is undesirablefor it leaves no degrees of freedomfor drag minimization; consequently, a progran; limit hasbeen set so that two degrees of freedomrerraifi freefor drag minimization (for small numbersof loadings, this is reduced to one). Thesecond linit imposedon the number of Fressureconstraints is dictatedby the number of loadinysthat are free to influence longitudinalFressure gradient. k maximum often lcadings does so - loadings 2, 5-9, ar,d 11-14. It has beenthought desirable to leaveone degree of freedomfor drag minimization for gradient constraints.The number of permissiblegradient constraints is reducedby one more wherever a Constraint is imposedon Cmo since Cm9 constraints are satisfildprimarily by the same x-depmdent loadingsused to satisfy gradiEnt criteria.
If the programreaches a liwit on t.he number of pressure constraints, it checks to see if gradientconstraints have been inposed.If one or more gradientconstraints have been imposed, the Frograxarbitrarily increases the gradier,t criterion table by 20 prcent, and begins anew withno pressure constraints. This process is also cyclicand can be repeatedup to 50 times before haltingwith the solution produced by the last cycle. If nc Fressuregradient constraints have been used, the program stops cyclingupon reaching either one of the two constraint limits, and retains the last solution.
“”””“”””_Ordinate constraints.- Linear theory producesor kinks, wrinkles, in thecomputed camber surfaceaft ofwing leading edge breaks. This is especiallynoticeable near the wing apex, or at the side- of-fuselage station if there is a fuselage. As a resul%, QI ccnstraintFrovisions are providedin the wing design input.
As many as five ordinateconstraints can be applied at arbitrary locations on thewing planform, provided ‘that each theof constraints be on a camber-calculationline, Theprogram tests the spanstations of the ordinate constraints and shifts them to thenparest calculated camber lineif they do not lie on one, (This hasbeen done to avoiddifficulties with two-dimensional inteqolation near the side-of-fuselage).
The Frogrannext checks constraint planform locations chordwise. If anyccnstraints are downstream of thetrailing edge, they are moved to the trailingedge, If anyare ahead of the leadingedge, the case is StOpFed. When crdinate ccnstraints are used,several cases shouldprobably be run. For thefirst case, both the maximum numberof component loadingsbeing considered and the maximum number of ordinate constraintsbeing considered should he used, and a RESTART deck should be generated,Given RESTART capability(described on page 35 a reduced pumber ofordinate constraints may be imposed. Any number ofordinate constraints can he deleted, startingwit.h the last one,(The ordirate constraint order cannot be changed, however).Cases without ordinate constraints can also be runfrom a’EESTART deck whichincludes the componentloading ordinatc data.
Giver. this logic, some thoughtshould he devoted to theorder in which ordinateconstraints are applied. Difficulties possibly requiringordinate constraints sometimes arise at thecenterline of a wingalone, -at theside-of-body wing station, and at wing stationshaving substantial change of sweep angle. One would be tempted,for example, to imFoseordinate Constraints at, say, 65 percent, 35 percent,and 85 percent chord ofthE side-of-body wing station, leaving two ordinateconstraints free for usefurther outbcard. It wouldthen be feasible,using RESTART, to run a case with oneordinate constraint at 65 percent chord on the side-of- bodystation, another case with two ordinateccnstraints at 65 and 35 percent,and so on,
The crdinateconstraint capability has one cther importantfeature with resFect to RESTART capability.Although planform location of ordinateconstraints carmot be changed in a RESTAFT deck, the constrainedvalues of ordinatecan be changedfrom case to case.
33 Ingeneral, ordinate constraints should be usedsparingly, since theycompromise the rumber of comgonentloadings available for dragninimization,
””””“””””“”Scluticn cver-ccns,traint.. - There are six tyFesof constraints that can be imposedon the wing design optimization. Constraints can ke inposed on:
(1) Lift.coefficient
(2) Pitching moment coefficient at zerolift
(3) Body buoyancy, body upwash,and nacelle buoyancy loadings
(4) Pressuregradierit on thewing upper surface
(5) Pressurelevel on thewing upper sucface
(6) Camber (E) ordinates
It is certainly Fossible to specifyan over-ccnstrained solution ” 5 camberordinate constraints for a casecombining three component loadings, for example,Consequently, a test has been
placed ’ in thefirst part of theoptimization program OPTIMTM to detect and correct this situation.The test first sums the number of ccnstraintstypes (1), (2), (3) , and {6) above. If the sum is eithfr greaterthan or equal to the pumber of componentloadings, the number of ordinateconstraints is reduced so thatthe sum is one less than the number of loadings. If +he altered number of .ordirateconstraints is negative,the program halts; otherwise it proceeds nornally.
&gaiihg--ne~~ct&a~z- The loadingdefinitions used in the program are tabulated on Faye 24, consisting of both analyticallydefined andconfiguration - dependenttype loadings. These may be input in anyorder, subject to thecondition that the camber-qenerating version (12- 14) of theconfiguration dependent loadings may notbe used without.also using thecorresponding configuration dependent loading (15-17).
Experiencewith the wing designprogram has shown that conbinations of thehigher order X term loadings (i.e:, the X term loadingsother than loading 2) tend to produceexcesslve twist or wavinessin the calculated camber surfaceunless constraints are imFosedon wing uFper surface pressure level andgradient.
I% is always good practice to runthe wing alone with threebasic loadings(ur,iform + linearspanwise f linearchordwise) in addition to any more sophisticatedwing design case. Although the three term casehas little capabilityfor handling multiple constraints, it serves as a check on theaverage pressure gradient
34 the wingcan be designed to, in addition to providing a quick aFFrcxirmticn to the optimumwing shape.
~?gst?gg-ce&ig~. - k "restart" optionhas besn providedin the program to minimizecomputer time on runsinvolving the same planformand Nach number. (i.e., differentdesign points in terms of CL t Cmo ordinate, or pressureconstraints) . The restart option works as follows: For a givenwing planform, Mach number, and set of loadings, most of thecomputer time is used in calculating the force coefficientsand interference coefficients associatedwith all the comFoneRt loadings.The calculations involvingthe solution of an optimumcombinaticn of loadings, with or withoutconstraints, are relatively quick, However, it may be
desirableto look at a nunberof ' differentoptimization or constrair,tsolutions. Therefore, on successive C~SPSinvolving the same basicloadings, it is possible to byFas3the component lcadingssolution dnd godirectly to theoptimization routines. This is done by setting RESTART= - 1. inthe program input for cases 2 and or..
If the Frcgram cases. are to b? inputat a later time, the corrpcr.entloadings data may be punched ir.+o cards(using RESTART= 1.) andread back in to thecompter through use of ZESTART= 2. The RESTART=2. datadeck includes, as well, the definition of anycorif iguration-dependent loadings that were presEnt in the wicg designprogram at the time thedata deck was puncked.
P.ZSTART=3.0 is a specialprovision in which the restart data are written onto a tape, which may latsr be reread in the wingdesign program. This feature is usefulin cases where the liftanalysis prcgran! may be runbetween successive wing designs. (RESTARP3.0 actuallyfunctions much the same as RESTART= -I., hut PESTART= - 1. is intendsd for useonsuccessive wing design cases without exitingthe wing design program).
The restart option also will work inthe case of a decreased nurrberofloadings. E.g., if a maximum (17) loading case were run,then thF forceand interference loading terms fcr all lesser corbinations .of loadings are available.Successive cases can then be run with differsnt loading combinations to check the design sensitivity to certainloadings, without repeating the basic loadingscalculations. Any combination of loadingsinvolving the set used when the RESTZ4S.T data were generated may be employed. The number and the order of the loadings may be altered as desired.
If the EESTART datadeck is used, it is notnecessary to recalculateany configuration-dependent data (since these are preservedalong with the basicwing loading data). With respect
35 I to the wing design5OlUtiOn that was possible at the time the RESTART data were generated,the FiESTART deck may be used for any wing design having:
1) The same or fewerloadings (order nayvary)
2) The same or fewer ordinate constraints(order cannot vary) . The value of Z at theselocations can be changed,however,
3) Same fuselagegeometry, anqle of attack,and side of fuselage station.
4) Any CL, Cmo, or pressureconstraint condition. glar.f_cg~considerations__a_nd spanwise integratiog. - The wing designprogram is a directtype solution, i.e., a wingshaFe is calculat2dfrom a known pressuredistribution. It is not necessary to calculatethe wing shape at all sFanwisestations in thegrid systen! used to representthe wing; only a representative set cfspanwise stations is, used.The lift,drag and pitching moment coefficientsare then computedfrom a spanwiseintegration of the characteristicsobtained at the selectedspanwise stations.
In the prograrinput, the camber surfacecalculations are performd at a standard set of 23 semi-spanstations unless otherwise specified. If the planform is irregular, particularly alcng the leadingedge, additional spanwise stations +hein viciaityof these irregularitiesshould be input to improvethe soluticnaccuracy. (This is donethrough inputs TJBYMX and TJEYS, as described ir! Section 4.)
Pr, addition, it hasbeen four,d that, for a basic wing cas€, the wirg root singularity aFd thecorresFonding root camber line can oftisr! be moderated by substituting a parabolicapex for the sharp apex coIsmon tosupersonic wing planforms. This will be performed autcrratically in the program if the input YSN00T is notzero. The program then fits a parabolatanger?.t to the wingleading sdge at YSNQPIT, with symIr;etry about Y=O.
Becauss the computedcamber surface slopes tendto exhibit some irregularity near the leadingedge (due to the sawtooth nature .of thegrid system), a sn!oothingoption is providedin tne program. This is activated by the code SMQQTH in the Frograrr inpt. The srrocthingtechnique involves averaging the computedsurface slopes of each grid e1emer:t with the slcpes of adjacent elements, which suFFrtlsses anyerracic slopes of individuale1ement.s.
36 Lift Apalysis
GiiJen a wingplanform, camber shape, and Mach number, thelift analysis program solves for the lifting pressure distribution and forcs coefficients for a rangeof angles of attack. As options, the Frogram will also includethe effects of: Fuselage(nominally circular in cross-section, arbitrary camber ar,d incidence) - Nacelles e Canardand/or horizontal tail Wing trailingedge flaps and/or incremental wing twist
"""Fuselage """"" soluticns, - Fuselage effectsare obtained by calculating the isolatedfuselage upwash field,then calculating the wing sclutioninthe preserxe ofthe fuselage upwash field,then calculatingthe fuselage forces in+he wing flow field, and combiningthe solutions by superposition,
The fuselage upwash field is calculatedfrom sler,der body theory. The input areadistribution of the fuselage is considered to be circularin cross-section. If a digitizedfuselage cross-section is inFutinto the basic geometry, the area and centroid of each section is comptedand used to definethe area 'and meacline distrikutionfor the analysis program. The liftanalysis program contains a wing-fuselageintersection oFticn.This feature tracks each wing percent chord line out throughthe side of thefuselaqe (again considered circular in cross-section), andbrzaks the wing solution into the proper exFosedandcarry-over typelifting Fressure calculations. Alternatively,the side-of-fuselage span station may beinput eithfr as a constant or as a tableof values to override the wing- fuselageintersection option.
Thelocal fuselage upwash angle is stror,glyaffected by span station andwipg height on the side of thefuselage. Theside-of- fuselagespan station must be carefullyiriput to avoidexposing any wing area to the upwash fieldthat is act.u.aliyinside the fus€lage.
The liftanalysis program cor.tains ap. option to calculatethr huoyaricy fielddue to unlqualfuselage area growth above and below the wing. This Fressuredistrihutioc, terrneu asymmetric fr;selage buoyancy, is calculated by splittingthe fuselage ar.ea intopisces abcveand below the wingand adding the resultant area growth onto thefuselage forebody area distribution.(The fuselage is again ccnsideredcircular, and +_he side-of-fuselage Z value is used to define the above-wir!gand belcw-wing area pieces). The asyrcmetric fuselage tlrm is z1ro, ofcourse, in the case of a mid-winq arrangement.
37 Theasynwetric buoyar,cy calculation is requested by input ' SYMN (valuegreater than zero). For a fuselagesignificantly non- circularin cross-section, use may be nade of two specialoptior_s to define the above-wingand below-wing area distributions and the correspondingwing-fuselage intersection:
SYMFI = 2.0 requiresinput of theabcve wing and below-wingareas.
0 ANYEOD = -10. allows inputof definition of the wing-fuselageintersection.
Both of theseoptions require input of thedata at khl same per centchords used in the camber surfacedefinition.
"""_Nacelles. - The nacelle calculations are very similar to the scluticnused in ths p+ar-field wave dragprogram. The pressure fieldsimposed by thenacelles on the wing,and wing-on-nacelles, are computedand their combirsd effect on the liftingsolu%ion oktainsdthrcugh superposition. The effect.of +he nacelles on the wixgdrag-due-to-lift can be substantialbecause of lift ccntributed by thenacelle pressure field. Both llwraplland llqlancelisolutions for the nacelleFressure field are available, as described inSection 3.6. Cazagd- andhorizontal-tail. - Canardand horizontal tail lifting pressuredistributions and force coefficients are calculated as for the wing case. Theprogram assunes that a canard is located forward of the wingand a horizontal tail aft of the wing,The effectsof doknwash from upstreamlifting surfaces (if any) are included ir. thesolution.
""""Downwash ,:shiftc' options. - The basic theoretical solution emFloyedfrcm canard or win9propagates directly aft. Sixe the dcwmashin the real flow case mustfollow the fuselagecon+our, a shiftfeature ic theprogram translates the downwash field laterally to accour?t . forfuselage radii change between a generating(canard or wing)and affected (wing or tail lifting surface. The downwasy shiftcan have an appreciable effect on the calculatedcharacteristics, as shown infigure 3.7-6.
The shift feature is controlled by inFut codes. If thecontrol codes are leftblank, the downwash will be shiftedaccording to theside-of-fuselage Y valuescanard,of wing, or tail. Alternatively,t.he downwash can be eitherunshifted, or shifted a specifiedamount, as described in +he datainput section (4.0).
"ExEerineEtal ""-"""" comEarisons.""" - Theoretical calculationsfor a typical suFerscnictransFcrt configuration are comparedwith corresFonding windtunnel data infigures 3.7-7 and 3.7-8 (wing-fuselage- nacelles)and figures 3.7-9 and 3.7-10 (incrementaleffects of
38 Legend
"-
I NO CANARD DOWNWASH SHIFT I
A cDcanard .10 */
wave + friction Am drag coeff.
OLP~l~ ~ I 1 -0- 0 .002 .004 .006 .008 .004 0
AC mcanard A 'mcanard
~~ ~ ~ WITH CANARD DOWNWASH SHIFT I ~-~ I Theory I
*" canard .0004 c
0L"J I I -I I ' 0- 0 .002 .004 .006 .008 .004 0 AC mcanard A 'mcanard FIGURE 3.7-6-EFFECT OF CANARDDOWNWASH SHIFT x M =2.7
39 P 0
.I 2
.I 0
.08
.06 THEORETICAL WAVE AND I\ /- SKIN FRICTION DRAG
CL Skin friction* Wing 0*00361 0.0023 Fuse lage 0.00185Fuselage } .04 Nacelles' * Ext 0.000280.00082 /IV Int 0.00086 \. * No allowance for transition @ W.B.N. i /o strip (grit) drag-believed to .02 I/ be approximately CD = 0.0001 I/ included drag internal** Nacelle !I in plot I1 0 I I I I I I I I 1 0 .010 .011 .012 .013 .014 .015 .016 .017 .016 .015 .014 .013 .008.012 .009.011 .010
CD
FIGURE 3.7-7. -DRAG POLAR COMPARISON Experiment Theory M = 2.7,Reynolds No. = 9.5 x 10 Nacelle off 0 - Nacelle on 0 "I
.12 .12
.10 .10
.a9 .08
CL cL
.06 .06
.04 .04
I .02 .02 i I
0 I 2 4 6 8 .008 .004 0 Obody, degrees Obody, C m0.60 Cr FIGURE 3.7-8.-FORCE COEFFICIENT COMPARISON P Horizontal Tail Setting Id - - - Theory -From experimental data -50 -30 -1 .o' +1.5' .I2
.10
.08
M = 2.7
Reynolds no. = 9.5 X 106 /ft
CL .06
.04 Tail setting = -5' .04 I .02
1 -6 0 -6 4 -.008 -.004 -.008 0 .004 .008 .012
deg cy body, Acm Note:Horizontal tail setting referred to wl. Angle ielative to wing z = 0 plane is 1.25 larger.
FIGURE 3.7-9.-HORIZONTAL TAIL EFFECTS r
Notes: 1. Lines are calculated values from lifting surface theory Experiment 2. Data points read from wind tunnel force plots 3. ACD, AC, are tail-on minus tail-off values.
0.1 0 0.1 2 M = 2.7
.0010
.0008
CL 0.06 .0006
.0004 1 Estimatedwave + friction drag of horizontal tail
.0002
.006 .004 .002 0 -.002 -.004 -.006 -.008 -.010
AC "'tail
FIGURE 3.7-10.-HORIZONTAL TAIL EFFECTS
43 horizontal tail) The theoretical buildup of the zero-lift drag coefficient is given in figure 3.7-7.
The lift analysisprogram contains an optional pressure limiting feature for the wing surface pressures which operates somewhat different fromthe one in tho design program. In the designcase, the lccal wingangle of attack is not allowed to exceedthe value associatedwith a Fressure limit condition. In theanalysis case, theFressure coefficient limit is imFosed,bu% thc local wing incidence may greatlyexceed the value at which a limit. is first encountered.
When thepressure limiting option is used, a set of configuration angles of attackfor the solution must be provided,and the configuration thickness pressures from the near-f ie&d programmust beFxcvided to Ferrnitlimiting of the total surfacepressure. A solution for a typicalwing through an arigle of attack series using the pressurelimiting feature is shcwn in figures 3.7-11 and 3.7- 12. The limitingfeature greatly improves the . lineartheory representationofthe wing pressure distribution as aqle of attack is ircreased.
Conr~quration-aeE~ndenthadir.crs. one mode of liftanalysis prcgramusage is to genlrate corifiguration-aependent datafor the wingdesign program. These data are producedas follows:
""CATA I?Z_SC_RIIPTFGN EELUIREMENTS Nacellepressure field Pressurefield caused Call fornacelles by nacelles on winq. (AJ3=1.0) Fuselagespwash field Pressure field induced Calculate fuselage onwing byfuselage effects or: wing UFWaSh.
Fuselag? buoyancy fieldPressure field induced SYiW .. 0. on wing by unequalfuse- lagevolune abcve and be- low wing.
UFcn execution,the program *hen lcadsthe Fressure fields into the _rrcFersysterr common blocks.
If tks fuselagebuoyancy field is notrequested (i.e., SYMN = O.), the program cornFut.es theFressure field due to a mid-wing arrangement.This i~ done so that a thicknesspressure field due to thefuselaae will be availablepressurefcr limiting calculations,if desired.
Ic calculatingthe fuselage upwash or buoyancyfields, it is important to remember thepowerful influence of wingheight on the
44 -M = 2.05 Experiment (NASA TND-1264,wing 2) Numerical method Upper surface No C limit 0 - P "- 0 Lower surface "" Limit at 0.7 vacuum (Co = -0.24)
b yI7 = b 0.1 yJ?= 0.5 y/$ = 0.7 yl; = 0.9 ~ I I C, = 0.13 -.2 c -.2 1 -.2 cp 0' 0
.2 .2Or 101=6"1 kexp= 0.253 -.4 t- -.4
"""" c 1: ~ooo~ooo -2 -2 -.4 "- 0
.: .: .: .2 cp-.2 OP
" 111111 0- .2 .4 .6 .8 1.0 0- .2 .4 .6 .8 1.0 0 .2 ..4 .6 .8 1.0 0- .6.4 .2 .8 1.0 0. .2 .4 .6 .8 1.0 XJC XJC XIC XJC XJC FIGURE 3.7-11.-PRESSURE COEFFICIENT COMPARISON- WING 2 (TWISTED AND CAMBERED WING. CI = 0.08) Experiment (wing 2, NASA TMX-332) ITwisted and cambered wing I -Theory, no Climit C = .08 P 1 x- -Limiter, C limit = 0.70 vacuum Ldes P .28 M =. 2.05
= .0081 'Dsym
.24
.20
'L .16
.12
.08
.04
0 .008 .016 .024 .032 .040 -2 .040 .032 .024 .016 .008 0 2 4 6 .02 0 -.04 -.08
Drag due to lift = CD-CD 'C ' svm m0.25E FIGURE 3.7-12. -TEST-THEORY COMPARISON, WING 2 f C, des = 0.08) side of thefuselage. This stronglyaffects bo+h the local upwash angles, ar,d theabove-and-helow wing area distributions.
Calculation of thefuselage upwash field may be rlor,e ir. either of twc ways; the Frir.cipa1condition is that theresultant Fressure field is that due t-o upwashonly. In the compter program, this is handled by inputting a camber surfacehaving approximately the correct wing-fuselagerelationship (wing height, etc.) , but then zeroingthe wing slopos in the cambersurface calculations (by setting WHIJP=l .O) . Ir itera%ivecycles, the wing carriber surface andfuselage relationship can be refined.
Alternatively, as a crudsstartirig p0ir.t in the fuselageupwash calculation, the flatwing option can be used. By settinq TIFE C=2.0, thewing slopes areautomatical.ly zeroed ar,d the wing height relative to %he fuselage will be controlledby the fuselage meanlineinput and the wingleading edge 55 definition (%LED and ZFUS in the &sic geometry).
4.0 INPUT FORYAT
Inputrequirements for the system are givenin this section and consist of: Executive control card summary Basic geometrydefinition .* Additional data input for programs of system
The usualinput format is 10 field - 7 digit, punched with decimals to the left in thecard fields. Some data (particularly thecontrol codes in the basic geometry) are inputin integer form,without decimal, t.0 theright in the card field. The formats are identified in all cases. To providedesign or analysisflexibility, there are numerous programoptions that are controlled by input codes. Where there is a @fnormallfway of handling the option, the code is defaulted to zero (i.e., if thefield contains a zero or is blank, the ffnormaltl solution will be calculated).
**NOTE**
The allowable input number sizes are: * No more than 5 digits to left of decinal a No more +.ban 4 digits to right of decimal
4.1 Executive Control card Surnrrary
Configuration input andprogram execution are crdered by means of controlcards read at the executive level. Thecontrol cards consist of a few alphanumericcharacters starticg in column 1.
Geometryinput.. - The configurationgeometry is read and manipulatedin the geometry module. Geometry my be input as all- new, or as a replscement or addition to existing geometry. The control cards for geometryinput are:
GE0M NEW All-newconfiguration description follows, andany previous geometry is purged. (Leave one column space between GEBM and N5W) . Inputgeometry is added to (or replaces) existingdescription.
49 Geometa""" update. - The basic geometry description contained in the geometrymodule may be updated using data contained in 0, 0 level common blocks. This applies to a new fuselagedefinition (io@., optimizedfusdage from thefar-field wave drag program) or a new wing camber surface definition. The control cards are: FSUP Fuselage will be updated tc definition contained in /QPBQD/. The /OPBQD/ definition is created each time the far- field wave drag programexecutes the optimum-fuselage-with-restraints case. If the fuselage update is requested, a second card,telling how to perform the update, is required,Punch (starting in column 1) the f ol lowing code:
-1. Fuselage is to be redefined at same x stations as previous definition.
1. Fuselage is tc be defined at 50 equally spacedstations.
WGUP with camber surface will be updated to thedefinition containzd in /CAMBEk/. The /CAMBER/ definition is createdeach time thewing design program executes .andpro- duces a cambersurface for a specified set of conditions.
Theuser must remember that the updatefor fuselage or camber surface- will require that tk.3 /QPSQD/ or /CWi3E€?/ definition be current. These corcirron blocks will contain the last definition produced by the far-fipld wave drag or wing designprograms. ?_rsqram ?xesg;jo_n. - Execution of theprograms in the system is ordersd by the following cards:
PL0T p 1ct program
SKFF skin friction program
FFWD far-field wave drag progran
NFWD near-field waveprogram drag
APU'LZ lift an3lysisprogram
WCEZ wing programdzsign
The controlcard for proqramexecution is the first card of the set describing the programdata input. Individual programinputs - areqiven on the followingpages.
50 IvIultiple case execution with the basic programs of the system is possible, as in the stand-aloneversions of theprograms. The datafor successive cases are stacked as ii2scribedin the program inputdescript.ion. At the end of the datastack, an END card is required to terminate the program. The END card is notused for the geometryinput., however.
""""""Interactive ~~g&~hlhs. - The graohics subroutines in the systemare activated by the executive card CRT (punched infirst three card co1urr;ns) . The CRT card may be placedanywhere in thedata deck that an executivecard may be read. If no CRT card is includ2d, the system will 2xecutewithout accessing any of thegraphics programs. h description of the interactivegraphics part of the designand, analysis systr-em is presentedin Appendix A.
4.2 Geomet-ryProgram
The geometryprogram stores the basic geometry data,and stacks it as required by the individualprograms of the system. Access to the gecn?e+ry program, to store or alter the configurationdescription, is through the GEQM or GEQM NEW control card (see execu+ivecontrol card summary)
'The format of thegeometry input uses both integer (control cards) andfloating point numbers, All integers are punchedright justifiedin their fields on the cards,without decimals. All floatingpoint numbers arcpunched, with decirrals, to the left of the field in 10 field -7 digitrormat. The prcqram logic uses the component. controlcodes (Jl, 52, etc.) on card 3 asfollows:
0 Component will not be input.However, ifthe cornpnent haspreviously been input (and not purged by a GEQM-NEW card), %he 0 is inter- preted as a 2,
2 Previously input component is left as is. Other New input for this componentreplaces previousinput. Thelogic of treating a 0 as a 2 for existingcomponents is to protectdata on thegeometry file from inadvertent loss. Then, if it is desired to add or change a configurationcomponent on successiveruns, only the new componentneed be addressed.
51 A control code other than 0 or 2 instructs the program to completely rep1.ace the previous comFonent description with a new one. It is not possible to add a fir, or nacelle to a previous fin or nacelle; the new description must be complete in itself. Deletion of a component is possible only through purgifig the entire configuration,using the GEBN NEW card.
Card Card Decimal Variable Description Number Column Required Name 1 GE~Mor GE@M NEW 1-4 GE@M = geometry addition 1-8 GE@M NEW = all-new geometry
2 1-70 Any desired title information.
3 1-3 NO JO Referencegeometry code.
0 = Reference geometry not required (plot program) 1 = Read reference area, -c, xcg 2 = Referencegeometry same as previous case. 3 4-6 NO J1 Wingcode input
-1 = Read uncamberedwing 0 = No wing 1 = Read camberedwing 2 = Wing same as previouscase,
3 7-9 NO 52 Fuselagecodeinput
-1 = Read circular fuselage 0 = No fuselage 1 = Read arbitrarily shaped (digitized) fuselage 2 = Fuselage same as previouscase 3 = Read circular fuselage and perimeter values.
3 10-12 NO 53 Nacelle input code
0 = No nacelles 1 = Read nacelles 2 = Nacelles same as previouscase.
52 Cad Card Decimal Variable Description Number Column Required Name
3 13-15 NO 54 inputFin code
0 = No fin 1 = Read fin data 2 = Fin data same as previous case.
3 16-18 NO 55 Canard (or horizontal tail) input code
0 = No canards 1 = Read canard data 2 = Canards same as previous case.
3 19-21 NO .56 FuselageSimplification code
-1 = Uncambered circular fuselage 0 = Cambered circular or arbitrary fuselage. 1 = Complete configuration is symmetrical with respect to X-Y plane, which implies uncambered circular fuselage if there is a fuselage.
3 22-24 NO WAFNumber of airfoilsdescribing wing. 2 -L WAF - 20. 3 25-27 NO NWAFOR Number of ordinatesdefining each airfoil section. 3 NWAFOR -L 20.
3 28-30 NO NFUS Number fuselageof segments. 0 = NFUS -L 4.
3 31-33 NO NRADX(1) Number of points defining half section of first fuselagesegment. If fuselage is circular, the program calculates the indicated number of Y and Z ordinates. 3 NRADX(~) A 30.
3 34-36 NO NFORX(1)Number of stations for first fuselage segment. 4 NFORX(~) 4 20.
3 37-39 NO NRADX(2) Same as above for segment2.
3 40-42 NO NFORX(2)Same as above for segment 2.
53 Card Decimal Variable Caxd Description Number Column Required Name
3 43-45 NO NRADx(3) Same as above for segment 3. 3 46-48 NO NFORX ( 3) Same as above for segment 3. 3 49-Sl NO NRADx( 4) Same as above for segment 4. 3 52-54 NO NFORX( 4) Same as above for segment 4. 3 55-57 NO NP Number of nacelles to read. Np 4 3. 3 58-60 NO NPOD OR Number of stations at which nacelle radii are specified. 4 L- NPODOR L- 20. 3 61-63 NO NF Number of fins to read. NF L- 6.
3 64-66 NO NFINOR Number ofordinates defining each fin airfoil section. 3 6 NFINOR 10.
3 67-69 NO NC AN Number of canards to read. NCAN L- 2.
3 70-72 NO NCANOR Number of ordinatesdefining each canard airfoil section. 3 L- NCANOR 10. If negative, airfoils are non-symmetric.
4 1-7 YES REFA Wing referencearea 4 8-14 YES CBAR Pitching moment referencelength. (Required for ANLZ Bnd WDEZ only)
4 15-21 YES XBARIN X value of pitching moment center (Required for ANLZ and WDEZ only)
Note: Omit thiscard if JO (Card 3) is 0 or 2.
54 Wing Description
Card Card DecimalVariable Description NumberColumn Required Name
Omit card sets 5, 6, 7, 8 and 9 if J1 is 0 or 2.
5 1-70 YES XAF Array of percentchords at whichwing airfoil ordinates will bespecified.
6 1-7 YES XLED X coordinateairfoilof leading edge, 6 8-14 YES YLED Y coordinateairfoilof leading edge, 6 15-21 YES ZLFD Z coordinateairfoilof leading edge, 6 22-28 YES CLED Airfoilchordlength
Note : Thiscard is repeatedfor each airfoil, ordered inboard to outboard.
7 1-70 YES TZORD Array camberof Z valuesreferenced to Z coordinate of airfoilleading edge, orderedleading edge to trailing edge.
Not e : This card is repeated for each airfoil, ordered inboard to outboard. Omit cardset 7 if wing not cambered.
8 1-70 YES WAFORD Array of airfoil u-pper surfacehalf thicknessordinates expressed in percent chord,ordered leading edge to trailing edge.
Note : Repeat Card Set 8 for each airfoil, ordered from inboard to outboard.
Not e : Card Set.9, an option in the plot program input to define .the lower surface.airfoi1 for an asymmetric airfoil shape, was deleted from the basic geometry to reduce core size,
55 FuselageDescription
Omit card sets 10-15 if 52 is 0 or 2. Thefuselage is input in segments.Complete input for each segmentbefcre going on to next segment. A segment may contain 520 definingstations. First segmentmust begin at x=O. If there is more thanone fuselage segment, the firststation of a segment repeats the definition of the laststation of the preceding segment {i.e., cross-section is againdefined at the same X station) . Otherwise, a gap in the fuselagedescription will occurbetween the last station of one segment. andthe first station of the followingsegment. Make the first x value of the succEedingsegment- slightly larger than the last x valu? of the precedingsegment.
Card Card Decimal Variable Description Number Column Required Name
10 1-70 YES XFUS Array of fuselage X stations
11 1-70 YES ZFUS Arrayof Z coordinates defining fuselageceriterline.
Note : Omit card set 11 if 56 f 0 or if 52 = 1.
12 1-70 YES FUSARD Array of fuselagecross sectional areas .
Not e : hit card set 12 if 52 not equal to -1 or 3.
13 1-70 YES FUSPER Arrayfuselageof perimeters.
Not e : Omit card set 13 if J2 not equal to 3.
14 1-70 YES SFUS Array of Y coordinatesdefining first station half section, ordered bottom to top.
15 1-70 YES SFUS Array of Z coordinatesdefining first station half section, ordered bottom to top.
Note : Repeatcard sets 14 and 15 for each station in segment 1. Omit card sets 14 and 15 if 52 is not equal to 1.
Not e : Foreach fuselage segment, repeat card sets 10 thru 15.
56 Nacelle Description
Card Decimal Variable Card Description Number Column Required Name Omit card sets 16, 17 and 18 if 53 is 0 or 2.
16 1-7 YES PODORX X coordinateoriginof of first nacelle
16 8-14 YES PODORY Y coordinateoriginof of first nacelle
16 15-21 YES PODORZ Z coordinateoriginof of first nacelle
16 22-28 YES POD ZW Z coordinateoriginof of first nacelle, referenced to lbcal wifig surf ace.
0. , program will calculatefrom POD OR Z +D, nacelle is located D unitsabove local wing surface -D, nacelle is located D units below local wing surface
Note: If PODZW f 0. , PODORZ is notrequired.
17 1-70 YES XPOD Array of X coorLinates,referenced to nacelle origin, at which nacelle radii will be specified.
18 1-70 YES RPOD Array of nacelleradii.
Note: For eachnacelle, repeat card sets 16 thru 18. If PODORY is non-zero, a duplicate nacelle is located symmetrically to the X-Z plane.
57 Fin Description
Card Cad Decimal Variable Description Number Column Required Name
Omit card sets 19, 20 and 21 if J4 is 0 or 2.
19 1-7 YES X coordinate of lower fin airfoil leadingedge.
19 8-14 YES Y coordinate of lower fin airfoil leadingedge.
19 15-21 YES Z coordinate of lower fin airfoil leadingedge.
19 22-28 YES Chordairfoillower oflength
19 29-35 YES X coordinate of upper fin airfoil leading edge.
19 36-42 YES Y coordinate of upper f,in airfoil leadingedge.
19 43-49 YES Z coordinate of upper fin airfoil leadingedge.
19 50-56 YES Chord lengthupper of airfoil.
20 1-70 YES XFIN percentArraychords, ofordered leadingedge to trailingedge, at which fin airfoil ordinates will bespecified,
21 1-70 YES FINORD Array of airfoilfin half thickness ordinatesexpressed as percentchord.
Note:Repeat card sets 19 thru21 for each fin.
58 Canard (Gr Horizontal Tail) Description
Card Card Decimal Variable Description Number Column Required- Name Program identifies horizontal tail or canard by location relative to wing. .Omit card sets 22-25 if 55 is 0 or 2.
22 1-7 YES X coordinate of inboard canard airfoil leading edge.
22 8-14 YES Y coordinate of inboard canard airfoil leadingedge.
22 15-21 YES Z coordinateof inboard canard airfoil leadingedge.
22 22-28 YES Chord lengthinboardcanardairfoil. of
22 29-35 YES X coordinateof outboard canard airfoil leadingedge.
22 36-42 YES Y coordinateof outboard canard airfoil leadingedge.
22 43-49 YES Z coordinp te of outboard canard airfoil leadingedge.
22 50-56 YES Chord length cf outboardcanardairfoil
23 1-70 YES XCAN Arrayof percent chords,ordered leading edge to trailing edge, at whichcanard airfoil ordinates will be specified.
24 1-70 YES CANORD Array of canardairfoil upper surface half-thicknessordinates expressed as percent chord, ordered leading edge to trailing edge.
25 1-70 YES CANORl Same as above for lowercanard airfoil Note: If canard is symmetric, omitcard set 25.
Note: For eachcanard, repeat card sets 22 thru 25.
59 4.3 Plot Program
This program draws a picture of the configurationdefined in the basic geometry, as requested by the codes on card 3. Views of %he configuration are controlled by the inputs on card 4. There will be 'as many drawings of the configurationas there are cards 4. Three different types of card 4 inputs are possible, for orthcgraphic, prspective, cr stacked three-view options.
Card Card Decimal Variable Description -Number Column Required Name 1 1-4 PL~T
2 1-80 Any desired title information.
3 1-7 YES AJ1 Wing inputcode.
0. = Ignore wing definition. 1. = Include wing definition.
3 8-14 YES A 52 Fuselageinput code.
0. = Ignore fuselage definition. 1. = Include fuselage definition.
3 15-21 YES AJ3 Nacelleinputcode.
0. = Ignore nacelle definitions. 1. = Include nacelle definitions.
3 22-28 YES AJ4 inputcode.Fin
0. = Ignore fin definitions. 1. = Include fin definitions.
3 29-35 YES A 55 Canard inputcode.
0. = Ignorecanard definitions, 1. = Includecanard definitions.
For OrthographicPro.jections 4 1 HOR Z X, Y, Z forhorizontal axis. 4 3 VERT X, Y, or Z for verticalaxis.
4 5-7 TEST1 @UT if deletion of hiddenlines required;otherwise blank.
60 Cad Card Decimal Variable Description Number Column Required Name
4 8-12 YES PHI Roll angle in degrees. 4 13-17 YES THETA Pitch angle in degrees.
4 18-22 YES PSI Yaw angle in degrees.
4 48-52 YES PLOTS 2 Length in inches of maximum config- uration dimension.
4 53-55 Punch @RT in these columns
For Perspective Views (See fig. 4.0~1)
4 8-12 YES PHI X ofeye point in data coordinate system.
4 13-17 YES THETA Y ofeye point in data coordinate system.
4 18-22 YES PSI Z of eye point in data coordinate system.
4 23-27 YES XF* X of focal point in data coordinate system.
4 28-32 YES YF* Y of focal point in data coordinate system.
4 33-37 YES ZF* Z of focal point in data coordinate system.
4 38-42 YES. DIST Distance from eye point to view plane in inches.
4 43-47 YES FMAG View plane magnification factor. Controls size ofprojected image.
4 48-52 YES PLOTS Z Diameter of view plane in inches. DIST and PLOTSZ determine a cone which is the field of vision,
4 53-55 TYPE The letters PER
* Inside configuration
61 eye polnr - I' / PLoTSZViewing plane -FMAG controls image size
FIGURE 4.0-1."PERSPECTIVE VIEW INPUTS For___"" Plan, FYont and Side. Views (Stacked)
Cad Card Decimal Variable Description Number Column Required Name 4 8-12 YES PHI Ordinate of viewcenterlineplan on paper(in inches).
4 13-17 YES THETA Ordinate of side viewcenterline on paper(in inches).
4 18-22 YES PSI Ordinate of front centerlineview on paper(in Inches).
4 48-52 YES PLOTSZ Length inches)(in of maximum plot dimension.
4 53-55 TYPE The letters VU3
Note:For each additional plot desired, card 4 will be repeated at this position in the data deck.
5 1-3 END
4.4 SkinFriction Program
codes on card 3 cont-rolinclusion of basicgeoretry as requested. Where additional input is required(e.g., fuselage perimeter option),input areas or lengths in unitsconsistent with the basic gecmetrydefinition.
Theskin friction coefficient subroutine in the programrequires lengths in feet.'The input lengths are ccnverted to feet, if necessary,usinq the factor SCAMOD on card 5 or 6. Inputson card 3 are integers,and must be right- justified ir, the field, withoutdecimal. The otherinput are 10 field -7 digit format,with decimals.
Decimal Card Card Variable Description Number Column Required Name
1 1-4 SKFR
2 1-70 Any desired TITLF: information
63
I Card Card Decimal Variable Descr,iption Number Column Required , Name
3 1-3 NO Jl Wingcode input
-1 = Wing defined in basic geometry. Make no correction for wing- fuselagejoint. 0 = No wing defined. 1 = Wing defined in basic geometry, Subtract wing root area from fuselage wetted area. 2 = Wing same as precedingcase.
3 4-6 NO J2 Fuselageinputcode
-1 = Wetted area and reference length will be input. 0 = No fuselage defined.
1 = Fuselage defined in basic geometry. 2 = Fuselage same as precedingcase.
3 7-9 NO 53 Nacelleinputcode
-1 = Wetted area and reference length will be input, 0 = No nacelles defined. 1 = Nacelles defined in basic geometry. 2 = Nacelles same as precedingcases.
3 10-12 NO 54 inputcodeFin
-1 = Fins defined in basic geometry. Make no correction for fin- fuselagejoint. 0 = No fins defined. 1 = Fins defined in basic geometry. Subtract fin root area from fuselage wetted area. 2 = Fins same as precedingcase,
64 Decimal Variable Card ,. Card Description Number Columns Required Name
3 13-15 NO J5 Canard (or horizontal tail) inputcode -1 = Canards defined in basic geometry. Make no correction for canard- fuselagejoint, 0 = No canardsdefined. 1 = Canardsdefined in basic geometry. Subtract canard root area from fuselage wetted area. 2 = Canards same as precedingcase.
4 1-7 YES AKI Mach number-altitudecombination code.
-AKI = Combination same as preceding case. 0 = Use Mach number-Reynolds combinations. AKI = Number ofMach-altitude combin- ations. AKIg 20.
4 8-14 YES AK4 Mach number-Reynoldscombination code.
-AK4 = Combinations same as preceding case. 0 = UseMach number-altitude combinations. AK4 = Number ofMach-Reynolds combin- ations. AKb 4 20.
4 15-21 YES AXTPT Miscellaneouscomponents code.
-AXTPT = Same components as preceding cas e. 0 = No miscellaneouscomponents defined . AXTPT = Number ofmiscellaneous components. NXTFT 10.
4 22-28 YES POVLP Totaloverlap nacellesareafor Subtract from wing wetted area.
5 1-7 YES AM Mach number
5 8-14 YES AL Altitude(feet /1000.)
5 15-21 YES DELT Temperaturedeviation standardfrom day (OF)
65 Card Decimal Variable Caxd Description Number Column Required Name
5 22-28 YES SCAMOD Scalefactorconvert to input dimensions to feet.
Note : There will be AKI of card(s) 5, Omit card set 5 if AKI is 0 or negative.
6 1-7 YES AM Mach number -6 6 8-14 YES RNPFL Reynolds Number per foot length x 10. 6 15-21 YES SCAMOD Scalefactorconvert to input dimensions to feet.
6 22-28 YES TOTEM temperatureTotal (OR) Note:There will be AK4 of card(s) 6. Omit card set 6 if M4 is O or negative,
7 1-7 YES SWETRB Fuselagewetted area
7 8-14 YES FUSL Fuselagereferencelength.
Note: Omit card 7 if 52 is 0, 1 or 2. 8 1-7 YES SWETNA Totalnacellewetted area
8 8-14 YES TODL Nacellereferencelength.
Note: Omit card 8 if J3 is 0, 1 or 2.
9 1-7 YES SWETXP Wetted areamiscellaneous of component. 9 8-14 YES RXLP Referencelength of miscellaneous component.
9 15-24 AnyPTITLE desired title information, Note:There will be NXTPT of card (s) 9. Omit card set 9 if NXTPT is 0 or negative.
Foreach new case, addCards 2 through 9 at this position in the data deck.
10 1-3 END
66 4.5 Far-Field Wave Drag Program codes on card 3 controlinclusion of basic geometrydata as requested. The case nurrber in first field of card 4 is an integer, and mu'st be right justified in thefield, without decimal, Other input are in 10 field, -7 digit format.
If the fuselagerestraint feature is used, the resulting fuselage definition for the last case will be storedand can be used to updatethe basic geometry (see executive control card summary, FSUP) . Multiple cases involving a givenconfiguraticn description {e.g., various Mach numbers) may be run by a card 4 series. If the geometry is to be changed, an END card must be input and the program re-entered by an FFWD or G%IM and FFWD set-up.
Card Card Decimal Variable Description Number Column Required Name
1 1-4 FFWD
2 1-80 information. Anytitle desired
3 1-7 YES AJ1 Wingcode. input
0. = Ignore wing definition. 1. = Include wing definition.
3 8-14 YES AJ2 code.Fuselageinput
0. = Ignore fuselage definition. 1. = Include fuselage definition.
3 15-21 YES AJ3 codeNacelleinput
0. = Ignore nacelle definitions. 1. = Include nacelle definitions.
3 22-28 YES AJ4 code.input Fin
0. = Ignore fin definitions. 1. = Include fin definitions.
3 29-35 YES AJ5 Canard (or horizontal tail) inputcode.
0. = Ignore canad definitions. 1. = Includecanard definitions.
67 Case Cards
Cards 4 input a series of cases of different Mach number, cut or theta variablss,and/or fuselage restraints, The solution is perfcrmed with the fuselage 2s input, and also for an optimum fuselage shape (subject to restraint points at which the fuselage shag€ must be as inpu?). If no fuselagerestraint is specified (NREST = 0.) , one will he assumed at the station of maximum overall area. Do notir,put restraint stations at nose or aft end of fuselage (those are automaticallyassumed). If NHEST>O., a restraintcard (card 5) will followthe case card, andthat restraintcondition will applyfor subsequer,t casesif NFEST is notchanged.
Card Card Decimal Variable Description Number Column Required Name 4 1-4 NO NCASE Case identification (right- justified) 4 8 -14 YES XMACH Mach number
4 15-21 YES Nx Number of equal intervals into which the portionof the X-axis, XA to XB for each roll angle, is to bedivided. NX 100. and an even number.
4 22-28 YES NTHETA Number of equal intervals into which the domain of theta (-90° to 90G) is to be divided, NTHETA 36. , and a multiple of 4. If the area distribution at -90° only is desired, input NTHETA = 1.
4 29-35 YES NREST Number of X stations for fuselage restraintpoints (=IO.), used for all subsequentcases if NREST doesnot change. If NREST = O., programassumes restraint points at nose,base, and station of maximum overall area.
5 1-70 YES XREST Arrayof fuselage stations, (including noseand base) at whichcomputed minimum dragcurve will be restrained to input area.
.Note:Repeat card 4 foreach new case. Only 1 card 5 may be input, after first card 4 with NREST # 0.
6 1-3 END
68 4.6 Near-Field Wave CragProgram
Codeson card 3 controlinclusion of basic geometryas requested.
Two optionsare provided for fairing the wing sectionshape at a givenspanwise station: linear or second order, controlled by TNOPCT on card 4- The code ANYBOD (on card 5) identifies the spanstation of the inboardend of the wing for calculatingwing thickness pressures and wave drag. This is the y value of the wing-fuselage intersection if there is a fuselage.
Card Card Decimal Variable Description Number Column Required Name
1 1-4 NFLII
2 1-72 Any desired TITLE information.
3 1-7 YES AJ2 Fuselageinput code.
0. = Ignore fuselage definition. 1. = Include ,fuselage definition.
3 8- 14 YES A 53 Nacelleinputcode.
0. = Ignore nacelle definitions. 1. = Include nacelle definitions.
3 15-21 YES A J4 codeinputFin
0. = Ignore fin definitions. 1. = Include fin definitions.
3 22-28 YES A 55 Canard codeinput
0. = Ignorecanard definitions. 1. = Includecanard definitions.
69 Card Card Decimal Variable Description Number Column Required Name
4 1-7 YES TNOPCT code,Fairing
-1. = Linearchordwise fairing, 0. = Second order fairing.
4 8-14 YES XM Basic Mach numbercase.this for
4 15-21 YES TNOM Numberadditional of Mach numbers. TNOM 4 5. 4 22-28 YES D ONT Wing printoutdata code.
0. = Minimal printout. 2. = Thicknesspressure coefficients at each grid element in the wing calculations will beprinted. 101. = Velocity potential will also be printed.
4 29-35 YES TNON Numbersemi-span of element rows in wing calculations, TNON 6 40. If blank, TNON set to 40.
4 36-42 YES TJBYMX Numberspanwise of stations at which wing thickness pressures are calculated. TJBYMX 24. Leaveblank if TNON not specified.
4 43-49 YES TNCUT Numberbody of stations at which pressure coefficients are calculated (‘-60). If blank, TNCUT set to 50.
5 1-7 YES ANYBOD Wing Y dimension at inboardedge. If negative, program will solve for wing- fuselage intersection.
5 8-14 YES WRAP Nacellepressurecode. field
-1. = Wrap solution for nacelie pressure field is desired. 1. = Glancesolution is performed.
70 CaXd Cad Decimal Variable Description Number Column Required Name
5 15-21 YES D LT2 Interferenceprintout code.
-1. = Summary table printout only. 1. = Details of nacelle/fuselage interference calculations will be printed . -
5 22-28 YES BCUT Numberdivisions of nacellesof used to define nacelle pressures and Whitham F(Y) function. BCUT 6 40. If blank, BCUT set to 40.
6 1-35 YES TXM additionalArray of Mach numbers, Solution will be performed for these Mach numbers after the solution for XM.
Note:There will be a total of TNOM values on thecard. Omit this card if TNOM = 0.
7 1-70 YES TYB2 semi-spanArray ofelement values of row at which wing thickness pressures are calculated.
Note:These values should be wholenumbers beginning with 0. andending with TNON. Up toten values per card. Up tothree cards. Omit these cards if TJBYMX was not specified.
a 1-3 END
71 4.7 Wing Desigr?. Program
TheKing design program principally requires a wing planform (supFliedfro= the basic geometry), a description of theloadings to be used in optimizing the wingshaFe, and sFecification of the designpoint aEd constraints to be applied to thz solution.
Punch alldata, with decimals, to theleft in the card colums (10 field -7 digitformat).
Defaultoptions are provided to help keeFir?put simple. These include:
0 'ILOADS This is tho number of loadings to be used infinding an optimum loadingcombination. If input as a positive Rumher, the specified riumber of1oa.dings will be taken, in order, from the table onpage 73. (A negative sign requires the user to list theloading numbers to be used.)
0 XOCI\iUlVl This is the number of percentchords used in printingthe camber surface output. If blark, standardpercent chords are used.
0 TJBYMX Standardsemi-span stat.ions are provided if TJBYMX = 0.
If program options are used that requirewing thickness pressures, nacelle buoyancyfield, fuselage upwash loading, or asymmetric fuselageloading, it is necessary to havepreviously run the near- field wave drag or lift analysisprograms to load the proper tables. This is done asfollows:
0 Nacelle buoyancy May be calculated by eitherwing loadinganalysisnear-fieldprogram or wave dragprogram.
* Wing 'hicknessobtained from near-field wave drag F ressures program, Fressures
0 Fuselageupwash Obtainedby running lift analysis loading programwith wing sloFes zeroed (WHUP = 1 .O) .
0 Asymmetric0bt.ained from analysislift program fuselageloading with SYMM # 1.0.
The most efficient way to obtain all of the configuration dependent data is to firstrun the near-field wave dragprogram,
72
". .. "" WING DESIGN LOADINGS
Loading Number Definition
~~~~~ 1. Uniform 2. Proportional to x, the distance from the leading edge 3. Proportional to y, the distance from the wing centerline 4. Proportional to y 2 5. Proportional to x 2 6. Proportional to x(c - x), where c is local chord 7. Proportional to x 2 (1.5 c -x) 8. Proportional to 2 (1 + 15 C 9.' Proportional to (1.05c-x)'.~ 10. Elliptical spanwise, proportional to d% 11. Proportional to x, the distance from the leading edge of an arbitrarily defined region 12. A camber-induced loading proportional to the bodybouyancy loading 13. A camber-induced loading proportional to the body upwash loading 14. A camber-induced loading proportional to the nacelltr I~uyerlcy loading 15. The body bouyancy loading 16. The body upwash loading 17. The nacelle buoyancy loading
73 without nacelles, to get the wingthickness pressures. Then run the lift analysisprogram, with nacelles, and with the zero slope option (WWP = 1.0) and asymmetricfuselage option (SYMM #O,) . Thefuselage upwash loading will .be thatobtained with the fuselage at the specified incidence. If the upwash fields corresponding to a series ' of fuselageangles of attack are desired, it will be necessary to rerun the lift analysisprogram to produceeach upwash pressure loading, Notice thatthe fuselage angle of attack in both the lift analysisand wing designprograms can be different-from the incidence in the basic geometry, In the wingdesign program, angle of attack is inputon card 3. Inthe liftanalysis program,angle of attack is input on card 24 as a special case ofpressure limiting option (requires FLIMIFl.0 on card 4, appropriate value of VACFR oncard 23) . Thenumber of elemefits inthe winggrid system is controlledby input TN0N (normally set at 40.). Depending uFon the Mach number andplanform, some programdimension limits may be exceeded with TNQIN = 40. If thisoccurs, the program solves forthe allowable TNBN, prints it, andstops the design case, It is thennecessary for the usertoreduce TNBN and the associated camber line variables TJBYMX and TJBYS.
**CAUTION**
Theloading options must be used wi.th some care, Loadings 12-14 cannotbeused without also using loadings 15-17. Loading 11 cannot be usedwithout specifying a correspondingplanform region (ANOARB>O). If allloadings are requested, the resultant optimum combination of loadings(and camber shape) may be physically unrealisticif no constraintsupperon surface pressure coefficiznt are imposed.
If the fuselage is included in the solution, it will be necessary to also use z constraintsat the side of fuselage station to obtain a directlyusable camber surfaceshape and a validdrag integration.
74 Caxd Card Decimal Variable Description Number Column Required Name 1 1-4 WDEZ
2 1-70 Any desired TITLE information
3 1-7 YES TNON Numberssemispan ofelements in wing gridsystem. 2. TNON 6 50. If blank, TNON set to 40.
8-14 YES TJBYMX Number semispanof stations at which camber surface is calculated. 2. 4 TJBYMX 25.
15-21 YES TIFAF plateFlatcalculation code
-1. = Use datafrom previous case. 0. = Flat plate calculation will be made. 1. = Flat plate data will beinput on card 9.
3 22-28 YES A J1 Fuselageinputcode
0. = Ignore fuselage definition 1. = Includefuselage
3 29-35 YES YS OB Y valuewing-fuselageof intersection
3 36-42 YES ALPB Fuselageangle Gf attack,deg. (relative to attitude in basic geometry).
4 1-7 YE3 APRINT Printedoutputcode.
-2. = Summary outputprinted. -1. = Inputdata (except large tables) and summary output printed. 0. = Inputdata, output summary and camber shapes at design condition, if requested, are prin-bed. 1. = Same as APRINT = O., plus some diagnostic data. 2. = All input,output and diagnostic data printed.
7s Cad Card Decimal Variable Description Number Column Required Name
4 8-14 YES SMOOTH Codedetermine to smoothing procedure applied to camber surface longitudinal slope at each span station.
0. = No smoothingperformed. 1. = Smooth-as-you-go technique used. 3. = Three pointsmoothing technique used.
4 15-21 YES RESTART Code determineto disposition of force and moment coefficients for component and interference loadings.
-1. = Data fromprevious case will be used. 0. = Data will be calculated byprogram for us e in current case and subsequentcases. 1. = Data will becalculated, and also punched on cards. 2. = Data areread from card sets 18 and 19. 3. = Data are read from tape 3 (written by previous case) Note: See page 35 for RESTART discussion.
4 22-28 YES YSNOOT Y valueforparabolic apex tangent to wing leadingedge. (Leave blank if not used.)
5 1-7 YES XM Basic Mach number 5 8- 14 YES CMO Designpitchingvalue of moment coefficient at zero lift,
5 15-21 YES CLDZIN designValue of lift coef:Bicient. If blank or zero, CLDZIN set to 1.0.
76 Card Card Decimal Variable Description Number Column Req uir ed Name
5 22-28 YES TLOADS Number ofloadings to be combined 2. TLOADS 4 17. TLOADS < 0. = Loadingnumbers will be input on card(s) 30. Loadingnumbers will be taken from table on page 73. TLOADS > 0. = Loadings will be in the order tabulated on page 73- E.g., if TLOADS = 3.0, first 3 loadings frompage 7'3 will be used.
5 29-35 YES XOCNUM Number ofchordwise locations at which camber ordinates will be printed, corresponding to options selected on card 7. (XOCNUM) f. 20.
0. = Defaultlocations of O., 5., lo., 20. , 30. , . . . 90. , 100. are used . Omit card 11. + = Values in percent of local chord will beinput (card 11). 5 36-42 YES ANOARB Numbers of points on cards12 and 13 used to define. the arbitrary region of the wing planform for loading number 11. ANOARB 20. If blank,cards 12 and 13 not read,
6 1-7 YES AXCPLIM Number ofchordwise locations(card set 14) used to specify wing upper surface limitingpressures, AXCPLIM '15.
"- Use values from previous case if /AXCPLIM/ same as previouscase. 0. = Card sets 14, 15 and 16 notread. +. = Card set 14, 15 and 16 are read. 6 8-14 YES AYCPLIM Number ofspanwise stations (card set 16) used to specify wingupper surface limitingpressures. Needed. only if AXCPLIM > 0. AYCPLIM 4 15.
6 15-21 YES TXCPT Code torequest use ofwing thickness pressures in pressure limiting e calculations.
0, = Wing thicknesspressures not used, 1. = Wing thicknesspressures used.
77 Solution and ConstraintOptions
Card 7 contains four inputs which control the extent of the solution and the constraints to be applied. Eachof the 4 inputs may take on 4 different values, as follows:
0. No solutionof this typedesired.
1. Calculatepressure distribution, drag, and pitching moment for optimum combinationof loadings,
2. Same as 1, plusalso calculate the wing shaperequired to support the optimum pressure distribution,
3. Same as 2,plus also punch the wing shape on cards. Order is percentchords for ordinates, percent span stations, and then theordinates in percent chord. 10F7.3 format. (May beinput directly into wing analysis program with TIFZC = 1.0).
Card Decimal Variable Card Description Number Column Required Name 7 1-7 YES CONSTR(1) Obtainsolution for minimum dragwith constraint on CL only.
7 8-14 YES CONSTR(2) Obtainsolution with constraints on CL and Cmo (requires Cmo value on card 5).
7 15-21 YES CONSTR(3) Obtainsolution withconstraint on CL and pressure limiting onwing upper surface.
7 22-28 YES CONSTR(4) Obtainsolution withconstraint on CL and C,,, pluspressure limiting on wing upper surface.
7 29-35 YES AN2 Number of 2 constraintlocations (65).
.
78 Cad Card Decimal Variable Description Number Column Required Name
7A 1-35 YES XZC~N X location of z constraint
7B 1-35 YES YZCdN Y location of Z constraint
7c 1-35 YES ZC @N Z values at XZC@N and YZC@N.
8 1-70 YES TJBYS semispanArray of stations at which the camber surface is calculated.
Note: Up toten values per card. There will be a total of TJBYMX wholenumbers which must beginwith 0.0 andend with TNON. If TJBYNX was blank,the following 23 values are used: O., I., 2., 4., 5., 6., 8., lo., 12., 14., 16., ..., 38., 40. Omit cards(s) 8 if TJBYMX = 0.
9 1-7 YES XF Wing aerodynamic center fraction. 9 8-14 YES SCL9 Flat wing lift-curveslope (per degree), based on the reference area for force and moment coefficients,
9 15-21 YES KF Flat wing lift-dependent drag factor.
9 22-28 YES AREA9 Flanform area in program units.
9 29-35 FACTOR Gross wing area/reference area.
Note: Omit thiscard if TIFAF (card 4) 0. The data on card 9 would normally be calculated by a previous run of the same planform at the same Mach number.
10 1-70 YES TLOAD Loadingnumbers pressureinuse for optimization,Integer numbersfrom 1.0 to 17.0, TLOADS (see card 5) in number,and inarbitrary order. Up to 10 valuesper card. Ofit card (s) 10 if TLOADS > 0.
79 Card Card Decimal Variable Description Number Co,lumn Required Name 11 1-70 YES TPCT Arrayof X/C (percentlocalchord)of values will be interpolated at each span station. Omit card(s) 11 if XOCNUM = 0.
12 1-70 YES YARB Array of Y coordinates which definean arbitrary planform region for loading number 11.
Note: Up toten values per card. Up to two cards. There will be a total of ANOARB values. If ANOARB (card 5) is blank,omit card set 12.
13 1-70 YES XARB Arrayof X coordinates which definean arbitrary planform region for loading number 11.
Note: Up toten values per card. Up to two cards. There will be a total of ANOARB values. If ANOARB (card 5) is blank,omit card set 13.
14 1-70 YES XCPLIM chordwiseArrayof locations (percent of localchord) used to define the wing upper surface limiting pressure coefficient. Needed if CONSTR(3) or (4) is f 0 on card 7.
Note: Up toten values per card. There will be a total of AXCPLIM values starting with 0. andending with 100. If AXCPLIM is notpositive, omitcard set 14.
15 1-70 YES YCPLIM spanwiseArrayof locations (percent or semispan)used to define the wing upper surface limiting pressure coefficient,
Note: Up to tenvalues per card. There will be a total of AYCPLIM values starting with 0. andending with 100. If AXCPLIM is notpositive, omit card set 15.
80 Decimal CaXd Card Varriable Description Number Column Required Name 16 1-70 YES CPLIMIT Arraylimitingof pressure coefficients on the wing uppersurface. All coefficients at a given semispan are input in the same order as XCPLIM, Begineach semispan set on a new card and in the same order as YCPLIM.
Note: Up toten values per card. There will be a total of AXCPLIM X AYCPLIM values. If AXCPLIM is not positive, omit card set 16.
17 1-70 YES CPGRAD Array of limitingupper surface pressure gradients (dCp/dx).Input at same X and Y stations as C'PLIMIT, in same order.
*18 1-80 TITLE cardTitle of RESTART data.
"19 1-80 RESTRT Array of force and moment coefficients for componentand interferenceloading, as punchedfrom a previous run, for restarting program execution.
Note: Omit cards 18-19 if RESTART (card 4) is notequal to 2.0.
20 1-3 END
The restast card sets 18-19 are printed on the Output file and identified by thestatement: RESTART DATA PUNCHED, DECK IMAGE FOLLOWS.
81 . . .. ” . . .- . - ....
4.8 Lift AnalysisProgram
Codes oncards 3 and 4 control the inclusion of basic geometry data as requested. Input is in 70 field -7 digit format. Note that the wingcamber surface may be defined in several ways, controlledby input TIFZC oncard 4:
””TIFZC 0. or 1. Input to liftanalysis Frogram on cards
2 Flat wing (Z = 0 everywhere)
3. As defined by wingdesign Frogram {which must have been run previcusly).
4. As definedbasic geometry.in The wing cambersurface input to the lift analysisprogram will automatically be used to update the basicgecmetry definition if TIFZC = 0. or 1 . By definition, a car.ard is required to be locatedforward of the wing,and a horizont,al tail aft of the wing. One each is allowsd, and they may both be input. at the same time. Both are assumed to be mcunted on the fuselage.
The fuselagedefinition (iR basic geometry) cannothave discontinuities in z (meanlineshape) between seqments, when used in the lift analysisprogram.
The downwash shift options (for canard and hing downwash) are controlled by inputsoncard 6. If the shiftcodes are left blank, t.he downwash will be shifted according to the respective side-cf-fuselagestations of canard,wing and tail.
If the pressure lixiting feature(controlled ]cy FLIMIT oncard 4) is used, it requires the wiRg thickness pressures from the near- field wave drag program,which must have been runpreviously at the same Mach number.
All anglesare input to the program in degrees.
82 Card Decimal Variable Card Description Number Column Required Name 1 1-4 ANLZ
2 1-70 Any desired TITLE information.
3 1-7 YES AJ2 Fuselageinput code.
0. = Ignore fuselage definition. 1. = Include fuselage definition.
3 8-14 YES AJ3 Nacelle inputcode,
0. = Ignore nacelle definitions. 1. = Include nacelle definitions.
3 15-21 YES AJ5 Canard inputcode.
0. = Ignorecanard definition. 1. = Includecanard definition.
3 22-28 YES Horizontal tail inputcode.
0. = Ignore horizontal tail definition. 1. = Include horizontal tail definition
4 1-7 YES TJBYMX Number ofspanwise stations defining camber surface. TJBYMX 25.
4 8-14 YES TNOPCT Number ofpercent chords defining each spanwisestation. TNOPCT 6 20.
4 15-21 YES TIFZC Code for camber surfaceordinate.
0. = Z is input. 1. = Z/C (percent) is input. 2. = Flat wing option (Z = 0). 3. = Camber surface is defined in common block /CAMBER/. 4. = Use definition contained in basic geometry.
Note: If TIFZC is 2., 3., or 4., inputs TJBYMX, TNOPCT,TPCT, TYB2 and WZORD are not required.
83 Decimal Variable Card Card Description Number Column Required Name 4 22-28 YES TNOM Number of Mach numbers in addition to basic Mach number XM. TNOM C 5.
4 29-35 YES F'NON Number of semi-span rows in wing grid system. FNON 6 40. If leftblank, will be set to 40.
4 36-42 YES FLIMIT Limiting pressure feature code. 0. = feature not desired. FLIMIT = number of configuration angles of attack for solution using pressure limiting.
1-7 YES TNFLAP Number of trailing edge flaps on right hand wing, TNFLAF C 5. 8-14 YES TNTWST Number of values (Y in percent, and angle)to define wing twist. Relative toinput wing shape. TNTWST 6 40.
5 15-21 YES TNALP Number ofcanard angles of attack (5 5). Not required if AJ5 = 0. 5 22-28 YES WRAF Code for nacelle pressure field solution
-1. = wrap 1. = glance
5 29-35 YES OXML Mach number input code for nacelle pressure field calculations.
0. = Free stream Mach number used. 1. = Mach number input on card 19.
5 36-42 YES DLT2 Nacelle pressure field calculation printoutcode.
-1. = summary only 1. = detailed printout 5 43-49 YES BCUT Number of cuts used to define pressure signaturefrom nacelles. If blank, will be set to 40. BCUT f. 40.
84 Decimal Variable CaXd Caxd Description Number Column Required Name 6 1-7 YES ANYBOD Wing/fuselageintersection Y value. If negative,solve for intersection. If ANYBOD = -10.0, intersection will be input on Card Sets 14-16. 6 8-14 YES THAL?? Number ofhorizontal tail angles of attack. THALP 6 10. Not required if AJ7 = 0.
6 15-21. YES SYMM Asymmetricbody volume termcalculation code.
0. = Do notcalculate 1. = Calculate 2. = Calculate using area distribution input on Card Sets 17 and 18. 6 22-28 YES SMOG0 Smoothingcode.
0. = Use 9 termsmoothing 1. = Use smoothing-as-computedpressure calculation.
6 29-35 YES WHUP Wing slope control cde.
0. = Wing slopescalculated from input camber surface. 1. = Wing slopes = 0. (usedfor fuselage . upwash field) ,
6 '36-42 YES FWSH Wing downwash shift, code
0. = Shift according to DYWH 1. = No shift
6 43-49 YES DYWH Shift of wing downwash at tail
0. = Use incrementbetween side-of-fuselage stations ofwing and tail DYWH = Y shift increment (+ = outboard)
85 Cad Card Decimal Variable Description Number Column Required Name
6 50-56 YES ' FCSH Canard downwash shift code
0. = Shift according to DYCW and DYCH 1. = No shift
6 57-63 YES DYCW Shift ofcanard downwash at wing
0. = Use incrementbetween side-of-fuselage stations of canardand wing DYCW = Y shiftincrement (+ = outboard)
6 64-70 YES DYCH Shift ofcanard downwash at horizontal tail
0. = Use incrementbetween inboard Y stations ofcanard and tail DYCH = Y shift increment (+ = outboard)
7 1-7 YES XM Basic Mach number for case. 7 8-14 YES TZSKAL Scalefactor for input Z ordinates. If blank, no scalingperformed.
7 15-21 YES CLIN( 1) Number of lift coefficients input for first Mach number (XM) at which the combined flat plate andcamber pressure coefficients will becomputed. . (CLIN(~) 5. )
22-28 YESB CLIN(2) Same as CLIN(1) for second Mach number ( TMACH( 1) ) .
29-35 YES CLIN(3) Same as CLIN(1) but for third Mach number (TMACH( 2) ) .
36-42 YES CLIN(4) Same as CLIN(1) but for fourth Mach number (TMACH( 3) ) .
43-49 YES CLIN(5) Same as CLIN(1) but for fifth Mach number ( TMACH(4) ) .
50-56 YES CLIN( 6) Same as CLIN(1) but for sixth Mach number (TMACH( 5) ) .
1-35 YES TMACH Array of additional Mach numbers for thiscase. TNOM values. Omit thiscard if TNOM = 0. Card Card Decimal Variable Description Number Column Required Name
WingCamber Surface Definition
Omit card sets 9, 10 and 11 if TIFZC = 2. , 3. , or 4.
9 1-70 YES TPCT chordArraypercentages of at which Z (or Z/C) ordinatesare input and pressure coefficients are evaluated and output.
Note: Up toten values per card. Up to two cards, There will be a total of TNOPCT values from 0. through 100.
10 1-70 YES ArrayTI32 of semi-spanpercentages at which Z (or Z/C) ordinatesare input.
Note: Up toten values per card. There will be a total of TJBYMX valuesfrom 0. through 100.
11 1-70 YES w zom Arrayof Z (or Z/C) ordinates of the right handwing camber definition. All ordinates at a givensemi-span are input inthe same order as TPCT. Begineach semi-spanpercent on a new card and in the same order as TYB2.
Note : . Up to ten values per card, There will be a total of TPCT x TYB2 values.
Wing Twist Definition
Omit cards 12 and 13 if TNTWST = 0.
12 1-70 YES YTWIST Arraysemi-span of percentages at which wing twist angles are input.
Note: Up toten values per card. Up tofour cards. TN'IWST values.
13 1-70 YES ATWIST Arrayof twist angles,degrees.,in corresponding to YTWIST. A positive angle means an increase in local angle ofattack. Linear interpolation is used for points between input points.
Note: Up toten values per card. Up tofour cards. TNTWST values.
87 Card Decimal Variable Card Description Number Colum Required Name
Wing-Fuselage Intersection
Omit cards 14-16 if ANYBOD f -10.
14 1-70 YES wx X values
15 1-70 YES WY Y values * 16 1-70 YES WZ Z values
Input X array defining wing-fuselage intersection, then Y and Z. Starteach array on a new card.Values are input at thepercent chordsof the camber surfacedefinition (Card 9), or basic geometry definition (if WZORD not input).
Asymmetric FuselageArea Input
Omit cards 17 and 18 if SYMM f 2.0
17 1-70 YES AOVR above-wing area 18 1-70 YES AUND under-wing area
Input area distribution abovewing, then below, Start each array on a new card,Values are input at thepercent chords of the camber surfacedefinition (Card 9), or basicgeometry definition (if WZORD notinput).
Alternate Mach Nos,For Nacelle Pressure Field Calculations
Omit card 19 if OXML = 0.
19 1-42 YES TMLOC Arrayof local Mach numbers fornacelle pressurefield calculations. First value correspondsto XM, successivevalues correspond to TMACH (if included).
Note: Up tosix values on thecard. There will be a total of TNOM + 1. values. Card Card Decimal Variable Description Number Column Required Name Wing Flap Definition
Omit cards 20 if TNFLAF = 0.
20 1-7 YES x1 Inboard Xleadingflap value ofedge.
20 8-14 YES Y1 Inboard Y leadingflapvalueof edge.
20 15-21 YES x0 Outboard X flapvalueleadingof edge.
20 22-28 YES YO Outboard Y value of flapleading edge.
20 29-35 YES DEFLAP Flapdeflection degrees.in A positive angle means the flap trailing edge is deflected downward.
Note:There will be a total of TNFLAP cards, one foreach flap.
21 1-35 YES TCA canardArrayangles of of attack. A positive angle means the leading edge is rotated upward.
Note:There will be a total of TNALP values on thecard. Omit this card if TNALP = 0. or AJ5 = 0.
22 1-63 YES THA horizontalArray of tail anglesof attack. A positiv.?angle means the leadingedge is rotated upward.
Note:There will be a total of THALP values on thecard. Omit this card if THALP = 0. or AJ7 = 0.
23 1-7 YES VACFR Fractionof vacuum pressurecoefficient for pressure limiting.
24 1-35 YES TLALF Arrayof &'s limitingfor pressure coefficient.
Note:There will be a total of FLIMIT values on thecard. Omit cards23 and 24 if FLIMIT = 0.
89 Decimal Variable CaXd Card Description Number Column Required Name
25 1-35 YES CLINP Arraysof lift coefficientsthefor input Mach numbers (XM and TMACH) at which the combined flat plate and camber pressure coefficients are computed . CL'S for each Mach number begin on a new card.
Note: Up tofive values per card. Up tosix cards, The number of values on eachcard will correspond with CLIN(1) on card 7. If CLIN(1) = 0, omitthe Ith card.
For a new case,input cards 2 through 25 at this place in the data deck,
26 1-3 END
90
- ... ". . ..- - .- ...... 5.0 TYPICAL CASE AND FROGIZAM OUTPLJT
A tyFicaldesign and aflalysis casejandassociated program output areFresented in this section.Given a configurationconsisting of wing,fuselage, nacelles and horizontal tail, thefollowing are obtained: Wing designat Mach number = 2.7 for CL = .10 and Cmo - .010, inpresence of fuselageand nacelles with pressure constraints. Analysis of configurationdrag-due-to-lift for a series of horizontal tail settings. Skin frictiondrag Far-fieldand near-field wave drag aralyses
* Drawing of configuration.
The inputcard listing for this case is shown cnpage 99.
The Frcgramoutput has been editedto reduce page count while illustratingoutFut format.
The outputbegins with a listing of the basic geometry,separated intc components(wing, fuselage, etc) (. An uxamberedwing was' specifiedin the basicgeometry, since the carrber surface will be deficed by the wingdesign program.
Configuration-DependentLoadings
Since tne wingdesign case is to be performedwith pressure lirriting,and inthe presence of fuselageand mcelles, the corr€s&ondingpressure arrays must becomputed. The near-field wavedrag Frouran; is runfirst, tc generate the wing thickness Fressuredata (page 109). Only the winggeometry is requiredfor chiscalculation; outFut for the complete configurationfrom the near-fieldprogram is illustrated later (page 234).
The liftanalysis program is executednext, to calculatethe nacellepessure field and the fuselageupwash pressure field. To obtain an approximateorientation between the fuselageand wing for the upwash fieldcalculations, a previouslydefined camber surface was inputusing the TICFBC = 1 .0 qption. The ANLB interfaceprogram inserts this definition into the basicgeometry andFrints it (page 110). The lift analysisprogram then computes the wingupwash field(page 1151, the nacelleFressure field (page 117) theasyrcmetric fuselage buoyancy pressure field(page 116), and the loadingon the wing due to the fuselage upwash field(page 121). Thewing upwash loading is that for the basic wingangle of
91 attackwith all wing slopes zeroed, i.e., as ccmputedwith input WHUP = 1.0.
Wing DesignSolution
Much diagnosticoutput is availablefrom the wingdesign module. However, printccntrols .3re used inthe progran (input A-PRINT) to provideoutput flexibility, In thetypical case shown, the print contrcl was Set at +2.0, to illustrateoutput format, and then editEd. The design case shown uses all seventeenloadings; first to generate a FESTART deck,then to obtain a Kingdesign for a sFecificdesign point. using the RESTART opticr.. ThE fuselage is includedin the solution.
TheKing designprogram first prints the input data andchecks the designand ccnstraint oFtioris (the card 7 inputs) for consistency. The serri-sFanstations, inprogram units, at which the camber surface will be calculated is nextprirtted, followed by a listing of the ccmponentloadings to be usedana the chordwiss locations at which the camber surface will be interFolatzd.Tables of the ccnf iguration depndent loadings are also output.
Five Z ordinat€consyraint locations are specified in theinput. These arechecked to see if the Y and X values of the constraints are cn a computsd Y stationand OR theplanform, respectively. (The Y values were shifted slightly and a noteprinted).
ThE Frogram rext cornputesand prints the flat wing solution(page 133). Thisincludes lift and drag coefficients, the lengthwise center of pressureposition (as a fracticn of overall wing lcng+h), and the drag-due-to-lift factor. Sincethe fuselage is used in thesolution, the aerodynamic center lccation of the .wing- fuselagocombination (computed in the lift analysis program at the tine ot thefuselage upwash calculation) is substituted for the aerodynanic center or thewing planform only age 133).
The Frcgraxrthen calculates the carry-over lift distribution of all tne camber-typeloadings (page 134), and the associated force coefficients.
TheFrogram next cycles through all the componentlcadings. For each, a table givingspanwise distributions of lift,drag, and pitching moment coefficients is printed. This is followed by the integrated values of lift coefficient, drag coefficiect, center of pressureposition, drag-due-to-lift factor, the ratio of input referencearea to grossplanform area (Sref/Spros), thepitching morrert slope with design CL, and the Cno associated with the compcnent cL. This is followed by the interference drag of the compcnentloading on the nacelle area distribution(if nacelles were input),and a tabulationtheinterferenceof drag coefficientsassociated with all othercomFcnent loadings. The
92 camber surfacefor the selected loading is ther!pr+nted, together with. the liftingpressure distribution and upperand lower wing surfacepressure distributions. The camber surfaceinboard of the side-of-fuselagestation is set to zero,since it is replaced by thefuselage shape. The individual camber surface data are. not shown, buthave the same formatas the finalsclution (page 182). TheFrogram next summarizes the force andinterference drag coefficientsof all the componentloadings (page 156). The order of the data are:
1) Lift coefficients for all loadingsand their respective C,, constributions (for theexposed wing part). 2) Interference drag coefficientsfor all wingloadings (page 156). 3) Drag coefficients of wing-on-nacelle(page 158). 4) Fuselageinput to wingdesign point (lift, drag, and pitchingmoment), transferred from lift analysis program (page 159) 5) Lift, drag, and Cmo contributions of the carry-overlift distributions(page 159).
All cf these data, plus the configurationdepen'dent pressure distributions,are then punched into a RESTART deck, and the deck imageprinted (page 159). Only a portion of the RESTART listing is shcwn since it is quitelong. (The siz.e cf th? RESTART deck is a furictionof the number of loadings,whether fuselage is used, nurber cf cocstraints, etc) . With all componentloading data defined, the Frograrcthen solvos for the wingdesigns requested on card 7. (If thedesign case is run from a RESTAkT deck, the Frogramsolution commences at this pcint.) The solutioncor.ditions are summarized (i.e. , CL, Cmo, 2 ccnsTraints, etc), followed by the optimizedvalues of CmO, KE (drag-due-to-lift factcr, cD/cL2), and the associatedloading conkination kactors AiC Li The respective contributions I of zxFosedwing, fuselage, carry-over lift, and nacelles to the configuration is thenprintEd.
The solutionpressure distribution is nextprinted ar,d scanned for pressureconstraint violations, If any occur (either in level or gradient),they are notedinthe right handmargin. At the ccnclusion of thepressure distribution print-cut, the locations' andmagnituds of the largestsolutior, Fressure level and gradient are ncted (page 171). If violations of inFutpressure limits OCCUT, tile solutior,repeats with a constraintadded at the location of worst violation. For the test case shown, a wingdesign bas obtainedin a suksequentrur. 'using RESTART. Since all of ths basic solution data were preservedin the RZSTAET deck, it was not necessary to
93 recalculate the configuration-dependentda%a. The RESTART deck is valid for any case having:
1) The same. or fewer loadings (order can be changed) 2) Same fuselagegeometry, angle of attack,and side of fuselage station 3) The same or fewer Z constraintlocations (in same order). The value of Z at theselocations can be changed,however. 4) Any CL, CmO, or pressureconstraint. In the particular test case shown, the value of Cmo was changed frcmthe case whichgenerated the RESTART data,and only four (of thefive available) 2 ordinatelocations were used.Scfution pressuredistributions were requested €or all four canber surface options (CL only, CL pluspressure constraint, CL plus Cmo, CL plus C,, + pressureconstraint). The resulting camber surfacefor
CL + Cmo + CP was requested to be output ar,d also punchedinto cards.
The solution commences forthe CL (and Z) case, It thencontinues byaFFlying the Fressure and Cmo constraints.
Incrder to illustrate programoutput, the solutioninvolving all constraintshas keen editedand is shown (beginningon page 160). The initialsolution has a numbEr of pressureviolations, the worstof which is identified(page 171), and a constraintapplied there. The solution , thenrecycles, and identifies a second ccnstraint to ke applied,
Subsequentsoluticn cycles build up to fourgradient constraints, one of which .is foundredurldant (i.e., made u.nnecessaryby a later constraint), as showl, on page 177. Thatconstraint is removed, together with thelast constraint applied (sixce it involved a redundantconstraint), as shown pageon 178. The solution continuesuntil the gradientconstraint is everywheresatisfied, and thenchecks Fressure level. In this case, level was already satisfied, so the finalsolution SUraiXy is Frinted,including a sunmary of the twelvelargest pressure gradients on the wing upFer surfacefor the final solution (page 181).
After thefinal solutior: is obtained, the programcalculates any requestedcanber surfaces, The sFanwisedrag sumary, force coefficient summary, Z values are printed as was done for eachof the ccnFonentloadings earlier, The canbcrsurface is then interpolatedat the requestedpercent chord values, printed and punckdinto cards (page 194).
94 Wing Camber SurfaceUpdate
In the illustrative case, the firial camber surface design was used to uFdate the basicgecmetry by means of theexecutive card WGUP. The updateddefinition is printedon page 196.
Lift Analysis
Thecamber surface definition punched by the wing designprogram was inputinto the llft analysis Frogram. The wing camberline definitionat ,075 semi-span(side-of-fuselage station in wing design Frograrr.)was substituted for the zerospunched by the wing designFrogram theinfuselage region, ir, order to allow calculation of the wing-fuselageintersection.
The fcrcecoefficient summary, tail-off, is shown on page 215. The Frogram first prints a table of lift,drag, and pitching morrent coefficients for the wing at the inputiccidence, and also perdegree angle of attack (FP at 1 degree). The incrementsdue to the nacellesare also printed. This table is thenrepeated with the fuselagecontributiori added. Tne drag terms are thon combinedinto two equations (nacelles onand off), and drag and pitching moment coefficientstabulated for a series of lift ccefficients.
The configuration streamwise lift distribution is next summed and printedand further broker! into separatesummations for wing- fuselage-canard,nacelles, and horizontal tail, These summations are cumulativeand are divided by the total lift of tne configuration.
The forcecoefficient and streamwise liftdistribution data are repeated for each tail angle of attack, togetherwith the ccntributionsdue to the horizontal tail.
95 The kinglifting Fressure distribution at a Specified CL is next printed (in this case, at a CL = .lo). These data(page 222) are fcr the Fressuresgenerated by thelifting surface, butdo Got include pressuresdue to the nacelles or asymmetricfuselage buoyancy field.
The spanwise lift distribution is printed last (page 224). This tabulation is for thawing-canard-nacelles combination only (excluding fuselage or horizontal tail) . .If the lirritingFressure oFtion of the Lift asalysisprogram is requested, the output is the same exceptfor tho alterations:
1. The data at tne configuration basic angle of attack become data at a specifiedangle of attack.
2. Notes areprinted to call attention to the pressurelimiting cption.
Addition of a canard to the configurationFrcduces an additional set cf force coefficient sununary data, i.e., data is printed both with and without the direct canardcontributicn.
Skin-Friction
The skin friction program printsinput, thec a table of wetted areas,drag/dynamic pressure (D/S) , anddrag ccefficient, fcr each inptflight condition (page 227). Far-Field Wave Drag
The far-field wave drag program prints ar! enrichedarea distributionfor the fuselage(page 2281, then the area distrlkution for differentconfiguration component buildups at a series of theta (cuttingplane inclination) values. The program next identifies andprints the arearestraint Foints corresponding to the case restraintcondition, followed by configuration data for the inptconfiguration and one optirrized subject to the restraintpcints. An optimizedfuselage area distribution corresponding tc the restraint case is thencalculated and printed, followed by a drag summary for %he configurationas-input ana with the optimizedfuselage (page 233).
Near-Field Wave Drag
The near-field wave drag module, for wing-fuselage-nacelles, was executed next. The programinput is first printed, followed by the wingfuselage intersection. (The E values of this intersectionare relative to the fuselagecenterline, .rather than
96 thecverall coordinate system.) Thickness pressure distributions for the empennagesurfaces are thenprinted (Fage 235). The nacelle terms are nextprinted. First the nacellepressure fieldacting on the wing is output(edit.ed out in this case since
it is ' ths same aspreviously illustrated in the lift analysis programoutput) . The interferenceFressure signatures associated withthe nacelles and fuselage acting on cneanother are next calculatedprinted,and including the "imaget1 signatures associatedwith reflections off the wingsurface. Thebuoyancy field of thefuselage acting on the wing is then suIcmarized,followed by the wing definition and isolated thic3mess pressuresolutions. The isolatedfuselage pressure distribution and the wing-on- fuselagesignature is next tabulated(page 244), .togetherwith a runningsummation of %he dragassociated with thesepressures. Each of these sums is divided by thetotal corresponding drag value.
The finaldrag summary (page 249) consists of kingsection data, takulatedfuselage and nacelle drag coefficiests, emperaage drag, total dragand wetted areas.
Tke kingsection data, at the solutiorrspanwise stations, consist of the isolated wing sectiondrag coefficient (CDW/C= drag of the element row divided by chord) , interferencedrag of fuselageon wingsection (CDBaWIC), interferencedrag of nacellesacting on the section (CDNQIWK), the sum of thosesection coefficients (SUM CWC), and the fraction of thetotal winq wave drag for the section.
Drag of the wing-fuselagecombination is nextprinted, including theisclated wing (CDW) , isolated fuselage (CDE) , fuselage-cn-wing interference (CCBIW) , wing-on-fuselageinterference (CDWIB) , and the total of those (CD WING-BODY).
A A tatle of nacelledrag terms is thenprinted, giving the isolated wave dragand the interference terms forthe nacelles at each inpt origin.
The total wave dragfor the configuration is pinted as T0TAL CD-
PlotProgram
The Elct Frogram prints the programinput andviewdata, A typicaldrawing frcm the program is presented on page 12.
97
4 40.0 60.4 70.0 5-1 5-2 6-1 6-2 b-3 64 6-5 6-6 6-7 6-8 1.2 B-€-1 8-1-2 1.2 8-2-1 1-2-3 1-213 8-3-1 8-3-2 1.235 1-4-1 1-4-2 1.219 8-5-? 8-5-2 1.343 8-6-1 8-6-2 1-32 8-7-1 8-7-2 . 1.205 8-6-1 e-e-2 100.0 133.33 L50cO 10-1 266.67 295.0 10-2 1.26 11-1 -13.2 11-2 11908 12-1 3T.O 12-2 16-1 35.04 17 3 -42 18 16-2 35.06 17 3-62 18 10. 19-1 20-1 21-1 -.9 19-2 20-2 21-2 -14. 22 23 24 1. 3'
99 0.0 2.7 0.0 4 4.969 -1. 5 END :, 8 ANLZ 1 CONFIGIJRATION-DEPFNDEMT LOADINGGFNERATION 2 1. 1. 3 It. 12. 1. 4 1. -1. 5 -1. 1. 1. 6 7.7 7 0.000 5.060 10.000 3o.ooo 70.000 40.00050.000 m.ooo 70.000 80.000 9-1 90.000100.000 9-2 0.000 5.noo to.ooo 20.000 3n.noo 4n.000 5o.ono 6o.noo 70.000 80.000 10-1
90~000l00~0@0~ 10-2 0.000 "093-e453 -1.478 -2.667 -3.896 -5.094 -6.213 -7.218 -8.082 11-2-1 -8.781 11-2-2 0.000 11-2-1 -8.781 11-2-2 0.000 11-3-1 -7.147 11-3-2 0.000 11-4-1 -5.676 11-4-2 0.000 11-5-1 11-5-2 11-6-1 11-6-2 11-7-1 11-7-2 11-8-1 11-8-2 11-9-1 11-9-2 11-10- -2.211 -2.lr56 11-10- 0.000 "136 "655 -1.241-1..760 -2.211 -2.557 -2.900 -3.266 -3.622 11-11- -3.970-4.308 11-11- 0.006 "'339"653 -1.208 -1.h66 -2.015 -2.238 -2.412 -2.542 -2.627 11-12- -2.646-2.597 11-12- END 26 YDEZ 969-50017 LOAD CHECK CASE 22 SPAN STA. *ITH FUSELAGE AND 2 TERMS 2. 4 0. 22. 40. 0. 1. 4.9688 0. 3 2. 1. 1. 4 2.7 e006 -1 e006 2.7 -17. 0. 2. 5 7. 2. 1. 6 1. 0. 0. 5. 7 130.85 189. 243.39189.189. 1A 4.96884.968@ 4.9688 6.625 8.2873 78 -4.07 -10.16-14.11 -8.32 -7.0 7c 0. 1. 2. 3. 4. 5. 6. 7. 8. 10. 8-1 12. 16. 19. 16. 22. 28. 25. 30. . 32. 36. 8-2 98. 40. 8-3 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 10-1 11. 12.16. 15. 17. 13. 14. 10-2
100 r
.0. . 66.25 12 207. 269.8207. 13 0. 100.. 14 0. 100. 15 "137 .-.137 16-1 -.IT7 -0197 16-2 rm2s .a025 17-1 .m2s *oom 17-2 FNO 20 UMZ 1 969-5OOA CHECK CASE RCSTARTCO 2 40. 22. -1. 3 2. 1. 2. .4 2.7 001 .I 5 7. 21 1. 6 €0 1. 1. 7 130.85 189.249.39 7A 4.9688 4.9688 4.9688 78 -4.07 -10.16 -19.11 7c ' 0. 1. 2. 7. a. IO* ' el 12. 14. 16. 30. 32. 340 8-2 40. 8-3 1. 2. 3. 8. 9. 10. 10-1 11. 160 17. 10-2 0. 66.2'3 12 207. . 269.8 13 0.0 100.0 14 0.0 100.0 15 -0.137 -0.137 16-1 -0.137 -0.197 16-2 .0025 .0025 17-1 00025 .Ob75 17-2 969-500 17 LOAD CHECK CASE 22 SPAN ST&. WITH FUXLAGE AND Z TERMS RESTART
+4C RESTART DECI 1NWT 64 (NOT LISTED) e*+ END 20 Wc.I)P 4NLZ 1 969-500 17 LOAD CHECK CASE 22 SPAN STA- WITH FUSELAGE AND Z TERMS 2 1. 1. 1. 3 720 12. 1. 4 1. -1. 5 -1. 2. 1. 6 P.7 1. 7 0.000 5.000 10.000 20.000 30.000 OOrbOO 50.000 60mOOO 70.000 80.000 90.000100.000 0.000 2.500 5.00@ 7,SOO 10.000 12eSr)O 15.000 17.500 20.000 25-000 +O.OOO r5.000 hO.000 47.TOO 55.000 FtoWO 70.000 T5.C)OO @O.000 ~s.001)100.006 90.006 0.000 -166 -182 -0624 -2.108 -3.584 -4.783 -5ob68 -6.254 -6062s -7.137 -8.500 0.000 .I66 ,182 "624 -2.108 -4.564 -6.783 -5066O -6.255 -6.625 -70137 -8.500
101 0.000 ,166 .182 "624 -2.108-3.584 -4.783 -5.668 4.255 -6.625 -7.137 -8.500 0.000 e166 -182 -0624 -2.108-3.584 -4.783 -5.668 -6.255 -6.625 -7.132 -8.500 0.000 -004 -e102 -e799 -1.8U7 -3.009 -4.001 -4.808 -5.397 -5.177 -66'161 -6.858 0.000 -a005 "098 -a707 -1.644 -2.632 -3.523 -4.251 -4.779 -5.085 -5.341 -5.742 0.000 "039 "185, -e879 -1.876 -2.936 -3.920 -4.754 -5.393 -5.807 -6.132 -6.500 0.000 -0032 -e175 -.E38 -1.799 -2.833 -3.806 -4.640 -5.284 -5.702 -5.974 -6.702 0.000 -.017 -e134 -e749 -1.657 -2.649 -3.592 -4.407 -5.036,-5.434 -5.688 -5.938 o.oo?) .020 "071 -e625 -1 -486 -2.447 -3.386 -4.212 -4.864 -5.280 -5.520 -5.770 0.000 .058 -e075 "562 -1.402 -6.028 -6.391 0.000 .OB5 .040 -0432 -1.214 -5.977 -6.609 0.000 a137 .176 -.10s-.?I4 -4.940 -5.431 0.000 -193 e316 -172 "315 -4.530 -5.190 0.000 0328 -532 e596 ,336 -3.157 -3.842 0.000 .403 1.087 1.156
22 25 1 2 ,3 4 5-1 5-2 10 1 2 3 4
i 02 . . . . ._.
EM 6 NFYD 1 NEAR-FtELD YAW DRIG 2 1. 1. 1. 1. 3 0. 2.7 4 4.97 -I. -1. 5 EN0 8 CLOT 1 969-500 CONFlGURATION 1. 1. 1. 1. 2 1. 3 xz 4-1 XY 4-2 YZ 4-3 END 5
I 03 L 0 P
.. . . I .. . ." x0 77.3280 . x0
" . . . .. yo 4.9688 Yn 20 = 0.0000 20 . . .. . - .. , ..C!iORD lb6.070.0 CHORD . . . .. ", .- .- - . .. - ." ... - .- - - ". . -. . CAMBER HALF-T-HICKNES.S CAfl8ER HALF-IHlW(WS . . .-.SAMBER HALF-J HICKNESS .( z I UPPER LOWER (21 UPPER Lnuu . IZJ ~ .UPP€E LOWER _" ." - 0.0. .o. 0000 0.0000 0.0000 0.0000 O.OQOO 0.000.0 . . .. . 4." D LMQO - 0.0000 2s.) 0.0000 e 5700 5700 0.0000 -5700 e.5700. . . . DLOOM .5.500 .5500 5.0 0.0000 e7140 .7140 0.0000 .7140 .7140 . .O.D.QQO -7120 e7120 . lor0 0.0000 ,8720 8 720 0.0000 .8720 ..072Q-. . o doao .a 720 -8720 20.0 0.0000 1.05001.0500 0.0000 1.0500 1.05.QO .O AQO0 1 a540 1.0540. 30.0 .o..oooo 1.14501.1450 0.0000 1.1450 i.1450 .O QOQO 1 1560 1.1560 40.0 0.0000 1.20001.2000 0.0000 1.2000 1.2000 . . OaOOQO 1.2130 1.2130 50.0 0.0000 1.23001.2300 0.0000 1.2300 1.23M. .. . D L QQQO 1.2150 1.2350 -60.0 P.0000 1.24901.2490 o..oooo 1 2490 llL49.0 .. - O..QOOD . . I.2370 . 1.2370 70.0 0.000.0 1.17001.1700 O.OOQ0 1.1700 1.1LPo . . D LQMO 1.1220 1.127.0 80.0 o*0000 ,9370 ,9370 0.0000 9370 93 20 D 90QO e8830 . a.8830 90.0 0.0000 e5460,5460 0 .oooo ,5460 e5460 0 .o.o.o.o e5.07.0 ,5070 100.0 D.0000 0.0000 0.0000 0.0000 .O I OOQO 0.~000 . . naaao .O ..O OQO o.aooo .
x0 116.9600 x0 F 148.9800. x0 =. .225.8100 YO = 16.3330 YO 8 31.2500 .Y 0 = -47.54.40 20 = 0.0000 20 ? 0.00.00 ZD . 3 -0.QQOO CHORD = 125.3500 CHORD * 77.2950 CH [w) > 32~6810
PERCENT CAM8ER HALF-TH IICKNESS CAH8ER HALF-THICKNESS CAMER HALF-THICKNESS CH0.R.D (2) UPPER LOWER (21 LOWERU.PPER 12) UPPER LOWER 0.0 0.0000 0.0000 0.0000 0.0000 0,0000 0.0000 0 ..oooo 0.0000 0.0000 2.5 0.0000 -5500 -5500 . 0.0000 5700 5700 0.00.00 5BQO e5800 5.0 0.0000 ,7150 .7150 0.0000 72 70 7270 0.0000 ,7290 a7290 10.0 0.0000 -8760 -8760 0.0000 e9020e9020 0.0000 ,9110 e9110 20.0 0.0000 1 1260 1.1260 0.0000 1.09801.0980 0.0000 1.1340 1.1340 30.0 0.0000 1 1740 1.1740 0.0000 1.22001.2200 . . .o.oooo 1.26.80 1.2.680 40.0 0.0000 1.2350 1.2350 0.0000 1.28901.2890 0.0000 1.3430 1.3430 50.0 0.0000 1.2500 1.2500 0.0000 1.31501.3150 0.0000 1.3750 1.3750 60.0 0.0000 1.2290 1.2290 0.0000 1.26201.2620 .O.O00.0 1.3200 1.3200 70.0 0.0000 1.0870 1 0870 0.0000 1.1050 1.10.50 0.0000 1.1550 1.1550 80.0 0.0000 -8400 8400 0.0000 -8420 ,8420 0.0000 .e800 .eeoo 90.0 0.0000 ,4740 .4740 0.0000 4730 . -1373.0 .D tQO.0.0 ..k95B , __ . 495.0 1.00,o 0.0000 0.0000 0.0000 0.0000 0.0000 0.0000 0 0000 0~0.000. 0.0000 +*** ****
= 2251II100.. 41~54SO . = 0,0000 . 32.6810 HALF-THICKMESS UP PER 10UER.UPPER ozoooo 0.0000 el340 -1340 e2610 e2410 94950 ,4950 .e800 ,8800 1.1550 1.1350 1.3200 1.3200 103150 1.3750 1.3200 1.3200 1.1550 .. .e800 .a800
._ ." 9OYO. " . . 0,0000 ,495014950 .100,0 0.0000 0.0000 0.OOOP
...... - - ..... -...... "" ...... *++* +++* *+** FUSELAGE.. e+.** ... W~L..-..-..~,W-._ -.**.*( ..... *.*e.+...... - - .....
... -. - ..- ..-. ." ... "...... -. .. .- ...... -...... -...... - ...... " ...... - .. -......
X 2 .... - ." ..... - ...... - .. . _ .... -. .." . FENTERIXNE .- CENTERLINE RADIUS AREA __ PERUETER - . . 0I0000 10.0000 0.0000 .O.OOOO ...... o. 0000...... - . . 16.6700 8.5500 2.7350 23.50OD ...... 17.1846 .... . 33.3300 7.1000 4.2782 . 57.5000 ..... - 2b8806 .... - .. .50.0000 5.6400 5.3226 89.000Q .. .. .33.4426 .. ... - .66.6700 4 1700 6.1026 117~0000...... 38.3429...... 83.3300 2 7300 6.3330 12b.0000 ...... 89.7911 .... 100.0000 1.2800 6.175238.8001 119.8000 . . 116.6700 -e 1400 5.8632 108.0000 ...... 36.8398 ...... 133.3300 -1 6000 5.7812 105.000Q- _ ..... 36.3245...... _..... 1.50.0000 -3 0400 5.8360 ...107.0000. lbr6688..- ...... 166..6600 . -4 e 5000 ...... 36.6688 - . " 5.8360 107.QOOO -. 1.83.3.300. -5.9000 5.8087 - 1Oh.OQQO .--l6.h9U 1" 200.0000 -7.4000 5.6980 102.0000. - - . ". ". -..35.8QlB - "- . - ..216r6700 -8 8500 5.4700 ..... 94.0000 34dbPL- - . . 213933QO -10.2500 . 5.0146 . "" 79.0m "" "3&0.28." ". , . . 250.0000 -11.7000 4.3336 ...... !i9aOOQO IZLzLpQ"--..-.-- .. - _. -.2.66*6700 -13.2000 .3 24 10 .... 33.OQM ...... Lam9 __... "" __. - ... 2839 3000 -14.6000 1.5918. " -. "_ -A"" -.-- .. .2.?5.0000 -15.7000 " 0.0000 ...... P.DQD~ -."~o-...... "
U ... ". X RADIUS X RADIUS "" . ." ....0.0000 ,.. 2..8650 0.0000 2.8690 2.0080 . 2.9830 2.0080 2.9830 15.4700 3.6330 15 4700 3.6330 " ..... 21.5250. 3.7700 21.5250 3.7700 . .28.0170. . 3.6.540 28.0170 3.6540 32.06.70 3.4200 32 0670 3.4209 " ... 3.5.5.0400. 3.4200 35.0490 3.. 4 2.00 "
**** FIN- ' , *+*+* **** **** *+** ++** --~K4uo-*,i* ' ****
X-1 . 22.5.80.00 XL = 270..0000 261.0000 YL s.. 47-~1500 YL = .0..0000 2..0000 , . ZL . .Q..OOOQ ZL -13.0000 -14.0000 CL CL = 24..2000 38.7500 .. 25.0000 XY - .26&5000 XU 282.5000 277.0000 , YU - 47a.500 . YU = o.ooao . .11.00.00. zu = 10... 00.00 ZU -a9000 -14.0000 CU .? .5&040. C.U * 9.ZOQO 9 0000
PERCENT UPPER LOYER- PERCENTHALF ... " ." .. " PERCENTHALF ..ORDCHORO . OR0
CHO R.0 THICK CHORO THICKCHOROTHICKCHOR.0 ... - .. -0.000.. -.... ~ ~.*OOO" - __0*.000.,. 0.00 0.. 0.0 0.00 0.00 -. 50.000 ... 1.500 ... 1.500 .... ". " . - -...... 3.2 r3P -1.. 50 3 2r 59 1 50 JL-O?-C!!O " O.?P_! _q,ooo . 67n.5.0 .L50 67.50 1-50 ." . . . - .. .. lOO.O.0 0.00 lOO..OO o..oo ... ___ "" -. .
- ...... , .- ...... "" .__ . -..... -. .. -. - ...... - ... ." PLANFORM BREAKPOINTS ...... " - ...... 8 Y CHORD XLE XTE...... Y...... ?7.3280 .. 0.0000 . 166.0700 0 77.3PBO ".243..3980 .... .orp.ooo. .??...a280 " 4.9688 166.0700 1 77.3280 Z4.3*39.89 1~6563 83 1040 6.6250 160.1330 77.3289 2 248.3980. . - 3.~3125. 93.1650 9.5100 149.7900 3 77.3280 .. 243.3980.. ..S.968B _. ... 116.9600 . 16.3330 125.3500 '4 83.1040 2433370 .. br 625.0. .. 168.9800 , 31.2500 77.2950 5 80.0799 ... 243-0751 ...... 8.2813 ... 22.5.8100 47.5440 32.6810 6 94.6559 242.91$6 -9.9375 225.8100. ... 27.5450 32.6810 100.4320 7 ... 24297580. .... 11,5938 ... 258.2100 66.2500 14.4450 8 106.2081 242*6014 .... 13.2500 9 111.9843 9 242*4449. . 1489063 1 0 117.7603 10 , 242.371Q . lb.5625 11 123.5362 242.8112. . 18r2188 . 1 2 129.3120 12 243.2515 .1988750 1 3 135.0878 13 . 2.4316917 . 21.5313 . 1 4 140.8637 14 23.1875244.1320 1 5 146.6395 15 244.5722 24.8438 1 6 152.4153 16 26.5000245.0124 1 7 158.1912 17 28.156324594527. 1 8 169.9670 18 245.8929 29.8125 1 9 169.7430 19 . 256.4390 . 31.4688 , 20 175.5196 247.6807 33.1250 21 181.2962 21 .34.7813248.9225 22 187.0729 22 250116f2 3615375 23 192.8495 23 25184059 36.0938.. 24 198 6262 252.6477 . . 39.7500 25 204.4028 . 253.8894 . 41.4063 . 26 210.1795 255.1311 43.0625 27 215.9561 256.3728 44.. 7188 28 221.7328 28 257r6146 . 4b.3750 . 29 226.6523 29 258.8592. . 58.0313 30 229.5211 260.1134 49.6875 31 232.3900 31 261.3675. 51.13438 32 235.2589 32 262.6217 53,0000 33 238.1278 263.8759 . 54.6563 34 240.9967 265.130056.3125 . 35 243.8656 35 266.3842 57.9688 36 246.7345 267.6383 59.6250 3? 249.6033 by.2813.268.8923 38 252.4722 38 270.1467 62.9173 ,. 39 255.3411 39 271.4008 . 64.5931). 40 258.2100 272.6550 66.2 500 .. INBQARD YING END DEFINITION v CHORD . , X Y 1 . ...0.,99 77.328698 4.969090 0.000000 - ...... 2-50...... 81.480430 4.969000 O.OOODO0 .. . . 5.00 . 85.632162 4.969000 0.000000 10100. , 4.96900093.935626 0.000009 . 20.00.' -. 110.542554 4.969000 o..oooooo 30.90 ., 1278149482 4.969000 0.000.000 .. -40..00 4.969000. .143.756411 0.000000 0 00
...... -- . 1501363339 . . 4.9.69000 - . -0sQOOOQQ . _. -. 4.rORIUI-- 176r979267 !t.P69000...... LOOQQQQ...... !t&kWl" 193,577196 . Qe96POOO ..Q~OOOOPO ...... 3LBBbQZL.- .... flQ~164124 .. 4.969000. . O.PQOOO0. .. 3.UB. . 90,OO 226r791052 4.969000 O.OOOOM .. 1~813472.. 4.969000 0. OQOPQQ .... QrQQQQPe...... ". . 10QrOO - Zf3r39.7981
......
TABLE OF INPUl .. ._. - XPCT .. - . 0.000000 2.500000 5.000000
. - - .- - ". . . . 60.000000. - - - - 70.000000 80.000000
. " ..... vti12
.. ._ - ." ...... , 0.0000- 0.000000 .570000 .714000 .8720001.050000 ...... 1- lcsoao ....~tooooo 1.230000. .. 1..24?000 1.170000 .9?7000 54 6OOQ 0.000000~ ...... :
.- - .- ".- .._ ...... - .... - - ..... - ..... -" . .A. . ... " ,0750 -. 0. oooooti 570000 714000 .672000~ 1.~050000 .. 1.145000 . 1.200000 1.230000 . - -. 1.249000 1.17000.0 e937000 546000 0.000000 .
. . " .. - ....
.*!OOO. " 0.. 0.00000 .570000 .714000 872000 ' i.0500oo. ... 1.1C5ooo ...... - 1.zooooo 1.230000 ... .- ... - ..11249000 1 170000 .93?000. 546000 0 .oooooo ". .. ~~72000 ... --.-. .. - .... 550000 1.054000. . - l..lZlbOOO 1.213000 1.235000 1.127000 507000.. . O.-OOOOpD_
550000 715000 1.25.0000 ~1.087000 .8,40000 .
...... - .. - ...... ; 2 7ooo _. ._ ...... - ...... "" .- .- ...... -4717 . -. 0.0.Q0~00 570000 .902000 1.098000 . 1.220000.1.289000 1.315000 ...... 10-2 6?00.0. . 1.105000 .842000 47300.0. 0.000000 . . - ". . - ..... 71.76 ... o:oooooo 580000 729000 .911000 . 1.334000 -. 1.268000 1.343000 1.375000 . . 1.320000 1.155000 .880000 49 5000 0.0.00000 , . . -. . - - .- .._ . - . . ,15 .- -- ." - . .134000 .261000 495000 .880000 1.320000 1.375000. 1.155000 .880000 a495000 0.09000.0..
13 4000 .261000. 1.375000.. 1.155000 ... .eeoooo. , ...... - ....
I
c j c 1 0 ." .osoois -.oleo24
.. " ...... "- .... *030173e027748. 1025354 .. -022959-no20565 LO~UZQ.--.AllS4bb.. .-OlZWh. COO8627 . .005208 -.005038 -e008434 -e011831-.015227 -a018624 -1022020 -.025417 ..
... ". .. -...... YING CA~I~ER SURFACE READ IMTO BASIC .GEOIIEZRZ. ...
.. - ...... - "_ ... ****. .... **** **** **** **** YING **** - ... *+**.. .. .****" ...".4*. . **** ...... REFA 9898.0000 . CBAR = 106,4100 .XBARU. .~BZ.OMP...... - ...... _ - ...... ". - ...... - .- . "- -. .. -...... ". . x0 9 77.3280 X0 * 83.1040 xo. *. 93.1650 YO = 4.9688 YO 6.6250 ..YO :* -_9,5100 ...... - . zo - ~0.0000 20 9 0.00.00 ...... IO ... 0,0000 ..... CHORD * 166.0700 CHORD 9 160.1330 . .. WRRD . * . l49.7900
...." .-. - ...... CAMBER CAMBER HALF-THICKNESS HALF-TH . HALF-THICKNESS ....tz, LOYERUPPER (2) UPPER .. LLPPER ... LOYER 0.0000 0.00000.0000 0.0000 0,0000 .0..0000 0.0000 ....-.OlZZ . .*.ST00 e.5700 . e0528 e 5700 -A~IIO .5500 ...... -.0244 ..7140 -7140 . ,1057 ..71*0 JUO ..1120 " r87-2.0 ,8720 2 3.0 6 ,8720 ",aUO .. .E720 ." i,.psoo ~0500. . , -1,372~ 1.. 0500 .. ..ZD.5.40 1.0540 1 .1450 1.1450 -2r79751.1450. 1.1450 1.1450 .--lL540 . -1.1560 l..2000 1.2000l..2000 -4.3476 1.2000 . ..L2130 . -1m2130 ._ 1..2300 . 1.2300 -5.9249 1.2300 . - LWD 1.2350 1.2490 1.24901.2490 -7.4654 1.2490 . .1..2310, . 1.2370 " " .1-*/700 . 1.1700 ..... :8..9226 .la 1709.. -.-LlZJO .. 111270 -9370 ..e9370 -1012581 s 9370 . -.-A.B~o. . .e830 .a5460 - -95460 -11r4383 .. 5460 . .5.070. . .5010 0.0000 0.0000 -12.4391 0.0.000 . OA1)90 O..OOOO
-...... ". -. .... x9 ...... YO _...... - ... 20 ...... CHORD .. - ...... - ...... -...... - ... " ...... PERCENT-... -. .. - ....CAMBER ..HALF-THICKNESS. ... CAMBER
""_CHOR_D_. "_ ur., ..... UPPER...... LOYER - - " "JZ!. . .. , .. "0.. 0 ...... 0.0000 0.0000 0.0000 0.0000 _. - . ....to ? - e0929. -5500 5500 -0918 ...... 5.0- ...... -1857 . .e7150 e7150 . - . ". 10.0...... ,1361 -8760 e8760 20.0 1.12601.1260-e2993 . -1*1740 ._ l.! i.740 . i.Zzoo .. ... 12350 1 2.350 11_28!?0 -1128.90 - 1.2500 1.2500 ...1.3150 .....1.3150 1.2290 1.2290 1.2620 1,2620 . 1.0870 1.0.870 LlPO... 1.1030 .a400 .e400 .a420 .e420 - .".+?.GO , .4740 ._" . 0 .oooo 0,0000 f .. i
e*** **** ****
..... HALF-THICKNESS CAMBER HILI-THICkSS .. U PP ER LOVERUPPER IZ) UPPER . . LnUER..-.- 0,0000 0.0000 0.0000 0.0000 0.0000 0.0000 ,0327 1340 s 1340 -.024.5 1340 .U+.O .. . ,0653 e2610 ,2610 -a0490 ,2610 ..ZblO e1308 04950 14P50 -no943 e 4910 a4910 r2159 .E800 .8800 1745 .E800 L8800 ,3359 1.1550 1.1550 -a2407 1.15501.1550 r3939 1.3200 1.3200 -e2911 1.28501.2850 e4173 1.3750 1.3750 -a3233 1.3750 1.3750 e4283 1.3200 1.3200 -.3484 1.3200 1.3200 e4242 1.1550 1.1550 -a3672 1.1550 1.1550 4158 .8800 ,8800 -.3795 ,8800 ,8800 e3966 ,4950 ,4950 -a 3822 e 4950 495P e3679 0.0000 0.0000 3751 0.0000 0.0000
-. .. CONFIGURATION DEPENDEAT LOADING GENERATION .. , ...... " ...... - ..- .. " .. - ..... 106.41000 XBAR= 187.00000 MACH NO.. 2.70000 XMAX. 272.65500 NON= 40 CiAR= .. - . .. . " .. TIFtC= -.1.00 TNOM= 0.00 SYMM. 1.00 SHOGO. -0.00 ...... - - - -. - . NOPCT= 12 JBYMAX. 12 RATIO. 4.153854 . .. - . . . . ._ ...... XPCT ye2 ...... 1 0.000 1 0.000 ..... 5 2 5.000 ...... - 2. . 000 .- .... - . - 3 10.000 3 10.000...... 4 .. 20.000 4 20.000 5 30.000 5 30 e000 " 6 40.000 . " b.. 40.000 .. 7.. - 50.000 7 50.000 .. 60iOOO 8 60.000 0 70.000 9 70 .ooo eo.ooo 10 . 80100.0 " ...... 90.000 11 90.090 100.000 12 100.00~0 ...... - ......
L c L c
...... - .. -" .... .- ...... " .-.
...... PLANFORH BREAKPOINTS - ...... - - - ...... -" . "" . .x .. Y 2 E HORD ...... _XLE-.- ..___ATE. .-AULXTL .....
...... " "- ...... "- . ." - .- ..
77;s2.eo" . " . ... 0.0000 0.0000 166.0700 166.0700. . 2 77.3280 2 4.9688 0.0000 166.0700 166.0700 3 83.1040 ..6.6250 0.0000 160.1330 160.1330.
93.1650 9.5100 0.0000 149.7900 149.7900 . ~~~ ~ ~ ~
5 116.9600 16.3330 0.0000 125.3500 125.3500 ... 4 ~ 83.10+0. . - zrJlza70 . __ .2.4i.2370 6 . 168.98.00 . 31.2500 0.0000 77.2950 77.2950 5 .... .BBr0.79.%~4L~L. ~--2.41975l.. .. 7 225.8100 7 41.5440 0.0000 32.6810 32.6810 6 . 94.6559 . 212..9146 .2k2*9146 8 225.8100 47.5450 0.0000 32.6810 32.6810 7 100.4320 .... 2424t.75.80 242-7580 . 9 . 258.2100 66.2500 0.0000 14.4450 14.4450 . -8 .. 106*2081..-24t?.bOlk ... .. 242.6014. 9 ..llls98$3. __ 24Z.*k49 . _.2+2.+.449 10 .. 117~7603... 2+2.3710..... 242.3710 ... 11. - . -12WiUE. LIZdU2 __ .-LCLdllE 12 129.3120 -_ 2132515 - 243.2515. 13 . 135,0878 .. --2$3.6917 . . 243.6917 14 . 140,8637.2!14.1320.. P44..132Q 15 145.6395 . . 2CZ15.722 . 244-5722 16 152.415.3245.0124 . 245.0121 12. .. 159..1912 . 245-45U-2.451i.4527. .... 18 ,163.9670 .. 24.5.8929 245.8929 19 . 169-7430 246.4390 ,246.4390 20 1.75.4196 ... 2C7.6807 . 247.6807 21 1.81.2962 .. 248.9225 24B-9225 22 187.0729 250.1642 250.1642 23 ...... 19.~.aw5. at'tw.9.... - -25~40~9.... 24 199.6262 . 252.b477 .252.bS7.7 25 204.5028 . 253.8894 .253.18894 26 . 210.1795.. . ..255.1311 ... 25511311 ... . 27 ..215>95bl ... L55.3728. ... .25-63728 28 _..22lr7.329, . -2579.6146 .... 257.06146 ... .. 29 .. . 224s6323-...... 258.a85P2 25B~r8592. . 30 229.3231 260.1134 .26O.1134 31 232.3.900 261.3675 261.3675 32 235.2589 ...262r.6217...... 242.6217 33 238.1278. 253r8759 253,8759 34 240.9967 265.1300 265.1300 35 -... .243.8616. .... __ 26613.842.. .. 264.3.8.42. .. 36 246.7345 261-5381 267.6383 37 249.6033 -218.8925 268.8925 38 252.4722 ... 270.1467 270*1467. 39 255.3411 271.4008 271.4008 ... 40 258.2100 .. 272.6550 , 272.6559
...... _. - -. .. - ..... - "" ._ -.... - - .- - ...... FUSELAGE DEFINITION ......
X . . RAD . AREA i! ._
0*00000~ 0.00000 0.00000 16.67000 23.500002.73501 ...... 33.33000 57.500004.27818 -. .. 50rP00D0. 5.32255. 89,.00000 . _. .46.67000 6.10264 117~00000 - . , .-..81.33000 .. 6.33301 126 00000
100.00000" ...... 6.17523 .. 119.80000 116.67000 5.86323 .108.00000 133.33000 -. 5.78122 . 105.00aoo ... 15O~POOOO ....107~000005.83602 . " .. ._ 16 . 5r8.3602 107~00000 - -4 a 5OOOO...... -.. - .- . -...... -18 .... 5.80869 106 .OOOOO . -5r90~~a ...... "-. __- " .. " ..... -20 ... .5.69804 102.00000 . -7-40000 . - "" - . __ 1 21 5.47002 94.00000 ...... -.. "...... ! -8.85000 " 33 5.01463 79 00000 -1Oa25000 ...... - ...... 23 4.33362 59 .OOOQO -11a7000Q...... 26 3.24102 33.00000 -13.20000 . 28 1.59577 8 .ooooo -14.60000 ...... 29 0.00000 0.00000 -15 70000 ..... "...... -...... -. ... NACELLE GEOMETRY . ._ .... ORIGIN '(XIYIZJ X
2 13 .420 00 16.33000213.42000 -5*8OOOO 0.00000 2.00800 15.47000 21.52500 28.01700 32.06700 35.04000
ORIGIN (XsYrZ J X .. 2 18 .670 00 31.25000 -4.9000031.25000 218.67000 0.00000 2.00800 15.47000 21 52500 28.03700 32~06700 35.04000
WING SLO?ESSET TO ZEROFOR UPYASH PRESSURE FIELD SOLUTIpN ..
...
FUSELAGE AREASABOVE AND BEL0.Y WING.
PERCENT CHORD x AREAABOVE , AREA BELOW .. 0.00 79.01 100.08 5.00 88.63 89.74 10.00 96.87 83.91 20.00 112.35 77.52 30.00 128.02 78 63 ...40.00 144.01 84.62 50.00 160.36 91.27 60.,00 176.97 103.09 . ... 70000 193.58 102.02 -. .. eo.00. 210.18 96.69 .... 90000 226.79 .84.77
" .. , .lOP,QP . . 243.40 . 67.e21 . . .- ......
L L w . - ". .- ---_I_.--- ~ . ". . .. FUSELAGEUPYASH ACTING ON YING AT ALPHA. 0.00 DEL XPCT 0.00 10.00 20.00 30.00 50.0040.00 60.00 70.00 w.oa 90.00 100.00 .. ' Y/8/2
0.000 -2.262 "634 -1.447-2.058"298-1.944 -1.834 -1.596 a070 a915 ,025 -2.262 -a634 -.298 -,1.447 -2.058 -1.944 -1.834 -1.596 070 ,915
,050 -. 3.278 4.120 3.761 2.752 2.180 2.164 1.964 21113 2.186 2.346 2.394 e075 3.355 4.123 3.980 3.525 3.356 3.564 3.543 3r882 3.839 3.680 3.433 .loo 3.169 3.464 3.290 3.087 3.066 3.230 3.207 3.473 .3 322 3.035 2.718 '.12 5 2.526 2.577 2.410 2.286 2.272 2.364 2.296 2.467 2.317 2.066 1.829 150 1.923 1.894 1.754 1.678 1.666 1.718 1.637 1.749 1.639 1.451 1.237 e175 1.463 1.408 1.300 1.254 1.244 1.276 1.194 1.264 1.194 1.057 ,869 ,200 1.130 1.066 -986 -960 .953 .974 a912 ,939 ,891 .?96 e645 250 -710 ,645 ,610 603 -602 -624 e581 ,560. s 568 515 .434
300 ,473 e418 e406 .408 ,413 -415 .399 ,373 m 370 .354 e316
350 -327 a289 -290 -294 -300 299 294 ,277 ,266 . .267 m241 -400 a229 ,210 -216 ,220 -226 e227 e226 .E14 ,201 e199 195
,450 160 159 -167 -171 176 ,179 -177 e172 . . 164 0.155 ~153 ...... 500 .122 e125 -132 ,135 e141 e145 ,143 141 135 e130 -123
e550 I097 ,101 .lo8 .llO e116 .120 .119 e117 . .llS * .110 e106
-600 .or9 ,084 -089 -091 -096 -099 .lo1 e099 098 a096 a092 - 700 ,058 062 064 -065 -067 .a70 072 ,974 074 - ,073 072 1 .BOO e039 a041 ,043 -046 ,048 -049 ;050 e051 053 . e055 .. e056
.900 ,024 e026 -027 .028 ,029 -031 -032 e034 035 .a37 1037
.. ". .. - ...... - .- 1.000 .025 -023 .020 .018 e016 -017 e017 .ole rOl8.. * 019 ,020 - ...... 'LIFTING
"- ...... - ...... KPCT , . 9.00 10.00 20.00 .. -...... YIBI2 . . . -...... "- ...... - -." ...... "" 0.006 '.' ,0164 .. ,0353 -mol27 -.0253 i.0243 -e0307 .-.0141. . ~0172 .0223 .... .(UBl. .. -0158 ...... - ...... " . ... . - - - . a.025 ..Ql64 . ,0353 0127 -e0253 -e0243 y.0.307 -a0141 . -&77 .. -0223 ... ..OZB1 -0158 ...... " . .. -...... 050 ... " - .. - - - - .. e0164e0353 -so127 -e0253-e0243-e0307 .-a0141 ,0177 . e0223.ozela0158 ... ,075 . e0164e0353 "0127 ,100 .0404 .0122 -e0176 . - .. -125 0379 -0084 0162 ... . -150 e0362 -0054 -.0153 "0216 -mol93-e0237-10211 ,0124 ,0163 e0261e0216 ...... - ...... -175 e0350 e0029 "0146 -a0200 "0179 -e0205-e0231 . ,011 2 a0148 . e0195 -0217
" . . "...... zoo e0340 e0006 -e0140 -e0188-e0168-e0179 -e0249 a0158a0134 .0101 -0192
.. _. . " .. .. e250 -0277 -e0030 "0133 -.0169-.0153 -mol47 "0233 -..0035 . .0101 . a0126.Ol65
-.0095
-e0085
.- ._ NACELLE PRESSURE..FIELO...... - .- ...... Yl8l2 X,PER CENT CHORD AND PRESSURE-CO€FF.ICIEHT ...... - ...... GLANCE SPLUTION.. . .- ...... _...... NACELLES BELOW UING .. ... 0.000 77.328 243.398
...0.000 100.000 ... 0.00000 0.00000 ,050 77.328 238.690 230.100 238.995 239.290 239.584 239.879 240.174 240.469 240.76k 241.058 241.353 241.648 241.943 242.238 242.533 242.827 243.122 243.417 243.712 - .. - .. 01 000 97.165 97.171 97.349 97.526 97. 704 97.881 98.059 98.236 90.41k 98.591 98.769 98.946 99.124 99.301 99.479 99.656 99.834 100.011 100.189 0.00000 0.00000 ,03894 ,03836 e03778 -03721 ,03663 ,03606 ,03550 ,03493 a03436 .03379 e03322 ,03266 e03209 e03153 -03097 ,03041 e02985 .02929 .loo 83.104 231.692 231.702 232.424 233.146 233.868 234.590 235.312 236.035 236.757 237.479 238.201 238.923 239.645 240.367 241.090 241.812 242.534 243.256 243.875 0.000 92.190 92.796 93.247 93.698 94.149 94.600 95.051 ‘95 502 95.953 96.404 96.855 97.306 97.757 90.208 98.659 99.110 99.561 100.012 100.398 _. 0.00000 0.00000 ,04458 ,04294 -04129 .03967 a03804 e03642 -03481 e03322 e03165 .03010 -02858 e02706 e02556 .02408 -02261 e02115 -01971 -01848 150 94.656 225.394 225.404 226. 499 227.595 228.690 229.786 230.882 231.977 233.073 234.169 235.264 236.360 231.455 238.551 239.641 240.742 241.838 242.934 244.029 0.000 08.182 80.189 88.920 89.667 90.406 91.145 91.884 92.623 93.362 94.101 94.840 95.579 96.318 97.057 97.796 98.535 99.274 100.013 100.752 0.00000 0.00000 e05210 e04913 e04616 -04322 ,04030 e03741 .03461 a03186 e02915 ,02648 e02385 -02226 e01949 -02071 ,.ole21 ,01377 -00937 .00503 0200 106.208 220.585 220.595 221.972 223.348 224.725 226.101 227.470 228.855 230.231 231.608 232.904 234.361 235.738 237.114 238.491 239.067 241.244 242.620 243.846 0.000 03.850 03.866 84.875 8 5.884 06.893 87.903 88.912 89.921 90.930 91.940 92.949 93.958 94.968 95.977 96.986 97.995 99.005 100.014 100.913 0.00000 0.00000 e06088 -05645 e05202 e04762 e04329 ,03906 e03500 03100 4 ,02709 ,02357 -02379 .02120 .0145O .0070 4 a00131 -.00484 -.01032 -.01497 ..
;216 116.926 218.815 210.825 220.294 221.763 223.232 224.701 226.170 227.639 .229.108 230.577 232.047 233.516 234.985 236.454 237.923 239.392 240.861 242.330 243.541 .. L‘ . 0.000 81.261 81.269 82.341 83.612 84.784 85.956 87.127 00.299 ..... 89.470 ... 90.642 91.814 92.985 95.157 95.329 96.500 97.672 98.844 100.015 100.981 ,’ ...... 0;00000 0.00000 e06530 Ob020 05508 .05001 -04503 .04025 .03m .a3100 e02653 -02382 a 92370 ,01693 .00914 e00153 -e00557 -e01187 “01037 -..02k50 .. . - ,292 . .. 1ii.971 218.815. 218.825 220.294 2211763. 232cQ47 ..233*516 234.985 236.454 237.923 239.392 240.861 - ._ 0.000 81.254 81.262 82.434 43.606 84.778 87.122a5.950 88.294 89.463. .90.639 91.811 92.983 94 155 95.327 96.499 97.671 98.843 100.015100.981 ...... 9.00000 0.00000 -06530 .Ob020 a05508 .05001 04503 .04025 - ..03557 -03100 -02653 -02382 .02370 e01693 e00914 ,00153 -a00557 -.01187 -.01837 -.02450 e250 117.760 218.826 218.836 220.308 221.780 223.252 224.724 226.196 227.669 229.141 230.613 232.085 233.557 235.029 236.501 237.974 239.446 240.918 242.39a 243.553 0.000 81.105 81.113 82.294 83.416 84. 657 85.839 87.020 88.201 89,383 90.564 91.746 92r927 94.108 95.290 96.471 97.652 98.834 100.015 100.949 0.00000 0.00000 ,06521 -06017 -05504 0499 6 ,04497 .04018 -03550 .03092 .O2645 e02389 e02360 .01674 ,00895 :00133 -e00575 -.01205 -.01860 "02449
r .3.00 129..312 221.119 221.129 222.710 224.292 225.873 227.455 229.037 230.618 232.200 233.781 235.363 236.945 238.526 239.499 239.509 241.091 242.673 243.534 243.534
0.000 80.575 80.584 81.972 83.360 84.748 86.136 87.524 88.912 90.300 91.688 93.076 94 -4 64 95.852 96.707 96.716 98.104 99.492 100.248 100.248
- .. 0.00000 0.00000 e05970 .05472 e04975 ,04483 ,04004 .03541 .03088 02648 .02271 e02319 r01775 ,01020 e00563 -04786 ,03721 e02746 -02239 a02239
r 350 1$0.864 226.214 226.224 227.505 228.785 230.065 231.346 232.529 232.539 233.819 235.100 236.380 237.661. 238.941 240.222 241.502 242.783 244.063 244.661 244.661
" . - 0*000 .. 8 2 t649 82.659 83.899 85.139 86.379 87.619 - B8.Zb4 88.774 90.014 91.254 92.494 .. . 93.734 94r974 96.214 97.454 98.694 99.934 100.512 100.512
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" ... 92.38? 93.821 95.255 96.689 98 123 99.557 100.992 101.101
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" ..91.833 92 448 92.460 94.503 96.546 98.588 100.589 .1OOe583 ..
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I - . .- " 2220012 223.741 xi5:4e1 227rt16 23.2a.625 242 I 829 242.899 242.e.909 24.4 *.6f4.
68e.596 70.841 13.086 75.331 88.800 9 21-335 95 .hi? 2 ..9 L6 16 979.603. .. - ...... "" ...... " 0.0000.0 0; 00000 .or180 e06511 cQ5840 ,e05180 -02 265. a02813 02116 ... .OlO?E .00069 -..00840 -.00874 -rQQ6V -tQLW ._ 222.794 222.804 224.582 226.361 22.8. 139 z33clau.... 237cQaQ 238rBOB 242.365 344.143 245.921 247q101 247.1l.I ....-...... 0;o-oo - 65.513 65.526 67.991 70.455 72.919 75.304 82.276 85.241 87.705 90..169 92.634 95.098 97.562 99.196 99.210 100.1.96
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187.013 , 227.153 227.163 228.602 230.041 231.480 232.918 234..357 235e.79.6 237.235 238.673 240.112 241.551 242.989 244.428 245.861 241.306 248.744 250.183 290r576
0 .ooo 63.528 63.544 65.824 68.104 70.385 72.665 74.946 77,226 79.506 81.787 84.067
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.. ... - - ......
198.626 233.415 233.425 234.627 235.830 237.032 238.235 239.438 240.640 241.843 243.046 244.248 245.451 246.654 241.056 249.059 250.261 251.464 252.667 253.869 ." ...... 0.000 64.397 64.416 66.642 68.868 71.095 73.321 75.547 17.773 79.999 82.226 84.452 86.678 88.904 91.130 93.357 95.583 91.809 100.035 102.261 -. ...
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.Of301. .02128 -01957 ,01788 -01657 e01662 ,01647 a01513 . " " .... 221.733 247,875 247.805 248.494 249.103 249.112 250.322. 252.15q 252.159 253.368 253.978 254.587 255.196 255.806 256.415 257.024 257.634 ...... - 0.000 72.855 12.883 74.581 76.219 77i97.7 79 L b76 88.166 8P 8 64 91.562 93 260 94.959 96.651 98.355 lQQ..O51 " .- . ... - Q*OOOOO 0.00000 ,03650 03540 ,03430 .0332i ,03212 .026?4 02570 e02467 ~02365 ,02263 e02162 .02062 . a01962
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. -050 -e00395 e00744 -022C7a02593e02319.01244 . . .a02428 . e03304 .0305e .OZ208 001599 ,01113 .. . .Ol288..Oll26 r075 no1662 mO2464 .02682 .02827-02528 mO2896 a02825 0209 3 e01320 ,00905e01320 .. .loo . . ,044~9 ,03321e02209 .02210 -02196 a02565 a02595. a02454 a02309 e01971 ,01148 e00787
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. .. *+*,************************************-**~.,*.*h~*~*.*..+.2.*.~**.~t******** TEA253,17 LOADING VERSION OF OCTOBER 901 1Q75. . - ...... -.
...... - - " ....- .... OPTIHUH COHBINATION OF 17 UING LOADU~G~...... - ...... 969-50017 LOAD CHECK CASE 22 SPANSTA. UITH FUSELAGE AND Z TERMS .. NUHBER OF PLANFORHBREAKPOINTS = 9.0 FLAT. PLATE UINTRRL.FLAG * 0.0 ._ NUMBER OF SEHISPANELEHENTS * 40.0 ...... PRINT-FLAG * . 2.0 NUHEER. OF SPAN STATIONS FOR CAHBER SURFACE = 22.0 .. ..SHOOTHING. FLAG * ... 1.0 SPANSTATION FOR PARABOLIC APEX = -0.0 . .RESTARTFLAG = . 1.0
. . -...... " .. BASICNUHBER HACH = 2 7000 ...... -.DESIGN f-L: .... .loo0 . . ". CBAR * 106.4100 ...... -. .... NUMBER OF LOADINGS. = - -. -17.0000 ...... PITCHING HOHENT ATCENTER 187.0000 NUMBER aF. _CAMBER ORDINATES. !. ... 12.0000 . - .. REFERENCE AREA = 9898.0000NUWER DF.?DlNiS DEFINING ARBITRARY.R€GlON. = . .2.0000 __.. C-H-0 CONSTRAINT = -0060 .. - __... .--EUSELU;E.ALPHI-* ..... 0.0000 . .- SPANSTATION FOR SIDE-OF-BODY . 4.9688 NUHBER. OF..BOOY LAMER ORDLNAES .= . . 19.0000
...... -...... "
......
CAMEER- SURFACE OPTION FLAGS = 1.0. 1.0 I*O. L,.O .... .- .... "" ...... t' .
c. N w c td P
. ." . ". . " -. "" - ...... - .. - ...... - . 5.CONSTRAINTS ARE APPLIED0.N-OROINATE ......
" CONSTRAINT LOCATIONS...... - . .- .-..:
...... 1 ...130.850000 . 4.968600-4..070000 2.-10.160000 . 4.968800.189.000000 .. 3 , 243.390000 4.968800 -14.110000 4 6.625000189.000000 -8.320000 . - ..... 5 ,. __ , 189.0000008.281300 -7.000000
.. " PLANFORMDEFINITION
X X .- .. (LEADINGEDGE1 Y CHORD EDGE1(TRAILING
.. ... " ...... " ...... 1- 59.999300 0.000000 243.88'1000.183.881700 .. -2 77.328000 4.968800 166.070000 243.398000 3 83.104000 243.237000160.1330006.625000
.... 4. 93.165000 9 510000149.790000 242.955000 -5 116.960000 16.333000125.350000242.310000 ..... 6 168.980000 31.250000 246.27500077.295000 7 225.810000 47.544000 32.681000 258.491000 .... 8 225.810000 47.545000 32.681000 258.491000 9 258.210000 66.250000 14.445000 272.655000
ORDINATESFOR BODY CANBER LINE ... I. X 2 I X z I X z . .I X .. z 1 0.00000 10.00000 2 16.67000 3 8.55000 33.33000 7.10000 4 50.00000. 5.64000 5 66.67000 5 4.17000 6 83.33000 7 2.73000 100~00000 . 1.28000 8 116.67000 . -a14000 9 133.33000 -1.60000 10 150.00000 -3.04000 11 166.66000-4.50000 12 183.33000 -5.90000 13 200.00000-7.40000 13 14 216.67000 -8,85000 15 233.33000-10.25000 16. 250.00000 -11.70000 17 264.67OQO -13.2000018 283.30000 -14.6000019 295.00000-15.70000...... 1
.. . . - - -- -. -. -. - "" - . . - - - -" "- ...... VALUES OF SEMISPAN LOCATION AT YHICHWING .CARBBEB~~~PEACL.HLLBL.ULCULAED .... _- ...... -...... ~0.0000 , 1.0000 2.0000 3 I 0000 4.0000.. 5.000D.. ... bO.QOQ- ... .L.OOOO .... .OORD. ...10.0000 __ 12qOOOO14.0000 16.0000 19.0600 22.0000 __ 25a00D0 ._~~20LooQO~~"1(2rpoOO. M..- ._ 36.0000
3B.0900 .- 40eOOOO ...... " ..... "...... -...... - - ...... WING GRID SYSTEMPUTS SIDE-OF-FUSELAGE AT Y= 4.14063 ATEDGE OF ELEHELLI_RRY= 3 ...... G SPANSTATION OF ORDIMATECONSTRAINT 1 IS CHANGEOFROM 4,96880 .TO ...... 4.96871 ...... __.I "" - . - _. ". . - . ." " . SPANSTATION OF ORDINATECONSTRAINT 2 IS CHANGED FROH .4.96880 TI1... 4196875 ...... SPANSTATION OF ORDINATECONSTRAINT 3 IS CHANGED FROM 4.9680.0~TO .. $196835 ". -. . .. SPAN STATION OF ORDINATE CONSTRAINT 5 IS CHANGED FROM e.zeno. TO . . e.zsits ...... - ... LOADING .I FOR THIS CASE IS UNIFORM OR CONSTANT (LOADING i IH~HELOADING DEFINITIONS) LOADING 2 FOR THIS CASE IS LJNEAR CHORDYISE (LOADING 2 IN THE LOADINGDEFINITIDNS) LOAOING 3 FOR THIS CASE IS LINEARSPANUIS€ (LOADING. 3 IN THE LOADING.DEFINITIONS) LOADING 4 FOR THIS CASE IS QUADRATICSPANYISE (LOADING 4 IN THE LOADINGDEFINITIDNS) LOADING 5 FOR THIS CASE IS QUADRATICCHOROYISE (LOADING 5 IN THE LOAOINGDEFINITIONSJ LOADING 6 FOR. THIS CASE IS PARABOLICCHOROYISE (LOADING 6 IN I.HE LOADINGDEFINITIONS) LOADING 7 FOR THIS CASE IS CUBICCHOROYISE (LOADING 7 IN THE LDADINGDEFINITIDNS) LOADING B,FDRTHIS CASE IS SIRILAR TO FLAT WING (LOADING B IN THE LOADINGDEFINITIONS) LOAOING 9 FOR THIS CASE IS SQ. ROOTFROM 1. E. (LOADING 9 IN THE LOADINGDEFINITIONS) LOAOING 10 FOR THIS CASE IS ELLIPTICAL C-SUB-P (LOADING 10 IN THE LOADINGDEFINITIONS) LOADING 11 FOR THIS CASE IS LINEAR IN ARBeREGIDN(LOADING 11 IN THE LOADINGDEFINITIDNSJ LOADING 12 FOR THIS CASE IS BODY UPYASH LOAUING (LOADING 16 IN 1HE.LOADINGDEFINITIONSJ LOADING 13 FOR THIS CASE IS NACELLE BUOYANCY (LOADING 17 IN THE LOADINGOEFINITIONS) LOADING 14 FOR THIS CASE IS NACELLEBUDY(CAMBER) (LOADING 14 IN THE LOADINGDEFINITIONS) LOAOING 15 FOR THIS CASE IS BODY UPYASH(CAMBER) (LOADING 13 IN THE LOADINGOEFINITIONS) LOAOIN6 16 FOR THIS CASE IS BDOY BUOYANCY TERM (LOADING 15 IH THE LOADINGDEFINITIONS) _... LOADING 17 FOR THIS CASE IS BODYBUOY. (CAMBER)(LOADING 12 IN THE LOADINGDEFINITIONS)
" . " XIC(PERCENT) FOR INTERPOLATED CAHBER SURFACE ORDINATES.
.. ". D€~ll4lTI~NOF ARBITRARYREGION FOR LOADING 11. . - . Y 0.00000 66.25000 x- 207.00000269.80000
,ARBITRARY RECXON DEFINITION(LOAOING 11)
F~ACTION~OFSEMISPAN 0.00000 1.OOOOO FRACTION.OFLOCAL CHORD . .79943 -80235 ...... "... - "...... NACELLE NUMBER 1, ORIGIN AT X . 213.42000000 ...... " ... -. Y. lbm33000000 -...... -.... - "....
2- -5m80000000 ...... " ......
...... " . - -...... NACELLELONGITUDINAL COORDINATES (X HAS BEENMULTIPLIED BY 1.004000401 ......
" . ._ 0.000000 2.008000 15.41000021.525000 28.017000 32.0671)OO . 15.040000 . ." ...... NACELLERADII (R HAS BEEN MULTIPLIED BY 1.000000001 ...... " ...... ". -
2.8650002.983000 3.6330003.7700003.654000 3.420000.. 3.420000 . . . " ......
NACELLE X AND RADIUS TABLESEXPANDED TO 40 ENTRIES BY LINEARINTERPOLATI.ON, AND x HASBEEN-TRANSLATED BY THE ORIGIN-X~.:
.. ~ ...... NACELLELONGITUDINAL COORDINATES (x HAS BEEN MULTIPLIED BY 1.00000000)~~
213.420000214.318462 215.216923 216.115385'217.013846 217.912308 218.810769219.709231 220.607b92 221.506154 222.404615223.303077 224.201538 225.100000 225,998462 226,896923 227.795385 228.693816 229.592308 230.490769 231.389231232.287692 233.186154 234.084615234.903077 235.881538 236.780000237.678462 238.576923 239.474385 240,373846 t41.272308 242.170769 243.069231243,967692 244.866154 245.76.4615. 246.66307_7 .247.561538 24B.460000
NACELLERADIIHAS BEEN(R MULTIPLIED BY 1.00000000l
2.8650002.9177982.9705963.0161903.0595713.1029523.146334 3.189715 3.2330973.276478 3.3198593.3632413.4066223.4500033.4933853.5367663.580148 3.623529- 3.648890 3.669219 3.6895473.7098763.7302043.7505333.7693203.753266 3.737212 3.721158 3.7051043.689050 3.b729973.6569433.6116043.5596933.5077823.455871 3.420000 3.4200003.420000
...... "" ...... " N4CELLE NUHBER 2, ORIGIN AT X 218.67000000 ...... - ". " ... Y= 31.25000000 ...... 1. -4~90000000 ...... - ...... - .. NACELLELONGITUDINAL COORDINATES (X HASBEEN NULTIPLIEO BY 1.040400001 . " . - ...... o.oooooo 2.00800015.4toooo 21.mooo 28.017000 3z.abtaoa .. -35.04ooaa ......
.... -. ." .... .__. , - ._. NACELLERADIIHASBEEN (R MULTIPLIED BY 1.00000000) ..... ". .... -. -- ..... - .. . .- ..- ... ".- - " 2L~65000 2.983000 3,633000 3.770000 .3.654000 ZOODO. DO. 3.424oon ...... "...... "......
" . .- . ... - . - -...... AT PRESSURESBODY BUOYANCY THE-EOLLOYING XIC_LP€RCENTTL .. ". . . 0.00000 10.00000 20.0000030.00000 .-+O~~OQOOO . .50~OQ000... bO9OQOQ-....~. 00 Q"... -BO~OPQOO 90.00MO .... 1.~.0.*.00P00._. . - _. - - - . -......
0.00000 2.50000 5.00000 7.50000 10.00000 12.50000 .l5.0POOO ' . i2.5ROOO -20;OMOo 25-00000 30.00000 35.00000 40.00000 45.00000 50.00000 55 .OOOOO 6000Q000 7OaOOOOO. BO*QOOOO 90.00000 10,0.00000 , ...... BODY PRESSURES ON THEYING UPPER SURFACE
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."
UPPE.R YING. SVRCACELIMITING CP TABLES .. X STAIIONS. 0.00000 100.00000
Y STATIONS 0.00000 100.00000 LIMIT E-P .... --.137000- .-.137000 -e137000 -.137000
.... " _. . *******************************************************************************************.+.*~#~#~~~~*~*~~**+*f~+#*+*#*~.*+*~*~*~~ THE MAXIMUM AND MINIMUM OF PRECEDINGTHE ARRAYARE -~13700000~-.-~13700000 ...... *. -: .c-P- CQNF~TLIDINALGRADIENT LIMIT . -. - .. , - .002500.002500 .. .002500 .002500 .... - ... - ... - ...... ~~***********++**.*i******.*********************+**********************************************~*~~~*+++~*~*~*****.~~~*********f+++*~_. . __ - ...... THE MAXIMUM AND RINIHUH OF THEPRECEDING ARRAY ARE._. ... __ .. _.QOZUU)PIL.-"-.-OD ...... c. w N .- . . . - . .- .- . .. .- .. - . . XICIPERCENT) FOR YING THICKNEG PRESSUR€. COEFFiCIENT.. _. . - . . ._ . .- . - " .. .. . ".. ~"" 0.00000 5.00000 10.00000 15.00000 . 20.00000 25~00000 30.oooo~. ~LOOOOO u).aonon.-. ~~ooooo. _...... _ . 50.00000 , 55.00000 60.00000 65.00000 70.00000 75~00000 8o.ooam. a5.0~00. . P~OOQM. -.~~.oooM ...... 100.00000 . .. - - .- . .. - . . - - .. .- -.. .
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FLAT WINGFORCE COEFFICIENTS C C XF M D c. e027674 C m .0004830 -- ,720744 -- 9 -e089414 "- = -630682 1 0 L C 2 L C L
CAR0 9 PARAMETERS.
XF - a72074445 SCL9 9 .02767366 KF - ,63068237 AREA9 781.13547925 FACTOR = ' 1.08589045 w P ...... - FUSELAGE CONTRIEUTION AND CARRY-OVER LIFT . . - ...... " FUSELAGECLO -mOOOOO COO .000001 CMO a00396 XACIXMAXO -12239 . CMOO .e00396 . . . _. - ...... CARRY-OVERCPFOR LOADING 1 OFTHIS CASE (UNIFORM DRCONSTANT J
XPCT 0.005.00 2.50 10.00 15.00 20.00 30.00 40.ilO 50.00 60.00 70.00 80.00 90.00 100.00
" YI012
0.000 0.00000 0.00000 0.00000 0.00000 .09609 62278 .97071 8652 3 .91b22.E7947 e9 2619 -92993 -93711-92993 e92619 .94577
025 0.00000 0.00000 0.00000 .Ob494 -42603 .78994 95960 86806 -91949e89054 ' ,929 96 ,93619,92996 .94577 ,94861
,050 0.00000 0.00000 ;12231 e51132 .68960 e92131 695010 -8985 1 -94254-91921 e94951 .9534b e95621 ,96248
CARRY-OVERCP FOR LOADING 2 OF THISCASE ( LINEARCHOROYISE 1
2 .50 15.00 20.00 15.00XPCT 0.00 2.50 5.00 10.00 30.00 40.00 50.00 60.00 70.00 80.00 90.00 100.00
Y1812
0.000 0.00000 0.00000 0.00000 0.00000 ,01658.00142 36442 -67677 1.27295,96826 1.58 644 1.90209 2.22009 2.525352.220091.90209 1.58644
-025 0.00000 0.00000 0.00000 e00096e45471 ,14744 ,04115 -75185 1.336791.03801 1.64 239 1.94942 2.258641.949421.64239 2.55742
050 0.00000 0.00000 .00180 .Ob077-59490 -28976 -15679 e88633 1.17484 1 47197 1.71 305 2.07520 2.37899 2.674472.378992.07520 1.71305
CARRY-OVERCPFOR LOADING 3 OFTHIS CASE ' LIN€ARSPANYISE
XPCT 0.00 2.50 5.00 10.00 15.00 20.00 30.00 40.00 50.00 60.00 70.00 00.00 90.00 100.00
Y1812
0.000 0.00000 0.00000 0.00000 0.00000 a02162 -3220514013e29815 -26940 e25787 ,34178 .35670,34178 ,30695e37042 ?: ?: , '
.025 0.000000.00000 0.00000 .01411 .09586 .18720 -26451 -27221 . .. -2969.6e.31657 -33648 -35249 ,37757.36906
-050 0.00000 Om00000 e02752 e11505 a22602,16473 .28e26853-26219 577 30381 -31805 e32886 e33591 ,34163 -. .. - ...... - ... ._...... - . .. - .. CASETHISLOADING CARRY-OVER4-OF FOR CP ( QUADRATLTfUtiUSLL. __ .. __ ...... - .. - ... - ...... -......
XPCT 0.00 15.0010.00 5.00 2; 50 2QmQQ-.. -30.00 -. - IO.& ~ . 150.00.. 60.00 70.00 80.00100.00 90.00 ...... - .....- ..
" ...... - ... Y1812 . - . -- - . - -...... -. .. - ......
0.000 0.00000 0.00000 0.00000 0.00000 e00324 e06160.-04770.02102 . e07942 -09385 e1 0975 el2440 .13857el2440 e10975 e15651
-025 0.00000 0.00000 0.00000 e00219 ,02997,01438 . a05147 . -06361 . .or849 09052 e10630-14856 ,13789 e12132
e050 0.00000 0.00000 ,01726e00413 e02662 -03812 a05191 e06051 07071 -08193 .ome 10402 .11211 12475 ... CARRY-OVERCPFOR LOADING 5 OF THISCASE (PUAO.RATIC CHDRDYISE I
X P.CT 0.00 10.00 2.50 5.00 15.00 20.00 30.00 40.00 50 .OO 60.00 70.00 80.00 90.00 100.00
Y!BlZ
0.000 0.00000 0.00000 0.00000 0.00000 -09921-00553 .00001 1.21733a70923 33698 1.86474 2.65610 3.59019 4.613683.59019 2.65610 1.86474
e025 0.00000 0.00000 0.00000 .00001 ,003371.33146a.81145 a41932 e15315 02257 1.98557 2.77348 3.69595 4.709293.69595 2.77348 1.98557
,050 0.00000 0.00000 ,02780-00531.00002 07416 1.575171.01419,57725 e26222 2.26305 3.07872 4.02271 5.059404.02271 3.07872 2.26305
CARRY-OVERCPFOR LOADING 6 OF THISCASE (PARABOLIC CHOROHISE I .: .
2.50 5.00 10.00 15.00 20.00 30.00 20.00 15.00XPCT 10.00. 0.00 5.00 2.50 40.00 50.00 60.00
70.00 80.00 100.00 90.00
Y1812 J
0.000 0.00000 0.00000 0.000000.00000 2.044781.28034a26154.00578 . 2.598062.43691 2.50492 2.12203 1.454582.12203 2.50492 .55339
025 0.00000 0.00000 0.00000 ,00390 -16177-556811.53498 . .2.178832.51116 . 2.62806 2.49415 2.08979 1.417602.08979 2.49415 -51820 ..... - .!. '. - -050 0.00000 0.00000,58449 ,23820 e00735 1.03169 2.71~732.673412.411~3 1.883a3 2.51129 2.05063 1.33468 .396421.33468 2.05063 2.51129 c w m
..... - ...... CARRY-OVER CP FOR LOADX~G. .. ._...... - ...... _ .... XPCT - 0.00 2.50 5.00 70 00. 80.00 90.00 . ._ YlBt2 .. ". . ._. 0 0 0.00 Q.OOO.00 0.00000 0.00000 2.80113 . 3.26304 3.97682
,025 0.00000 0.00000 0.00000 -...... 2.92903 . 3.56362 4.0513L ,050 0100000 ,050 0.00000 .00005 3.2387 5 3.849293.23875 4.30423
CARRY-OVER CP FOR LOADING ....
..... - " XPCT 0.00 2.50 5.00 10.00 20.0015.00 . 39.00. ... 40.00 .. 50.00 60.00 . .70*00 80.00 90.00 100.00 ...... - ...... YlBl2 ..... -.
.- - - .- . ._ . " .....- . " .. - ...... -...... - ...... 0,000 Q100000 0.00000 0.00000 0.00000 e232011.27193.1.45810 ...93603 . .a5827 82254 ,74669 068248 e63953068248 ,74669 -61009 ... a025 0.00000 0.00000 0~0000~a90573i.32356 1.45715 ,88187 ,15679 e8035785082 ...... *nbt5 . .6m2 ,65017 -61093 - ..... - ......
e050 0.00000 .0.00000 ,29531 1.181721.497401.213631.02276 90264 84564 .80039 e73832 a64407e68818 e61603
...... " - .....
...... - .... "" ...... CARRY-OVER CP FOR LOADING 9 OF THIS CASE (sa. ROOT FROM T. E.. 1...... - - ...... - ..... - - .. - ... - .... - -...... XPtT '_ . 0.60 2.50 . . 5.00 60.00 TO. 00 80.00 90.00
'_ YI0i2 .. -. .- .. _.._ .... - . - .... . : "C!..OOO 0.~0OQOb 0.00000 - 0.00000 1.29083 , 1.16160 . 1.00145 , ,81854 ...... rO25 .9:+09.099... OtOOOOO. . QaQOOQO. .. 1.2741% ...... - .Ll470.7 . . -9933.7.81614 ...... ". - -. ." -. .." . .- ...... -- - __ - ...... -.... " ... 005Q . - 0cO.OQOD . 0.00000 .2 4008 1i27407 ...... - 1*.1406B 1.98235 -79139...... - ... 1
...... - "...... __ "" ...... €AR_RY-OYER CP FUR LOADING 11 OF THIL CASE . LLWEAR. IN-ARBIRLGLII~(L.-.L ..
...... - .- ...... "...... - ...... - ...... I ..
. _" -.. -...... - .... "" ...... -. , . ." . . - .- . .- - .-. - . ." -". .- . . 0.000, ,_0,00000 - 0.00000 0.00000 0.00000 . - .. o.oooo.o . O_.OOOOQ .. ~o_Iooooo.... Q..QQQOO. .. oroooao. o.ooooo ... , ...0.00000 __ 0.00000 e02361 e 32718 ...... - ...... --. - ...... __... I.
- ...... - .- .... -...... - .. -. "...... - .... - ..... - - __ .- ... - ...... - ... * . . --O~DOODO. o.ooooo- . . 0.ooOoo . . o.00000 0.00000.. . D*nnOM . -9L000nQ.-.AQOnan .... aAnooo1. 0.00000 ..... -...... PrQP0.00 . 0.00000 e04937 a34620 ...... __.._ ...... ~ ...... - ...... -. - ...... -- ...... -- ...... "... "-rO3Q .--B,QOOO& O..OOOOO . 0.00000 0.00000 0.00000 - - -Q.M000"hWDQD--"O~O000~ -Qco0000 0.00000 .... - Q.00009. . 0.QOOOQ. ... -12444. . e41148 ...... ". - - ..... - ... -... "" .. -- ...... -...... - ...... -. .... CARRY-OVER te .FOR .LOADING 14 b~ THIS
...... - ..... "_ ... -...... - ...... U..Oe. 2Oa"4UU----SLOD--- 60.00
- ...-...... - ......
...... -...... " .. - -.PrPPO Q.MPDD .. ..O~OOOOO 0.0.0000 .O.OOOOO ...... 9.0.00(10 . 0~00000 0.OOOOQ . ,00544 ...... - ...... - .- .. .. lQZ5- ... .QrO.QOW . .0.00000 ... 0.00000 ... Q.OOOOd - ...... PLO.QQOO ...... O..O.ODPQ . .o.aoooo ... _.ooan -...... - -...... -...... "" .. -. - .... "_" ... -XPCTLQL~&O LOQ ______LQrQO 30~00...... 90.OQ .lO.O&Q"-. -
"...... ---__..____. ..
.. ~ ...... - ...- .... "... " ... -.... - ... -. " .. ,......
... -. - ...... - ...... ". -. ... ____ _" -- . - .... - ...... ""
...... ____ ...... " "_ "" ......
. " . - . , ...... - . . ... - - . .. . ."" .- .. . .._ ...... YING DATA FOR UNIFORH OR CONSTANT LOADING . -. Q69-500..1? LOAD.CHECK CASE 22 SPAN STA. YITH €USELAGE AND 2 TERM . ... .- SPANYISE OISTRIBUTION OF SECTION DRAG, LIFT, AN0 PITCHING HOHENT
Y SECTION SECTION STC TION -" b C C 812 CHORD 0 L M .. . 0.0000000 183.8817000 0.0000000 7409501 -.6845282 .0250000 177.9445264 0.0000000 e 7758395 -a7372558 .0500000 172.0073528 0.0000000 e8387086 -a8141165 .0750000 166.0701792 3.3007611 1~0000000 -.9656284 .1000000 160~1330000 1.6568419 1~0000000 -1.0192347 e1250000 154.1951859 1.2350506 1~0000000 -1.0764594 .1500000 148.2586941 1.0146702 1~0000000 -1.1383784 .1750000 142.3260032 .8688242 1.0000000-1.2058524 ,2000000 136.3933123 e7616182 1~0000000 -1.27'e7008 .2300000 122.6106675 ,6020349 1~0000000 -1.4451936 .3000000 113.9394744 4860737 1~0000000 -1.6352923 e3500000 103.2682813 ,3920641 1~0000000 -1.8640200 4000000 92.5970882 e3116704 1~0000000 -2.1461007 e4750000 76.6960487 e1999098 1~0000000 -2.7130599 e5500000 63.0912977 e1024884 1~0000000 -3.4670189 6250000 49.4865467 .0054721 1~0000000 -4.6326307 .?000000 35.8817958 -e1934528 1~0000000 -6.6821074 .7500000 30.5922200 -.0261647 1~0000000 -8.0063824 .8000000 27.3627760 .1156497 1.0000000 -9.0987203 .9000000 20.9038880 .2639022 1~0000000 -12.3067406 .9500000 17.6744440 .2982005 1~0000000 -14.7858920 1.0000000 14.4450000 e1423389 1.0000000.. -18.3990688 X CP c - ,938399 c- -647121 --- * .714229 K -734870 L 0 L E S C REF M "-- - e920569 -- * -e072721 C * .019612 S C M PROG L 0
INTERFERENCE DRAG OF LOADING 2 ( LINEAR CHOROYISE 1 ON LOADING 1 (UNIFORM OR CONSTANT 1 IS .10231462E+01 INTERFERENCE DRAG OF LOADING 3 ( LINEAR SPANYISE ) ON LOADING 1 (UNIFORM OR CONSTANT 1 IS ,35038758Et00 INTERFERENCE DRAG OF LOADING 4 ( QUADRATIC SPANYISE 1 ON LOADING 1 (UNIFORH OR CONSTANT IS .20308055E+00 .INTERFERENCE DRAG OF LOADING 5 (QUADRATIC CHOROYISE I ON LOADING 1 (UNIFORH OR CONSTANT 1 IS .12990764E+01 INTERFERENCE DRAG OF LOADING 6 (PARABOLIC CHOROYISE 1 ON LOADING 1 IUNIFORM OR CONSTANT 1 IS .10857848EtOl 1NTER.FERENCE DRAG OF LOADING 7 I CUBIC CHORDYISE ON LOADING 1 IUNIFORM OR CONSTANT 1 1s .14564008Et01 INTERFERENCE DRAG OF LOADING 8 (SIMILAR TO FLATYING) ON LOADING 1 (UNIFORH OR CONSTANT 1 IS .54114581E+00 INTERFERENCE DRAG OF LOADING 9 (sa. ROOT FROM T. E. ON LOADING 1 (UNIFORM OR CONSTANT 1 IS .72564265E+00 INTERFERENCE DRAG OF LOADING 10 ( ELLIPTICAL C-SUB-P I ON LOADING 1 (UNIFORH OR CONSTANT 1 IS ..69049248EtOO INTERFERENCE DRAG OF LOADING 11 (LINEAR IN ARBeREGION) ON LOADING 1 (UNIFORM OR CONSTANT 1 IS .47908493E-01 INTERFERENCE DRAG OF LOADING 12 (BODY UPYASH LOADING 1 ON LOADING 1 (UNIFORM OR CONSTANT 1 I5 .12048131E-01 INTERFERENCE DRAG OF LOADING 13 ( NACELLE BUOYANCY 1 ON LOADING 1 (UNIFORH OR. CONSTANT I IS .38898304E-02 INTERFERENCE DRAG OF LOADING 14 (NACELLE BUOY(CAM8ER)) ON LOADING 1 (UNIFORM OR CONSTANT ) IS .38898304E-02 INTERFERENCE DRAG OF LOADING 15 IBODY UPYASH (CAMBER)) ON LOADING 1 (UNIFORM OR CONSTANT 1 IS .12048131E-01 INTERFERENCE DRAG OF LOADING 16 ( BODY BUOYANCYTERM 1 ON LOAOING 1 (UNIFORM .OR. CONSTANT IS -.12190312E-02 INTERFERENCE DRAG OF LOADING 17 (BODY BUOY. (CAMBER) ) ON LOADING L (UNIFORM OR CONSTANT 1. IS -.12190312E-02 L P 0
...... - ...... Y .. SECTION .. ..." .- C . ." 012 CHORD D ...... - .." . "" . -. - " ... ,.0.0000000.~ 183.8817000 0.0000000. . .le0303936-1.0852D44...... 0250000 177.9445264 0.0000000 1.0841903rl.1771041.. .. . -...... -. "" ..0500000 17210073528 0~00D0000 . 1,2019202 -1.140DO52 ...0750000 166.0701792 9.2051767 . 1.3353591 ... .-1.7384780 ...... 1000000 160.1330000 4.2757079 1.4812924 .-.1.7556970
- . ._" .125QOOO . 154.1951859 2.9089002 1.4257.731. .. .-1.7721179 .... -- .. . .1500000. 148.2586941 2.1906517 1.3705690 . -1.7887482 . .... "...... 1750000 142.3260032 1.6984455 1.3164540 -1.8061001 .... .2000000 136.3933123 1.3400430 1.2610898 .. .-1.822712Q ...... 2500000 124.6106675 ,8456530 . 1.1525092-1.8573530 ... .3000000 113.9394744 .5429794 1.0541730.. -1.8985884 q3500.900 . 103.2682813 .3337833 ,9545366...... -le9387352 ...... 4000000 92.5970882 -1806995 . .8562689. . -1.9807214 .4750000 76.6960487 ,0447455 .7089300 -2.0419749 .5500000 I 63.0912977 -e0230816 e5855307 -2.1250335- . .6250000 49.4865467 -.OS32899 ,4595165 ,. -2.2041253 . . ,7000000 35.8817958 -.Ob37878 . -3334628. . ...- 2.2033479 ___ ..*7500000 30.5922200 -.0215506 ...... 2845141 .. -.-2.3231477. .. . .8000000 27.3627760 e0018451 -2534699 -2.3491939 ~9OOODOO 20.9038880 .0104310 . ..1946102 -2.4256708 .. .9500000 17.6744440 .0121128 -1638388 . . -2.4521495 1.0000000 14.4450009 .0058128 . .138771e.- . WQO~.+...... "_ . X
" ce . - ..- ...... - c. = ...... L ...... _. P. . L - .. REF
-a155923 C -e056583 . "
C M . ", .. L 0
TNTERFE REN C E DRAG OF LOADING 1 (UNIFORM OR CONSTANT 1 ON LOAOING INTERFE RENIC E DRAG OF LOADING 3 ( LINEARSPANYISE I LOADING -. ON INTERFERENIC E DRAG OF LOADING 4 ( QUADRATICSPANYISE I ON LOADING . INTEPEESENIC E DRAG OF LOADING 5 (QUADRATICCHORDYISE.) ON LOADING . . INTERFERENIC E DRAG OF LOADING 6 (PARABRLICCHOROYISE 1 ON LOADING .. - INTERFEREN'C E DRAG OF LOADING 7 ( CUBICCHORDYISE I ON LOADING INTERFERENC. E DRAG OF LOADING 8 (SIMILAR TO FLATYINGI ON LOADING - INTERFERENCE DRAG OF LOADING 9 (SQ..ROOTFRON T. E. 1 ON LOADING, INTER_FERENCE DRAG OF ,LOADING 10 ( ELCIPT1CAL.C-SUB-P L OM LOARXNG ___ WIEREERENCE. DRAG OF LOADING 11 (LINEAR IN.ARBeREG.IOH1 W .LMPLHG !N.TEEFERENCE DRAG UF LOADING 12 (BODY UPYASHLOADIH6.I.OH-LOADIN& INTERFERENCE ORA6 OF LOADING 13 ( NACELLEBUPYANCY. .I .ON LOAD.ILIG . LNTEREEBENCE PRAG OF LOADING 14 (NACELLEBUOY(CAMRER)I 0N.LOAQING .. INTERFERENCE OR4G OF LOADING 15 (BODY UPYASH(CAMBER)) 0N.LOARING . ..INTERFERENCE DRAG OF LOADING 16 ( BODY BUOYANCY T€RH .I ON..LOADING .INTEMEREHCE DRAG OF LOADING 17 (BODY BUOY. (CAHBEEI .1.OK.L"G 1 .. . -.. . . ." . . . .. I -- . .. - - . .- .. . . - " . -...... - ...... - - - .- .. . .. -. .. -. .. . , YINb DATA FOR LINEAR SPANYISE . LOADING . - . .-.. +..- . , -...... - v 9.69-500 17 LOADCHECK CASE 22 SPAN STA. YITH..FUSELAfiEAND I TERBS. _. . . . SPANYISEDISTRIBUTION OF SECTION DRAG# LIFT# AND PITCHING nbnEHi ...... - Y SECTION SECTION . SLCTION . .- -" C C c sit- CHORO D L H
0.0000000 183.8817000 o..ooooooo ,2524992 -.2420649 .0250000 177.9445264 0.0000000 e2592031 -a2555230 .0500000 172.0073528 0.0000000 .2596443 -.2602768 e0750000 166.0701792 -e0996978 .2250000 -a2172664 .1000000 160.1330000 e0476099 .3000000 -.3057704 .1250000 154.1951859 ,1467936 .3750000 -e4036723 .1500000 348.2586941 a2434399 4500000 -.5122703 .1750000 142.3260032 .3415410 .5250000 "6330725 .2OOOOOO 136.3933123 a4407091 6000000 7672205 .2500000 124.6106675 e6381523 7500000 -1.0838952 3000000 113.9394744 a8306856 9000000 -1.4717630 350.0000 103.2682813 1.0074875 1.0500000 -1.9572210 .4000000 92.5970882 1.1627786 1.2000000 -2.5753208 .4750000 76.6960487 1.3200773 1.4250000 -3.8661103 .5500000 63.0912977 1.3938920 1.6500000 -5*7205812 .6250000 49.4865467 1.308 54 55 1.8750000 -8.6861825 .7000000 35.8817958 -5532453 2.1000000 -34.0324256 7500000 30.5922200 1.2209408 2.2500000 -18.0143604 .8000000 27.3627760 1.9458385 2.4000000 -21.8369286 .9000000 20.9038880 3.0278847 2 7000000 -33.2281995 .9500000 17r6744440 3.3840285 2.8500000 -42.1397923 1.0000000 14~4450000 2.1109021 3 0000000 -55.1972064 X CP C 1.020416 L S REF ---" -920569 -.147457 S PROG
.. INTERFERENCE DRAG OF LOADING 1 IUNIFORtl OR CONSTANT 1 ON LOADING LINEAR SPANYISE INTERFERENCE DRAG OF LOADING 2 ( LINEARCHOROYISE 1 ON LOADING LINEAR SPANYISE INTERFERENCE DRAG OF LOADING 4 ( QUADRATICSPANYISE 1 ON LOADING LINEAR SPANYISE INTERFERENCE DRAG OF LOADING 5 (QUADRATICCHOROYISE 1 ON LOADING LINEAR SPANYISE INTERFERENCE DRAG OF LOADING 6 1PARABOLICCHORDYISE ) ON LOADING LINEAR SPANYISE INTERFERENCE DRAG OF LOADING 7 ( CUBICCHORDYISE ) ON LOADING LINEAR SPANYISE INTERFERENCE DRAG OF LOADING 8 (SIHILAR TO FLATYING) ON LOADING LINEAR SPANYISE INTERFERENCE DRAG OF LOADING 9 (SO. ROOTFROH E.T. 1 ON LOADING LINEAR SPANYISE INTERFERENCE DRAG OF LOADING 10 ( ELLIPTICAL C-SUB-P ) ON LOADING LINEAR SPANYISE INTERFERENCE ORAG OF LOADING 11 (LINEARIN ARBeREGION) ON LOADING LINEAR SPANYISE INTERFERENCE DRAG OF LOADING 12 (BODYUPYASH LOADING 1 ON LOADIWG LINEAR .SPAMYISE. INTERFERENCE ORAG OF LOADING 13 ( NACELLEBUOYANCY 1 ON LOADING LINEAR SPANYISE INTERFERENCE DRAG OF LOADING 14 (NACELLEBUOY(CABBER1) ON LOADING LINEAR SPANYISE INTERFERENCE DRAG OF LOADING 15 (BODYUPYASH (CAHBER)) ON LOADING LINEAR SPANYISE.
L INTERFERENCE DRAG OF LOADING 16 ( BODY BUOYANCYTERH 1 ON LOADING LINEAR SPANYISE . P INTERFERENCE DRAG OF LOADING 17 (BODY BUOY. (CANBERJ 1 ON LOADING LINEAR SPANYISE c - ...... - .. - ...... - ..... ". . .J1ME .DATA FOR . QUADRATICSPANYISE . LOADING __. - ...... - ...... 969-599 . '1.7TOiQCHECK CASE 22 SPAN STA. YLTH FUSELAGEAND 2 TERMS, ...... _" _" - . __ ..__ - " " _.- ... CP - ...... C = . 1.033028. c- 1.386804 "- a829257 K 1.299244...... L. .. D L E. -e367456 C -e282861 ... M 0 ...... - .- ...... I,.NTEREERENC.E DRAG OF LOADING 1 (UNIFORM OR CONSTANT I ON LOADING ..4 i .PUADRATlCSPANYISE 1 IS . .62736474E+00 INTERF-ERENCEDRAG OF LOADING 2 ( LINEARCHORDYISE ONLOADING 4 i QUADRATICSPANYISE I IS ... .39317316Et00 1NTEREE.RENCEDRAG OF LOADING 3 I LINEARSPANYISE ONLOADING 4 ( PUADRAUCSPANYISE 1'1s . ,10849319EtOl - .IKL€RF_ERENCE DRAG OF .LOADING .5 (QUADRATICCHORDYISE 1 .ON LOADING __ .4. i QUADRAUC SPANYISE 1 .IS . .18173140Et00 . lHTLREERTNCEDRAG OF LOADING 6 (PARABOLICCHORDYISE I ONLOADING .. 4.1. QUADRATIC.SPANUSE ) IS .... .26964457Et00 INTEREERENCEDRAG OF LOAOING 7 ( cuuc CHORDVISE I ONLOADING 4 ( QUADRATICSPANYISE ) IS .76256456E-01 INTERFERENCEDRAG OF LO4OING 8 (SINILARTO FLAT YING) ON LOADING 4 ( .PUADRATLC SPANYISE IS . ..55597137Et00 I!ilX.RFER.ENCE DRAG OF- LOADING 9 (sa. ROOT FROM T. E. 1 ON LOADING.. 4 .( PUAORATlC SPANYISE ) IS .48665544EtOO .INTERFEPENCEPRhG OF LOADING 10 ( ELLIPTICAL c-sue+ I ONLOADING 4 L PUAIIRATIC SPANYISE I ,IS .- .53074380Et00 . - IHIER€U!EHCE_ DRALOF. LRADING 11 (LINEAR IN ARB.REGION) ON LOAOING. .4 .(. .PUADRhUC .SPANMILE I 1.5' . - -.*12634418E-01 ..INTERFERENCEDRAG OF .LOADING 12 (BODYYPYASH .LOADING ON LOADIHG. .. 4 L auruunc SPANUISE. I IS . .10336168E-01 .WlERF€RENCE ORAG OF LOADING 13 NACELLEBUOYANCY ON LOADING 4 .(. .QUADRALLC SPANYISE 1 .IS . .56882088E-O2 1NTER.FERENCE .QRAG .OF. LOADING 14 (NACELLE BUOY (.CAMBER) 1 ON LOADING 4 ( QUADRATIC SPANYISE 1 IS .56882088E-02 INTERFERENCEDRAG OF LOADING 15 (BODYUPYASH (CAMBER)) ON LMDING 4 ( PUAORATIC SPANYISE ) IS .10336168E-01 INTERFERENCEDRAG OF LOADING 16 ( BODYBUOYANCY TERM ON LOADING 4 QUADRATIC SPkNYXSE ) IS -.69055808€-02 JNTERFJREN4E .DRAG. OF LOADING 17 (BODY Buoy. (CAMBER) I ON. L.OADING 4 LQUADRATLC SPANYISE-l..IS - . .~..690558OBE-02 1 ...... _. . Y SECTION SECTION SECTION. . . -" C c c . .. __ . , .... " el2 CHORD D L n - 0.0000000 183.8817000 0.0000000 1.2783719 -1.4312949 .0250000 177.9445264 0.0000000 1.3542289 -1.5638501 .0500000 172.0073528 0.0000000 1.5393599 -1.8291181 00750000 166.0701792 20.5465001 2.0953535 -2.5471187 .1000000 160.1330000 8.7261141 1.9402178 -2.4720552 .1250000 154.1951859 5.4330349 1.0063386 -2.3957348 .1500000 148.2586941 3.6945433 1.6699737 -2.3184911 .1750000 142.3260032 2.3937813 1.5391427 -2.2403421 .2000000 136.3933123 1.8400386 1.4135907 -2.1609479 e2500000 124.6106675 -9155049 1.1800841 -2.0003908 *3000000 113.9394744 .4417276 -9861979 -1.8587434 .3500000 103.2682813 -1832124 .8103071 -1.7132505 .4000000 92.5970882 ,0494787 e6521124 -1.5629463 e4750000 76.6960487 -.0238774 a4471943 -1.3252851 5500000 63.0912977 -e0336361 3030265 -1.1260259 .6250000 49.4865467 -.0235780 -1867411 -.9117735 .7000000 35.8817958 -.0118632 .0985107 -.6831024 7500000 30.5922200 -.0037483 .0709289 -.5852814 .8000000 27.3627760 -.0006476 e0573165 -a5360979 .9000000 20.9038880 .0002949 -0331481 -.4160066 .9500000 17.6744440 .0003210 a0243550 -e3666617 1.0000000 14a4450000 .0000840 e0159019 -.2960677 " - . x- CP C = -917149 C 2.443731 "- = .771585 K = 2..905105 L 0 L E S C REF ,n ------920569 -- "219684 C = -a115619 5 C H PROS L 0 INTERFERENCE DRAG OF LOADING 1 (UNIFDRN OR CONSTANT ON LOADING 5 (QUADRATIC CHORDYISE ) IS INTERFERENCE DRAG OF LOADING 2 I LINEAR CHDRDYISE 1 ON LOADING 5 IPUADRATIC CHDRDYISE 1 IS INTERFERENCE DRAG OF LOADING 3 1 LINEAR SPANYISE 1 ON LOADING 5 IQUADRATIC CHDRDYISE 1 IS INTEEF€R€NCE DRAG OF LOADING 4 ( QUADRATIC SPANYISE ) ON LOADING 5[QUADRATIC CHORDWISE J -IS INTERFERENCE DRAG OF LOADING 6 (PARABOLIC CHORDYISE ) ON LOADIMG 5 (QUADRATIC CHORDWISE 1 IS INTERFERENCE DRAG OF LOADING 7 I CUBIC CHDRDYISE 1 ON LOADING 5 (QUADRATICCtlDRDYISE 1 IS INTERFERENCE DRAG OF LOADING 8 (SINILAR TO FLAT WING) ON LOADING 5 (QUADRATIC CHORDYISE I IS IITERFERENCE DRAG OF LOADING 9 (sa. ROOT FRON T. E. 1 ON LOADING 5 (QUADRATIC CHORDYISE 1 IS INTERFERENCE DRAG OF LOADING 10 ( ELLIPTICAL C-SUB-P 1 ON LOADING 5 (QUAORATICCtlORDYIS€ 1 IS INTERFERENCE DRAG OF LOADING 11 (LINEAR IN ARBeREGIONI ON LDAOING 5 LQUADBATIC CHORDWISE I 1s INTERFERENCE DRAG OF LOADING 12 [BODY UPYASH LOADING 1 ON LOADING 5iQUADRATIC CH0RDYISE.J IS INTERFERENCE DRAG OF LOADING 13 ( NACELLE BUOYANCY ) ON LOADING 5(QUADRATIC CHDRDYISE J:.IS INTERFERENCE DRAG OF LOADING 14 (NACELLE BUOY(CAHBER11 ON LOADING 5 (PUADRATIC UORDYISE I IS. ) II INTERFERENCE DRAG OF LOADING 15 (BODYUPYASH ICAhBER)) ON LOADING 5 [PUADRATIC CHORDYISE IS P INTERFERENCE DRAG OF LOADING 16 I BODY BUOYANCYTERH 1 ON LOADING 5 (PUADRATIC..CHORDYISE .1 .IS w INTERFERENCE DRAG OF LOADING 17 (BDDY BUOY. (CAHBERI ) ON LOADING 5 (QUADRATIC CHOROYISE 1 IS ...... KINGDATA..FaR PARABOLIC CHDRDYLSL LOAOING . -. . .- --- .- - .. 969-3QC. 17 ~LOAD.I.HEtK. CASE. 22 SPAN .STAr-. YITH:.FYS€LIG€.b4D2-TEEH.X -- -- - . .... - ... .. - ._. - spAyY.I$.E 01SJRIBUT.IONSECTIbNnF DRAG,.'LIF.T;.-~D.PITCHING. hOMENT. .... '1. .... Y SECTION.SECTION ... SECTIDN "- C c .. c . . _" 812 .. CHORD 0 L .n ...... - ..... - ... 1...... 0~0000000 183.8817000 -1.39185890~0000000 1.4917912 ...... 0250000 177.9445264 0~0000000 1.5615864 -1.4985080 " .. . qO500000 172.0073528 0~0000000 1.6968809 . -1.6570225 .~ - .- - -. .0750000 . 166.0701792 . 20.1820783 2.0958671 . -2.0240129- .1.000000160.1330000 7.83611321.9519239 -1.9876386 . __ - ...... lt50000 154.1951859 4.3180221 1.8077514 . .-1*9456747.. - . . . s1500000 148.25869412.5981171 1.6700038-1.9015728 .1750000152.3260032 1.4855479 1.5412931-1.8573958 . r2OOOOOO 136.3933123 -7664999 1.4136547-1.8075831 .2500000124.6106675 -e0557157 1.1807185-1.7056169 .3000000 113.9394744-e2671609 ,9890139-1.6160763 ._ .. .3.frOOOOO 103.2682813-e3572653 e8097533. . -3.5097411 .COO0000 92.5970882-1,3961343e6506543"3646834 .47500 00 76.6960487.4750000-e2814541-1.2107498e4461790 .550000 0 63.0912977.5500000"1735138 .3047960-1.0547725 .625000 0 49.4865467.6250000 -.0971234 -e8657318a1871792 .70000 00 35.8817958.7000000 -.0440600 -.6538292e0979868 ... .7!jOOOOg 30.5922200-.0224898e0727414- .. -.5817040 .8000000 27.3627760-.5126662-mol19420e0563681 .90000 00 .20.9038880.9000000 -.0035512 .0340047 . -e4181634 ...... - ...... 95OOOOO 17.674444Q-.0015462 .0226862 . -e3352519 . ". . 1~0000000~ 14.4450000-.0008254 . .017619C-.3235225 . - ...... X .-. - . " - ...... CP ...... - - ...... C 9 e917731 C le898925 "- = e649283 K = 2.254638 .. 1 D L E .s . C . ." REF .. n ...... -"" . .920569 -- * -093692 C e171901 ...... L C n ..... PROG L 0 ....- ...... 1.Nl.ERFERENCE DRAG OF LOADING 1 (UNIFORM OR CONSTANT 1 ON LOADING b lPARA8oLIC CHWYLSE 1- LS-. LNTERFERENCE DRAG OF LOADING 2 I LINEARCHOROYISE ON LOADING b IPARAEDLICCHORRYISE 1 IS ... INTERFERENCE DRAG OF LOADING 3 ( LINEARSPANYISE 1 ON LOADING 6 (PARABOLIC CHORDYLSE 1. IS . . INTERFERENCE DRAG OF LOADING 4 ( QUADRATICSPANYISE 1 ON LOADING .AL (PARLBOLLC.CHOR0YISE .)-.IS - INTERFERENCE DRAG OF LOADING 5 IQUADRATICCHORDYISE ) ON LOADING .. 6 LPARABOLIL CHORDWISE .). IS INTERFERENCE DRAG OF LOADING 7 I CUBICCHORDYISE ) ON LOADING 6 (PARABOLICCHORRYISE 1 IS INTERFERENCE DRAG OF LOADING 8 (SIflILAR TO FLATWING) ON LOADING 6 IPARABMICLHORRYISE ).IS 1.NTERFERENCE DRAG OF LOADING 9 (SO. ROOT FROfl E.T. I ON LOADING 6 (PARABLILIfCHORDYISE I IS .. INTERFERENCE DRAG OF LOADING 10 I ELLIPTICAL C-SUB-P ) ON LOADING. 6 ~PARAB0UC.CHORDYISE. ).IS.- .... I.ITERFERENCE..DRAG OF LOADING 11 (LINEAR IN ARBaREGION)RH -LOADING_. 6 .(PARABU. CHOURYlSL .LIS-...- IYTERFERENCE DRAG OF LOADING 12 (BODY UPYASHLOADING ) ON LOADING 6. LPARABOLILCHOBDYIS€. ) IS. - .I.NTERFERENCE DRAG OF LOADING 13 I NACELLE BUOYANCY ) ON LOADING 6. LPARABMUCHORDUSE-J-L- . XNTERFERENCE DRAG OF LOAOING 14 (NACELLE BUOY (ChMBER)) ON LOADLKG. b (.PARABU-CHQtUWLIS" - INTERFERENCE DRAG OF LOADING 15 (BODYUPYASH (CAMBER)) ON LOADING . h ~PARAUtLIC_CHORI1ULSE..J1s" . 1.N.TERFERENCE DRAG OF LOADING 16 ( BODY BUOYANCY TERfl 1 ON LOADING 6_LPARABOUC~LIIIRIIY~€.J.15... - .INTERFERENCE DRAG OF LOAOING 17 (BODY BUOY.. (CAMBER). 1 ON LMDIPG-.4..LEAaaenLrcRERdLULl-LS" ...... - - - . . Y SECTION SECTION SECTIOLL . e" C ... c c ...... BIZ CHORD 0 L n - . , ...... _. . 0,0000000 183.8817000 0.0000000 1.6742763 -1.BlO4941 .0250000 177.9445264 0.0000000 1.7663932 -1.9714472 .0500000 172.0073528 0.0000000 1.9896429 -2.2781662. - . .0750000 166.0701792 35.3653518 2.6814571 -3.1256052 .1000000 160.1330000 12.7152031 2.4041490 -2.9304524 . .1250000 154.1951859 6.5658502 2.1464063 -2.7395879 - .1500000 148.2586941 3.5714341 1.9079036 -2.5535465 e1750000 142.3260032 1.8842747 1.6879747 -2.3726189 .2000000 136.3933123 .8868117 1.4855144 -2.1966251 .2500000. lZ't.6106b75 -.Ob63247 1.1328514 -1.8636165 . .3000000 113.9394744 -a2547880 e8660750 -1.5890976 .3500000 103.2682813 -.t77258a .64'17524 -1 3309269 . .4000000 92.5970882 -.2212113 .4648162 -1.0906226 . .4750000 76.6960487 -.1185498 e2641315 -.76951,72 .550000O 63.0912977 -.0514031 .1471426 -.5393039 .6250000 49.4665467 -.0185476 .0710040 -e3430296 .7000000 35.8817958 -.0049046 .0270615 -.1862225 .7500000 30.5922200 -.001892? a0167955 -.1377510 '~.8000000 27.3627760 -.0008447 e0119949 -.1115649 .9000000 20.9038880 -.0001467 .0053626 -.Ob70419 .9500000 17.6744440 -.0000544 e0032269 -.0463692 1.0000000 14~4450000 -~0000161 .0016156 -a0337047 ... .- A CP C .9$1350 C. L D S C REF M " = ...... -"- - m...... -920569 "135793 C 9 -e039700 s C M PROG L 0 ...... "-. . - - . . - ." " . . INTERFERENCE DRAG OF LOADING 1 LUNIFORM OR CONSTANT I ONLOADING CUBIC CHOROYLSL INTERFERENCE DRAG OFLOAOING 2 LINEARCHOROYISE ON LOADING CUBICCHOROYISE INTERFERENCE DRAG OFLOADING 3 ( LINEARSPANYISE ) ONLOADING CUBIC .CHOROYISE. INTERFERENCE DRAG OFLOAOING 4 ( PUADRATICSPANYISE ) 0N.LOADING .CUBIC CHORDVISE INTERFERENCE DRAG OF LOADING 5 (OUADRATICCHOROYISE J ON LOADING CUBIC. CHORDYISE INTERFERENCE DRAG OF LOADING 6 (PARABOLICCHORDYISE 1 ONLOADING CUBIC .cnnRoYIsE- INTERFERENCE DRAG OFLOADING 8 (SIMILARTO FLAT WING) ON LOADING CUBIC .CHOROYISE INTERFERENCE DRAG OFLOADING 9 (Sa.ROOT FROM T. E. I ON LOADING CUBICCHOROYISE INTERFERENCE DRAG OF LOADING .O ( ELLIPTICALC-SUB-P 1 ON LOADING . CUBIC CHORMIS INTER€EPENCE DRAG OFLOADING 11 (LINEARIN ARB.REGION1.ONLOADING CUBIC LHORLULX- 1.NTER.FERENCE .DRAG OFLOADING .2 (BODYUPYASH LOADING LON LOADING .CUBIC. MORDULSE. INIERFERENCE DRAG OFLOADING .3 NACELLEBUOYANCY 1 ON LOAOING CUBIt. CHORQYISL.. INTERFERENCE DRAG OFLOADING .4 (NACELLE6UOYlCAH6ERIJ ONLOADING. CUBICCHORDYLSL. INTERFERENCE DRAG OFLOADING .5 (BODYUPYASH (CAIEERII ON LOADING CUBIC cHmauIsE. INTERFERENCE DRAG OFLOADING .6 I BODY BUOYANCYTERN 1 ONLOADING CUBIC..CHORDYIS€. 1tl.TERFEREN.CE DRAG OFLOADING .7 (BODYBUOY. (CANEERI J ON. LOADING CUBLC ..CtLOBMISE.- Y SECTION SECTION -" C L 8/2 CHORD D L c.. 0000000 183.8817000 0.0000000 .7511525 -.6347600 .0250000 177.9445264 0.0000000 .7866845 -.6827532 . .0500000 172.0073528 0~0000000 ,8458607 -.7473634 .0750000 166.0701792 3.7361336 1.0000451 -.E656535 .1000000 160.13300001.7742890 .9961879 -.917558D .1250000 154.19518591.3331175 .9988378 -.9758993 . . .1500000 148.25869411.0168e03 1.0007716 -1.0388276 .1750000 142.3260032 -8616729 ,9968933 -1.1039064 .2000000 136.3933123 ,7536686 ,9997471 -1.1785240 .2500000 .5591254124.6106675 .99?8125 -1.3432789 3000000 113.9394744 .3860239 .9948416 -1.5298483 .3500000 103.2682813 -2351520 1.0021850 -1.7667940 ,4000000 92.5970882 .181e981 1.0008105 -2.0472644 .4750000 76.6960487-.0151047 1.0033199 -2.6200887 .. . . .5500000 63.0912977 -.lo60708 .9835553 -3.3191919 .6250000 49.4865467-e1596947 .9835181 -4.4653032 ,7000000 35.8817958 -.3177683 .9839211 -6.4831123 .7500000 30.5922200 -* 1770755 ,9710085 -7.6907130 . .8000000 27.3627760 .Ob09439 .9978718 -8.9798354 . .9000000 20.9038880 ,2549810 ,9677070 -11.8278473 .9500000 17.6744440 ,2206933 1.0165444 -14.9184371 1.0000000 14.4450000-e1545978 .9109414 -16.7074209 . 1 X CP .. . c= -933996 C = ,600950 --- 8 i675634 K= L 0 L E S C REF M "-" e920569 -- = .026172 C = e111885 S C M PROG L 0 . INTERFERENICE DRAG 0 F LOADING 1 (UNIFORM OR CONST ANT 1 ON LOADI NG ILAR TO FLAT YIN I GJ IS e66895198EtOO INTERFERENCE DRAG OF LOADING 2 ( LINEARCHORDWISE ) ONLOADING 8 (SIMILAR TO FLAT YING I IS ~10010980E+01 INTERFERENCE DRAG OF LOADING 3 ( LINEARSPANYISE 1 ON LOADING 8 (SIMILAR TO FLAT YING) IS .29189613Et001 INTERFERENCE DRAG OFLOADING 4 ( PUAORATICSPANYISE 1 ON LOADING 8 (SIMILAR TO FLAT YING 1 .IS ~10375287Et00 INTERFERENCE DRAG OFLOADING 5 (QUADRATICCHORDYISE 1 ON LOADING 8 (SIMILAR TO FLAT YING) IS .12407691E+01 iNTERFERENCE DRAG OFLOADING 6 (PARABOLICCHORDWISE 1 ON LOADING 8 (SIHILAR TO FLAT YING) IS .l2016898Et01 iNTERFERENCE DRAG OFLOADING 7 ( CUBICCHORDWISE 1 ON LOADING 8 (SIMILAR TO FLAT YING) IS .14425E19€+01. INTERFERENCE DRAG OFLDADING 9 (Sa. ROOTFROM 1. E. I ONLOADING 8 (SIMILAR TO FLAT YINGl IS .02715982E+00 INTERFERENCE DRAG OF LOADING 10 ( ELLIPTICALC-SUB-P 1 ONLOADING 8 (SIMILAR TO" FLAT WING) IS .72627049Et00.2. INTERFERENCE DRAG OFLOADING 11 (LINEARIN ARB-REGION) ONLOADING B.(SIMILAR TO FLAT Y1NG 1 IS- ..41371082E-41.. INTERFERENCEDRAG OF LOADING 12 (BODYUPYASH LOADING 1 ON LOADING 8 (SIMILAR TO-. FLAT YING 1 IS .lzn56143~-01. INTERFERENCE DRAG OFLOADING 13 ( NACELLEBUOYANCY 1 ONLOADING 8 (SIMILAR TO FLAT UING) IS .2945?836E-02 INTERFERENCEDRAG OF LOADING 14 (NACELLEEUOY(CAfl8ER)) ON LOADING 8 (SINILAR. TO FLAT YING) IS. ..29457836E-D2. INTERFERENCEDRAG OF LOADING 15 (BODYUPWASH (CAMBER)) ON LOADING 8 (SIMILAR TO FLAT YING) IS .12856143E-01 ZNTERFERENCEDRAG OF LOADING 16 I BODY BUOYANCYTERM ) ONLOADING 8 (SIMILAR TO FLAT.YING) IS -.16301299€-02 1NTERFERENCEDRAG OF LOADING 17 (BODYBUOY. (CAMBER) I OH .LOADING 8 (SINLLAR To FLAT.. YINGJ. .IS - -Jb?iO.UPPE-02.. . ~...... YING DATAFOR SQ1 ROOT FROM T. EA LOADING 969-500 17 LOAD CHECK CASE 22 SPANSTA. YITH FUSELAGEAN0 2 TERHS .. . . SPANYISEDISTRIBUTION OF SECTION DRAG, LIFT# AND PITCHING MOHENT . ." ". Y SECT.ION SECTION SECTION "- C C C BIZ CHORD 0 L n . ". 0.0000000 183.8817000 0.0000000 1.0131671 -.87139r7 .0250000 177.9449264 0.0000000 1.0608846 -.937Qb73 . .0500000 172.0073528 0.0000000 1.1433267 -1.0265615. .0750000 166.0701792 7.2748445 1.3630046 -1.1985779 .1000000 160~1330000 3.3366139 1,3381735 -1.2487639 .. . ~1250000 154.1951e59 2.2714838 1.3133294 -1.3005662 .1500000 148.2586941 1.6893780 1.2878869 -1,3549958 .1750000 142.3260032 1.3009306 1,2615795 -1.4126423 .2000000 136.3933123 1.0139926 1.2351348 -1.4727876 .2500000 124.6106675 e5768761 1.1803914 -1.6040857 3000000 113.9394744 -3325216 1.12877r2 -1.7488981 .3500000 103.2682813 -1328114 1.0748411 -1,9105994 .4000000 92.5970882 -no174514 1.0172594 -2.0948971 .4750000 76.6960487 -.2040746 .9261461 -2,4324628 .5500000 63.0912977 -a2685511 8385092 -2.8352564 e 6250000 49.4865467 -a3159346 .7420212 -3.3734019 .7OOOOOO 35.8017958 -e3682630 ,6307752 -4.1594252 ~7500000 30.5922200 -a2153391 a5844066 -4.6303483 .8000000 27.3b27760 "1155193 ,5517096 -4.9707846 9000000 20.9038880 -.0216555 ,4831362 -5.9061697 I .9500000 17.6744440 -e0062481 .4402294 -6.4658459 1.ooooooo 14.4450000 -.0315881 ,3973887 -7.2827496 X CP C * .986059 C9 .955578 L. 0 S C REF M -----". 9 e920569 * .084568 C * m175704 .s C M PROF L 0 INTERFERENCE DRAG OF LOADING 1 (UNIFORH OR CONSTANT 1 ON LOADING 9 (SO. ROOT FROM T. E. ) IS ~71557952Et00 INTERFERENCE DRAG OF LOADING 2 I LINEAR CHORDUISE 1 ON LOADING 9 ISP. ROOT FROM T. E. ) IS .12438662E+01 , INTERFERENCE. .~ .- - " DRAG. OF-. LDADING" . 3 ( .LINEAR SPANYISE ) ON LOADING 9 [sa. ROOT FROM T. E. I IS .13596999E+OO INTERFERENCE DRAG OF LOADING 4 ( QUADRATIC SPANYISE 1 ON LOADING 9 (Sa. .ROOT FR.0M.T.a E. 1 .IS. . -.12756533€+00 INTERFERENCE DRAG OF LOADING 5 (QUADRATIC CHORDYISE I ON LOADING 9 (SO. ROOT FROM T. E. I IS .lbC18523E+Ol INTERFERENCE DRAG OF LOADING 6 (PARABOLIC CHORDWISE 1 ON LOADING 9 (sa. ROOT FROM T. E. I IS .15489058EtOl INTERFERENCE DRAG OF LOADING 7 ( CUBIC CHORDYISE ) ON LOADING 9 (sa. ROOT FROH T.. E. 4 IS .19773895EtOl INTERFERENCE DRAG OF LOADING 8 (SIHILAR TO FLATYING) ON LOADING 9 (Sa. ROOT FROM 1. E. I IS r61885110Et00 INTERFERENCE DRAG OF LOADING 10 ( ELLIPTICAL C-SUB-P 1 ON LOADING 9 (sa. ROOT F.ROM.T.. .E. .I IS . ..802435Z7Et00 I-. .._ __._INTERFERENCE DRAG .OF LOADING 11 (LINEAR IN ARBaREGION) ON LOADING 9 (SQ..ROOT-EMM.T..La 1. IS. ..,49048708E~Ol . INTERFERENCE DRAG OF LOADING 12 (BOOY UPYASH LOADING 1 ON LOADING 9 (sa. ROOT FROK I*.€. ) IS . .14383!07E-01. INTERFERENCE DRAG OF LOADING 13 ( NACELLE BUOYANCY ) ON LOADIblG 9 ISP. ROOT EROM TL.€. .)..IS. . .20855132-E-02 INTERFERENCE DRAG OF LOADING 14 (NACELLE BUOY(CAMBER)) ON LOADING 9 ISP. ROOT.FRDM 1m.E. I .IS .ZO~S~~E-OZ INTERFERENCE DRAG OF LOADING 15 (BODY UPYASH (CAMBER)) ON LOADING 9 (sa. ROOT. FROH T.. E- IS .. .14383407E-01 INTERFERENCE DRAG OF LOADING 16 ( BODY BUOYANCY TERM ) ON LOADING 9 (SP. ROOT FRM Ta.E. .I .IS . -.84264670E-03 INTERFERENCE DRAG OF LOADING 17 (BOOY BUOY.(CAMBER) ) ON LOADING 9 (sa. RnoT-EBnn LE. .IS ..1.0dZb467OL-03 ...... - YINGDATA FOR ELLIPTICAL C-SUB-P LOADING .. .. 9C?-500 17 LOACCHECS CASE 22 SPANSTA. UITH FUSELAGE AN0 2 TERMS . . -. SPA!IdISfDISTkIEUfION OF SECTION DRAG, LIFT,AND PITCHLNG HOHENT Y SECTION SECTION SECTION -" C C C P, IC' CHORD 0 L M L - oc;'Looo lJ3.8817000 0.0000000 .8080698 -a7459070 aC25G000 177.9445264 0.0000000 ,8464148 -.8036624 .0500005 172.Ct73528 0.0000000 ,9161870 -.8887469 .07:COOC 166.0701792 4.0274333 1.0969019 -1.0591997 .lOtOOOO 160.1330000 2.0130837 1.0944862 -1.1155383 .1250000 154.1951859 1.4895591 1.0913724 -1.1748181 1150000C1 146.2586941 1.2109220 1.0875546 -1,2380487 . .175OOOG 14i.3260032 1.0225564 1.0830253 -1.3059686 .2000000 136.3933123 ,8809192 1.0777755 -1.3781524 ,2500000 124.6106675 -6642529 1.0650704 -1.5392330 .3000000 113.9394744 ,5013291 1.0493331 -1.7159663 .350000C 103.2682813 ,3671192 1.0304247 -1.9207322 .4OCOGOO 92.5970882 -2523484 1.0081667 -2.1636271 .4750000 76.6960487 .lo12056 e9679844 -2 6261996 . .5500000 63,0912977 -.0200900 .9186811 -3.1850848 .6250000 59.4865467 -.1188605 ,8586872 -3.9779808 .7000000 35.8817958 -.2495885 .7855571 -5.2491771 .75COOOO 30.5922200 "1327435 .7275816 -5.8252966 . .EO00000 27.3627760 -.OS59525 .6600000 -6.0051554 .9000000 20.9038880 -.0171421 .4794789 -5.9008222 .9500000 17.6744440 -.0202062 .3434749 -5.0785826 1.0300000 14.4450000 0.0000000 0.0000000 0.0000000 X "" CP_. . . .. C 9 .919594 C 9 .704516 --- .698575 K = .a33103 L 0 L E C n 9 -e032610 C * -056104 C U PROG L 0 .. ItiTERFERENCECRAG OFLOADING 1 (UNIFORfl OR CONSTANT 1 ON LOADING 10 I ELLIPTICALC-SUB-P I IS . .64299531EtOO INTERFERENC'ORAG OFLOADING 2 I LINEARCHOROUISE 1 ON LOADING 10 I ELLIPTICALC-SUB-P I IS .10658583Et01 rNTERFERENCEDRAG OFLOADING 3 1 LINEARSPANYISE I ON LOADING 10 I ELLIPTICALC-SUE-P 1 IS .25876723EtD0 it!TEF.FER€NCECRAG OF LOADING 4 1 PUADRATICSPANYISE I ON LOADING -10 I ELLIPIICAL C-SUE-P. 1 IS .5097a3~4~-01 1;:TERFEREKCE DRAG OF LOADING 5 PUACRATICCHORDWISE I ON LOADING 10 ( ELLIPTICAL C-SUB-P 1 IS .14114779E+01 INTERFERENCE DRAGOF LOADING ..6 PAAAaOLICCHOROWISE 1 ON LOADING 10 I ELLIPTICAL C-SUE-P 1 IS .llb83273Et01 . !NTECFERENCE ORACOF LOADING I CUBICCHOROYISE 1 ONLOADING 10 I ELLIPTICALC-SUB-P I IS .15949953Et01 :NTEkFEKEECE CRAG OF LOADING 8 SIMILAR TO FLATYING 1 ON LOADING 10 I ELLIPTICALC-SUE-P 1 IS ,53394344EtOO !NTEEFER5NCE CRAGOF LOADING 9 sa. ROOT FROM T. E. 1 ON LOADING 10 I ELLIPTICALC-SUB-P 1 IS .74733741E+0OF-.. I%TEEFERENCE CRAGOF LOADING 11 LINEAR IN CPB.REGION I ON LOADING 10 ELLIPT1CAL.C-SUE-P 1 IS .51419306E-01 INTERFERENCE lEAG OF LOADING 12 BODYUPYASH LOADING 1 ON LOADING 10 I ELLIPTICALC-SUE-P 1 IS .121076E2E-01 IPITEEFERENCE CRAG OF LOADING 13 NACELLEBUOYANCY I ON LOADING 10 I ELLIP.TICAL C-SUE-P I IS .37141478E-02 IXTEEFLSENCE CRAGOF LOCOING 14 FiACELLEBUOYICAMaER) I ON LOADING 10 ( ELLIPTICAL C-SUE-P I IS .37141478E-02 ISTERPEPENCE DRAGOF LOADIidG 15 BODYUPYASH ICAMBERI 1 ON LOADING 10 I ELLIPTICAL C-SUB-P I IS .321076E2E-01 1!4TEl!FEREN!:E: L'EAGOF LOADING It 50OY BUOYANCYTERM I ON LOADING 10 ( ELLIPTICAL C-SUB-P I IS -.60321758E-03. . INTFEFiRENCE ORAG ilF LOADING 17 BODYBUOY. ICAUEERI I OK LOADING 10 I ELUP.TLCAL C-SUE-P I IS -.6932174BE-03 " ...... ". ... bbn-500~ii LOAD CHECKCASE zz SPAN STA. .mi PUSELAGEAND z Tuns ...... - . -. . -..-. SPANUISE DISTRIBUTION OF SECTION DRAG, LIFT, AND PITCHING. MHENT -. .... Y SECTION . SECTION .. SECTION -. . -" C C c ." . BIZ .. CHDRD D L: .I I .... . O~OOOQOOO 183*8817000 0~0000000 -.0240441.0187135 . .Qf50000 177.9445264 0~0000000 .OZZZ419 -e0293964 .0500000 172.0073528 0~0000000 a0330258 -.0448299 ~0730000 166.0701792e0297939.Ob16348 -.0862040 ,1000000 160~1330000 .0177053 a0592769 -.0860747 .1250000 154.1951859 e0144896 -0569657 . -e0859933 .1500000 148.2586941.0125877 -0548954 -.0862554 .- .1750000 142.3260032.0112873 .0527201 -a0863849 .ZOO0000 136.3933123 .0106112 . .0504032-.C862948 .2500000 12406106675 .0091111 e0461476 -.0866876 . ,3000000 113,9394744 ,0071755 .0417704 -e0863995 - .. a3300000 103.2682813 e0057944 -0379834 -e0873057 .4000000 92.5970882 e0044758 -0344417 -no888993 .4750000 ' 76.6960487 e0033489 e0283823 -a0893197 ..... ". .550000063.0912977 .0022243 -0233725 -.0911995 .625000049,4865467 e0014684 .0186231-.0944147 .7000000 35.8817958 .0008995 .0140130 -.0998602 .750000030.5922200 .0004769 . .0105987 -a0894721 1 .BOO0000 27.3627760.0004990 e0100896 -.0963497 .. .9000000 20.9038880.0002184 e0080284 -.1021017 e9500000 17.6144440 .0003026 .0077910 -a1190176 1~0000000 lCe4450000 .0000724 .0031771 -e0599916 .. - - X _. . . .- CP .... C = ,036338 C = .007826 --- 9 e867120 K 5,926973 .. L D L E " . S C REF M " = -e464472 C 8 -.013476 ---- .920569 -- " . C M PROG L 0 -" .. INTERFERENCE DR~GOF LOADING 1 IUNIFORn OR CONSTANT 1 ON LOADING L1 (LINEAR IN ARB.REGION) IS .23637590€-01 INTERFERENCE DRAG OF LOADING 2 I LINEARCHORDUISE 1 ON LOAOING 11 (LINEAR IN ARB.REGION1 IS e52556546E-01 INTERFERENCE DRAG OF LOAOING 3 ( LINEARSPANYISE I ON LOAOING 11 (LINEAR IN ARBmREGLOtU IS .17.462539E-Olr 1,NTERFERENCE DRAG OF LOADING 4 I QUADRATICSPANUISE ON LOADING 11 (LINEAR IN ARB.REGION1 IS .12610489€-01 INTERFERENCE DRAG OF LOADING 5 (QUADRATICCHUROYISE ) ON LOADING 11 (LINEAR IN ARB.REGION1 IS .83848568E-01 INTERFERENCE DRAG OF LOADING 6 (PARABOLICCHOROYISE 1 ON LOADING 11 (LINEAR IN ARB.REGI0NI IS . . .10234734E-01 INTERFERENCE DRAG OF LOADING 7 I CUBICCHOROUISE ) ON LOADING 11 (LINEAR IN ARB.REGION1 IS .68675837E-01 INTERFERENCE DRAG OF LOADING 8 (SIMILAR'TOFLAT YING) ON LOADING 11 (LINEAR IN ARBmREGION) IS .15157762E-01 . INTERFERENCE DRAG OF LOADING 9 (Sa. ROOT FROM 1. E. 1 ON LOiOING 11 (LINEAR IN ARBmREGIONJ IS .12692167E-01 - LNTERFERENCE DRAG OF LOADING 10 I ELLIPTICAL C-SUB-P 1 ON LOADING 11 (LINEAR IN ARB.R€GIONJ-IS .. .24741670E-01 INTERFERENCE DRAG OF LOADING 12 (BODY UPYASHLOADING 1 ON LOADING 11 (LINEAR IN ARBeREGION) IS a29265031E-03 ItiTERFERENCE DRAG OF LOADING 13 ( NACELLE BUOYANCY ON LOADING 11 (LINEAR IN A!?BcREGIONI IS ..522143ClE-03 INJERFERENCE DRAG OF LOADING 14 (NACELLEBUOY(CAH8ER)I ON LOAOING 11 (LINEAR IN ARB.REGION1 IS .52214341€-03 INTERFERENCE DRAG OF LOADING 15 (BODYUPUASH ICAHBER)) ON LOADING 11 (LINEAR IN ARBeREGION) IS .29265031E-03 INTERFERENCE DRAG OF LOADING 16 I BODY BUOYANCYTERH ON LOADING (LINEAR IN ARBaREGION) P 11 IS e33628352E-03 \o !NTERFERENCE DRAG OF LOADING 17 l80DY BUOY. ICAtiBER) 1 ON LOADING 11 (LINEAR IN ARB.REGION) IS . .mza352~-03 . .. ,. - .. . "- .-. - VI 0 YINGDATA FOR BODY UPYASH LOADING LOADING -. .." .~. . . . p69.-50o 17. COAD CHECK CASE 22 SPAN STA. YITH FUSELAGE AND .z TERMS.. .. _. . . SPANVISE DISTRIBUTION OF SECTION DRAG, LIFT, AND PITCHING MOMENT. Y SECTION SECTION SECTION "- C C C B/2 CHORD .D 1 M 0.0D00000 183.8817000 0.0000000 .0199352 -.0171715 .0250000 177.9445264 0.0000000 -0203094 -.0183712 .0500000 172.0073528 0.0000000 e0217256 -a0203907 .0750000 166.0701792 0.0000000 -0222595 -.0212843 .1000000 160.1330000 0.0000000 .0221333 -e0212475 .. . . .1250000 154.1951859 0.0000000 e0216492 -e0216256 .1500000 146.2586941 0.0000000 -0211931 -.0222990 .1750000 142.3260032 0.0000000 -0203227 -e0228787 ...2000000 136.3933123 0.0000000 .0200290 -.0239821 .. .250ODOO 124.6106675 0.0000000 ,0191149 -.0262668 ...3000000 113.9394744 0.0000000 .O182813 -.0287738 . ". . .3500000 103.2682813 0.0000000 .0182212 -.0326650 .4000000 92.5970882 0.0000000 .0177127 "0366937 .4750000 76.6960487 0.0000000 -0174484 -.0458375 .5500000 63.0912977 0.0000000 e0171522 -.0580714 .6250000 59.4865467 0.0000000 e0182157 -.0829274 .7000000 35.8617958 0.0000000 .0209287 -.1379502 . . . - . . .- 7500000. 30.5922200 0.0000000 .0218173 . -e1728406 .8000000 . 27.3627760 0.0000000 .0220862 -a1994720 .9000000 20.9036880 0.0000000 .0201745 -.2461318 .9500000. 17.6744440 0.0000000 .0173637 -.2569930 - . - .. . . 1,0000000 14.4450000 0.0000000 e0112932 -.2076505 X CP C 9 .016252 c9 0.000000 --- 9 684021 K 9 0.000000 .. .. - L D L E S C . . - REF M ----- 8 ,920569 " 8 e004682 C 8 e001794 S C M -. . PROG L 0 .~ INTERFERENCE DRAGOF L OAOING 1 (UNIFORM OR CONSTANT 1 ONLOADING 12 (BODY UPYASH LOAOING 1 Is 0. INTERFERENCE DRAGOF 1OADING 2 ( LINEARCHORDWISE ONLOADING 12 (BODY UPWASH LOADING 1 IS 0. INTERFERENCE DRAGOF L OADING 3 ( LINEARSPANYISE ONLOADING 12 (BODY UPYASH LOAOING 1 IS 0. INTERFERENCE DRAGOF L OADING 4 ( PUADRATICSPANYISE 1 ONLOADING 12 IBOOY UPYASH LLIAOING. I. IS. 0. INTERFERENCE DRAGOF L OADI NG 5 (PUADRATICCHORDUISE 1 ON LOADING 12 (BODY UPYASH LOADING 1 IS. 0. INTERFERENCE DRAGOF L OADING 6 (PARABOLICCHORDYISE ONLOADING 12 (BODY UPYASH LOADING I IS 0. INTERFERENCE DRAGOF L OADING 7 1 CUBICCHORDWISE 1 ONLOADING 12 (BODY UPYASH LOADING I IS. 0. INTERFERENCE DRAGOF L DADING E (SIMILAR TO FLATYING) ONLOADING U (BODY UPYASH LOADING I IS 0. . . 1.NTERFERENCE DRAG OF L OADING 9 ISP. ROOTFROM T. E. I ONLOADING 12 (BODY UPHASH LOADING I 11 a -- -.. .INTERFERENCE DRAGOF L OADING 10 I ELLIPTICALC-SUB-P ) ONLOADING 12 l8DDY UPHASH LOADlNG I- LT Q. INTERFERENCE DRAGOF L OADING 11 (LINEARIN ARB.REGION1 ONLOADING 12 (6OD.Y UPYASH LMO.ING. I. IS 0. INTERFERENCE DRAG OF L OADING 13 ( NACELLEBUOYANCY 1 ONLOADING 12 (6OD.Y UPYAW LOADING Is. 0. lNTERFERENCE DRAG OFLOADING 14 (NACELLEBUOY(CAMBER)) ONLOADING 12 (BODY UPYASH LOAQING 1 IS 0. INTERFERENCE DRAG OFLOADING 15 (BODYUPYASH (CAMBER)) ON LOADING 12 (BODY UPYASH LOADING .J IS 0. INTERFERENCE DRAG OFLOAOING 16 I BODYBUOYANCY TERM 1 ONLOADING 12 (BODY UPYASH LOARING 1 IS 0. INTERFERENCE DRAG OF LOADING 17 (BODY BUOY. (CAMBER) ON LOADING 12 (BODY YPYASH LOADING 1 IS - -0...... -. "- .. . YINGDATAFOR NACELLE BUOYANCY . LOADING .. .. " -.- .. . - .. - .. ,969-500 17 LOAO CHECK CASE 22 SPAN STA. YITH"FUSELAGE.ANO L IEMS .- .. _. SPAHWISE.RISTRIBUTION OF SECTIONORAGa LIFT0 .AN0 .PITCHING HMWT SECTION SECTION. .SECTION C C C CHORD 0 L n .. . . .. 0.0000000 183.8817000 0.0000000 .0001173 . . . -.0001540" . .0250000 177.9445264 0.0000000 .0004694 . -.0006370 .0500000 172r0073528 0.0000000 e0009759 . -moo13649 .0750000 166.0701792 0.0000000 a0017408 -e0024903 .1000000 160.1330000 0.0000000 .0020751 -e0030665 .. .l250000 . 154.195~59 0.0000000 .0029497. "0044669 .1500000. 148.2586941 0.0000000 .0032910 -.0051611 . .1750000 142.3260032 0.0000000 .0030770 -e0062537 .2000000 136.3933123 0.0000000 ,0045904 -.0016394 . .2500000 124~6106675 0.0000000 .0054555 -e0098394 .3000000 113.9394744 0.0000000 .0072213 -a0145887 .3500000 .103.2682813 0.0000000 .0080303 -e0200926 .4000000 92.5970882 0.0000000 e0105369 -e0267138 4750000 76.6960487 0.0000000 ,0093157 -.0277060 .5500000 . 63.0912977 0.0000000 .0096136 -a0355341 .6250000 59.4865461 0.0000000 -0092359 -e0454217 .7000000 35.8817958 0.0000000 ,0081753 "0574132 .7500000 30.5922200 0.0000000 e0043955 -a0370185 ,8000000 27.3627760 0.0000000 ,0003417 0032440 .9000000 20.9038880 0.0000000 0.0000000 0.0000000 .9500000 . i7.6744440 0.0000000 0.0000000 0.0000000 1 .ooooooo 14.4450000 0.0000000 0.0000000 0.0000000 X CP C = ' .005585 c m 0.000000 --- = ,057603 K 9 0.000000 L 0 L E S C REF n ---- m .920569 -- = -e440087 C 9 -e001935 S C H PROG L 0 INTERFERENCE DRAG OF LOADING 1 (UNIFORM OR CONSTANT J ON LOADING NACELLE BUOYANCY 1 IS 0. INTERFERENCE ORAG OF LOADING 2 I LINEAR CHOROYISE ) ON LOADING IACELLE BUOYANCY 1 IS 0. INTERFERENCE DRAG OF LOADING 3 LINEARSPANYISE 1 ON LOADING NACELLE BUOYANCY IS 0. INTERFERENCE DRAG OFLOADING 4 ( PUADRATICSPANYISE 1 ON LOADING NACELLE BUOYANCY J IS 0. INTERFERENCE DRAG OF LOADING 5 (QUADRATICCHOROUISE J ON LOADING NACELLE BUOYANCY I IS 0. INTERFERENCE DRAG OF LOADING 6 (PARABOLICCHOROYISE ) ON LOAOING NACELLE BUOYANCY I IS 0. INTERFERENCE DRAG OFLOADING 7 I CUBICCHORDYISE ) ON LOADING NACELLE BUOYANCY I IS 0. INTERFERENCE DRAG OFLOADING 0 (SIMILARTO FLAT UING) ON LOADING NACELLE BUOYANCY IS 0. INTERFERENCE DRAG OFLOADING 9 (SP. ROOTFROM T. E. OM LOADING NACELLE BUOYANCY IS 0. INTERFERENCE DRAG OFLOADING 10 I ELLIPTICALC-SUB-P J ONLOADING NACELLE BUOYANCY J IS 0. INTERFERENCE DRAG OFLOADING 11 (LINEARIN ARB.REGION) ON LOADING NACELLE BUOYANCY I IS 0. INTERFERENCE DRAG OFLOADING 12 (BODYUPUASH LOADING ON LOAOING NACELLE BUOYANCY IS 0. INTERFERENCE DRAG OF LOAOING 14 (NACELLEBUOY(CAMBER1) ON LOADING NACELLE BUOYANCY 1 IS 0. INTERFERENCE DRAG OF LOADING 15 (BODY UPYASH(CAHBER)) ONLOADING NACELLE BUOYANCY IS 0. INTERFERENCE DRAG OF LOADING 16 I BODY BUOYANCYTERH I ON LOADING NACELLE BUOYANCY 1 IS 0. INTERFERENCE DRAG OFLOADING 17 (BODYBUOY. (CAMBER) 1 ONLOADING NACELLE BUOYANCY 1 IS 0. YINGDATA FOR NACELLEBUOY (CAMBER)LOADING . .. i69-50017 LOADCHECK CASE 22 SPANSTA. UITHFUSELAGE AND 2 TERMS SPANWISEDISTRIBUTION OF SECTIONDRAG , LIFT,AND PITCHING HOHENT Y SECTION SECTION SECTION "- C C C 812 CHORD D L M G.OCOOOO0 183.8817000 0.0000000 .0002720 -.000351B .0250000 177.9445264 0.0000000 .0004138 -.0005529 . .0500000 172.0073528 0.0000000 .DO09311 -.0012B72 .0750000 166.0701792 .000024'1 .0017408 -.0024903 .1000000 160.1330000 .0000020 .0020751 -.0030665 .1250000 154.1951859 .0000473 ,0029497 -e0044669 .1500000 148.2586941 .0000281 .0032910 -.0051617 .1750000 142.3260032 .0000616 .OOM~~O -.0062537 .2000000 136.3933123 .0001236 .0045904 -a0076394 .2500000 124.6106675 .0002436 ,0054555 -.009e394 .3000000 113.9394744 .0002085 .0072213 -.0145887 .3500000 103.~682~13 .00011e9 ,0088303 -.0200926 .4000000 92.5970882 .0002700 -0105369 -e0267138 .475DOOO 76.6960487 .0003084 ,0093157 -a0277860 .5500000 63.0912977 .0002672 e0096136 "0355341 .62500CO 49.4865467 .0000543 -0092359 -e0454217 .7000000 35.8017958 .0000691 .0081753 -e0574132 .7500000 30.5922200 .0000369 -0043955 -.0370165 .BOO0000 27.3627760 0000009 .0003417 -.0032440 .9000000 20.90388~0 0.0000000 0.0000000 0.0000000 .9500000 17.6744440 0.0000000 0.0000000 0.0000000 1.OOOGOOO 14.445GOOO 0.0000000 0.0000000 0.0000000 X CP C 9 e005585 c 9 .000128 "- . .e57603 K = 4.112798 L 0. L E S C .UEF -. n ----- * ,520569 -- * -.440087 C - -.001935 s C H PROG L 0 " OCAGOF (.OAOING l (UNIFORM OR CONSTANT 1 ON LOADING 14 (NACELLE BUOYLCAl¶BERl) IS .37395446E-02 DRAGOF LOADING 2 ( LINEARCHOROJISE ) ON LOADING 14 (NACELLE BUOY(CANBER1) IS .59821625E-02 CRIG OF LOADING 3 LINEARSPANYISE 1 ON LOADING 14 (NACELLE BUOYLCAHEERI) IS .42389670€-02' D6E.GOF LOADING 4 ( QVADRATICSPANWISE ) ON LOAOING 14 (NACELLE BUOY(CAM8ERl) IS .37495071E-02 OEfGCF LOLOING 5 (PUCORATICCHOROYISE ) ON LOADING 14 (NACELLE BUOY(CAMBERI) IS .729.49422E-02 DRAGOF LOADING 6 (PAhABOLICCHORDWISE ) ON LOADING 14 (NACELLE BUOYLCAMBER)) IS .17314144E-02 ZECG OF :OACING 5 ( CUBICCHOROUISE 1 ON LOADING 14 (NACELLE BUOY (CAMBER) 1 IS .53332648E-Q2 GFAG OF LOACING 8 (SIMILARTO FLAT WING) ON LOADING 14 (NACELLE BUOY(CAMBER1) IS .25098526E-02 ClEAG OF LCACING 9 (sa. ROOT FROM T. E. ON LOADING 14 (NACELLE BUOY(CAMBER)l IS .22607B46E-02? DRAG .OF LOADING LO ( ELIIPTICAL C-SUB-P I ON LOADING 14 (NACELLE EUOY(CAMBER)) IS .3741B132E-02 DRAG OF LOAOING 11 (LINEAR IN ARBIREGIONION LOADING 14 (NACELLE BUOY(CAMBER)) IS ,57341500E-03 !NTERFEREKCECRAG OF LOADING 12 laOD'( UPUASHLOADING ON LOADING 14 (NACELLE EUOY(CANBER)I IS .5225236?€-05 ZNTERFERCkCEDRAG OF LCADING 13 ( NACELLEBUOYANCY 1 ON LOADING 14 (NACELLE BUOY(CAMBER1) IS .12627839E-03 Ib!TEkFEP.ENCE DRAG OFLOADING 15 (BODY UPYASH(CAMBER)) ONLOADING 14 (NACELLEEUOY(CAMBER)L IS .52252367€-04 INTERFEREKCE DRAG OF LOADING 16 ( BODYBUOYANCY TERM 1 OM LOADING 14 (NACELLEBUOY(CAH8ER)) IS e94075333E-05 INTEPFCRENCECRAG OF LOADING 17 (BODY BUOY.(CAMBER) 1 ONLOADING 14 (NAC€LLEBUOY(CAflBER11.IS -94075333E-05 .- - ...... - .. -, . . -.. .. - .- -...... YING DATAFOR.BODY UPYASH LCAHBERJ LOAILING . - - - - . . . .- - - - ..- .n -. 969~50Q.17 LOAD CHECK CASE 22 SPAN STA..YITH FUSELAGE AN0 Z XEMS...... "" SPANYISE DISTRIBUTION OF SECTION DRAGS LIFT, AND PITCHING MOMENT . . . . - . - . , ...... Y SECTION SECTION . SECTION . .~. "- C C .. c ... 81 2 CHORD D L n . .. .. "" .- . 0.0000000 .. 183.8817000 0.0000000 -.0150016.e0165623 . .0250000177.9445264 O.DOOOOO0 -.Dl61667,0173935 .0500ODO 172.0073528 0.OOOOOOD "0177343.0186505 . .075.0000 166.0701792 .0012257 .0222595.-.0212843...... 1000000160.1330000 .0012041 -.0212475.0221333. -... .l25000D . 15S.1951859.0008419 .D216492-.0216256 - . .150 00 00 146.2586941.1500000 .OD06261 -.0222990mO211931 .1750000142.3260032.0003266 .0203227-.0228767 . . .2ODOOOO -136,3933123.0003031.0200290-.0239621 : .25OOOOQ 124.6106.675 .0001082.0191149 -.02626EB . . r3000000. -113,9394744.0000016 -.0287738m0182813 .3500000. 103.2682813 0000806 .0182212 ._ -.0326650 _. .4000000 -e0366937a017712792.5970882.0000677 q475OOO.Q 76.6960487.0000285 . e0174484-a0456375 . . ' .550000063.0912977-.0000430 .0171522-.0580714 .. *62 50 000 49.4865467*6250000-.0000590 a0182157 -a0829274 .700 00 00 35.8817958.7000000 -.0000608 .0209287 -.1379!51)2 .3500000 30.5922200-.0000107 .0218173 . . -.1728406. .8000000 27,3627760.0000405 .0220862 -.1994720 . e900000020.9038880 .0000421 e0201745 -.248131E .950000017.6744440 .0000266 -0173637 -.2569930 1~0000000 14.4450000-.0000333 -.2076505e0112932 X CP C = -018252 C 8 .000260 --- = ,684021 K = .780685 L 0 L E 5 C REF M ----- = -920569 -- = .004682 C 9 e001794 S C M PROG L 0 INTERFERENCE DRAG OF LOADING 1 (UNIFORM OR CONSTANT ON LOADING 15 IEOOY UPYASH (CAMBER)) IS -13207657E-01 INTERFERENCE DRAG OF LOADING 2 ( LINEAR CHCROYISE ) ON LOADING 15 (BODY UPYASH (CAM8ER)l IS .20614524E-01 INTERFERENCE DRAG OF LOADING 3 ( LINEAR SPANYISE 1 ON LOADING 15 (EODY UPYASH (CAHEERl) IS .57260391E-02' INTERFERENCE DRAG OF LOADING 4 ( QUADRATIC SPANYISE 1 ON LOADING 15 (BODY UPYASH (CAREER)) IS .22282178E-02 INTERFERENCE DRAG OF LOADING 5 (PUADRATIC CHORDYISE 1 ON LOADING 15 (BODY UPYASH (CAHBERI) IS e26842756E-01 INTERFERENCE DRAG OF LOADING 6 (PARABOLIC CHORDYISE 1 ON LOADING 15 (BODY UPYASH (CAMBERII IS .23686154E-01 INTERFERENCE DRAG OF LOADING 7 ( CUBIC CHORDYISE 1 ON LOADING 15 (BODY UPYASH ICAMBER)) IS e31360523E-01 INTERFERENCE DRAG Oi LOADING 8 (SIMILAR TO FLAT UING) ON LOADING 15 (BODY UPYASH (CAMBER)) IS e11730779E-01 IETERFERENCE DRAG OF LOADING 9 (Sa. ROOT FROM T. E. ) ON LOADING 15 l8OOY UPYASH IC.AMEER)) IS e16014873E-01. IFTEPFERENCE DRAG OF LOAOING 10 ( ELLIPTICAL C-SUB-P 1 ON LOADING 15 caoor UPYASH (CAMBER)) IS al4316308E-01 i::TEEFERECICE DRAG OF LOADING 11 (LINEAR IN ARE.REGION) ON LOADING 15 (BODY UPYASH (CAHBERI) IS .86849194E-03 IVTERFEREHCE DRhG OF LOADING 12 (BODY UPWASH LOADING ) ON LOADING 15 (BODY UPYASH LCAMBER)) IS .260067D7E-03 ihTCPFERENCf DRAG OF LOADING 13 NACELLE BUOYANCY 1 ON LOADING 15 IEOOY UPYASH (CAMBER)) IS .49983912E-04 IFTEKFERENCE DRAG OF LOADING 14 (NACELLE BUOY(CAMBER1) ON LOADING 15 IEOOY UPYASH (CAMBER)) IS e49983912E-04 fr:rEF.FEKENCE @LAG CF LOAPING 16 ( 6UDY EUGYANCYTERM 1 ON LOADING 15 (EOOY UPYASH fCAI8ER)I IS -.13606499€-04 IKTCRFERENCE DRAG OF LOADIN: 17 (BODY 6UCY. ICAEEER) ) ON LOADING 15 (BODY UPYASH (CAMBER)) IS -.13606499E-O4 . . . ". . - .. . __ __ _. UINGDATAFORBODY .BUQYANCY .TERK LOADING - .- 969-500 17 LOADCHECK CASE 22 SPANSTA. UITH FUSELAGE AND Z'TERHS SPANWISEOISTRIBUTION OF SECTION DRAG> LIFT,AN0 PITCHING HOHENT Y SECTION SECTION SECTION "- C C C 012 CHORD 0 L n O.GOOOOC0 183.8817000 0.0000000 .0012633 -e0027446 .0250000 177.9445264 0.0000000 .0012696 -.0028101 .0500000 172.0073528 0.0000000 ,0012584 -.0028545 .0750000 166.0701792 D.000000D .0012441 -.0029140 . .1000000 160.1330000 0.0000000 -.0000400 -.0023109 .1250000 154.1951859 0.0000000 -e0006373 -.0015437 - .1500000 148.2586941 0.0000000 -.0013829 -.0004346 .1750000 142.3260032 0.0000000 -.0019972 .0005587 .2000000 136.3933123 0.0000000 -.0026088 ,0017498 .2500000 124.6106675 0.0000000 -.0045431 .0052926.. .3000000 113.939$744 0.0000000 "0059995 .0087494 ,3500000 103.2682813 0.0000000 -.0071245 ,0125372 .4000000 92.5970882 0.0000000 -.0084617 .0176057 .4750000 76.6960487 0.0000000 -.0103943 .0282688 .5500000 63.0912977 0.0000000 -.0112420 ,0393547 .625000D 49.4865467 0.0000000 -.0105717 .0491104 .7000000 35.8817958 0.0000000 -.0096770 ,0645111 .7500000 30.5922200 0.0000000 -.0095236 -0761343 .8000000 27.3627760 0.0000000 -.0093318 ,0848577 .9000000 20.9038880 0.0000000 -.0083788 .lo34734 .9500000 17.6744440 0.0000000 -.0066879 .0994734 1 .ooooooo 14.4450000 0.0000000 -.0053462 ,0989333 X CP C "005353 c- 0.000000 --- = -687736 K = 0.000000 L D L E 5 C REF H "_" = .92C569 " * -.DO4838 c- -.000475 S C n PROG L 0 INTERF FR ENCE DRAG OF LOAOING 1 (UNIF0P.H OR CON STANT I ON LOA.DIN IG BODYBUOYANCY. IERH I IS 0. INTERFER ENCE DRAG OF LOAOIHI6 2 ( LINEAR CHORD UISE I ON LOPIOIN IG BODYBUOYANCY TERH I IS 0. INTERFERENCE DRAG OF LOADING 3 ( LINEARSPANYISE I ON LOADING BODYBUOYANCY TERH 1 IS 0. ,INTERFERENCE DRAG OF LOADING 4 ( QUADRATICSPANYISE I ON LOADING BODYBUOYANCY TERH I IS . 0. INTERFERENCE OItAG OF LOADING 5 (QUADRATICCHOROYISE I ON LOADING BODYBUOYANCY TERH I IS 0. INTERFERENCE DRAG OF LOADING 6 (PARABOLICCHOROYISE 1 ON LOADING BODYBUOYANCY TERH I IS 0. INTERFERENCE DRAG OF LOADING 7 ( CUBICCHORDWISE I ON LOADING BODYBUOYANCY TERM 1 IS 0. INTERFERENCE DRAG OF LOADING 8 (SIMILARTO FLAT YINGI ON LOADING BODYBUOYANCY TERrl I IS 0. INTERFERENCE DRAG OF LOADING 9 (sa. ROOT FROM T. E. .I ON LOADING BODYBUOYANCY TERH I IS 0. INTERFERENCE DRAG OF LOADING 10 ( ELLIPTICAL C-SUB-P I ON LOADING BODYBUOYANCY .TERk.J. .IS 0. INTERFERENCE CRAG OF LOADING 11 (LINEARIN ARB.REGIONI ON LOADING BODYBUaYANCY ZERn 1. .LS 0. INTERFERENCE ORAG OF LOADING 12 (BODYUPYASH LOADING I ON LOADING BODYBUOYANCY IERM I IS 0. INTERFERENCE DRAG Of LOADING 13 ( NACELLEBUOYANCY 1 ON LOAOING BODYBUOYANCY TERtl .I.~LS- Q. INTERFERENCE DRAG OF LOAbING 14 (NACELLEBUOY(CAH8ER)l ON LOADING BODY.BUUYANCY T€RH. .J IS. 0. INTERFERENCE DRAG OF LOADING 15 (BODYUPWASH (CA4BER)I ON LOADIIIG BODYBUOYAMCY.. TERH -1. LS . .... Q. XN-TERFERENCE DRAG OF LOADING 17 (BODYBUOY. (CAHEERI I ON LOAO-ING .BOOY..BUOYANC.Z..TE~..L1L 4.. 1 ...... - . . . . - , - .- .. , .. .-. .. .. ". . UING DATAFOR EDDY BUOY. LCAHBERL.LOADING - - .. .. :.. . --969-500 .17 .LOADCHECK .CASE 22 SPAN-STA.YLT . . SPANYISEDISTRIBUTION OF SECTION ORAGJ LIFTJ..ANDPITCHING MHENT " Y SECTION SECTLON SECTION -" C C BIZ CHORD D L H .. . . 0.0000000 183.8817000 0.0000000 -.0010249 -.DO01639 . .0250000 177.9445264 0.0000000 -.00066BB -.0003048 .0500000 172.0073528 0.0000000 -.0002469 -e0009537 e0750000 166.0701792 .0005866 .0012441 -.0029140 .1000000 160.1330000 .0005146 -e0000400 -e0023109 . .1250000 154.1951859 .0002164 -.0006373 -a0015437 .1500000 148.2586941 .0001168 -moo13829 -.0004346 .1750000 142.3260032 .0001143 -.0019972 .0005567 .2000000 136.3933123 .0001008 -a 0026088 .DO17496 .2500000 124.6106675 .0000502 -.0045431 a0.052926 .3000000 113.9394744 .0000442 -.0059995 e0087494 .3500000 ,103.2682813 .0000425 -.0071245 .0125372 .4000000 92.5970882 .0000489 -.0064617 e0176057 .4750000 76.6960487 .0000450 -e0103943 .OZBZbBB .5500000 63.0912977 .0000269 -.0112420 -0393547 .6250000 49.4865467 .0000108 -.0105717 e0491104 ,7000000 35.8817958 -.0000128 -a 0096770 .0645111 .7500000 30.5922200 .0000004 -e0095236 ,0761343 .8000000 27.3627760 .0000162 -e0093318 .0846511 .9000000 20.9038880 .0000176 -.0083788 -1034734 .9500000 17.6744440 .0000120 -.0066879 ,0994134 1.0000000 14.4450000 .OGOOO64 -a0053462 ,0989333 X CP C * -e005353 c * .000110 -" * L 0 L S C REF fl ----- .920569 -- * -a004838 C * -e000475 S C fl PROG L 0 .. . . -. . ... " INTERFERENCE DRAG OF LOADING 1 (UNIFORfl OR CONSTANT 1 ON LOADING 17 (BODY BUOY. (CAHBERI b IS --.26418061E-02 INTERFERENCE DRAG OF LOADING 2 ( LINEARCHOROYISE J ON LOADING 17 IBOOY BUOY. (CAflBEPJ 1.1s -.16409644E-02 INTERFERENCE DRAG OF LOADING 3 LINEARSPANYISE ) ON LOADING 17 (BODY BUOY, UAMBERJ .I 'IS. - -.ZBE23kC9E-02' INTERFERENCE DRAG OF LOAOING 4 I PUADRATICSPANYISE ) ON LOADING 17 IBOQY EUOY.L..(CAHBER&..J-IL _..-..23852bl7E.-02. INTERFERENCE DRAG OF LOADING 5 IOUAORATICCHORDYISE J ON LOADING- 17 IBOOYBUOY* ICAABERJ J If- - -12B955CbE-02 INTERFERENCE DRAG OF LOADING 6 (PARABOLICCHORDYISE ) ON LOADING 17 (BODY BUOY, lCAHBER1 L1S .-.b026757BE-02. INTERFERENCE DRAG OF LOADING 7 I CUBICCHOROYISE 1 ON LOADING 17 (BODY BUOY. ICAHBER) 1 IS -.370456?.7E-03 INTERFERENCE DRAG OF LOADING 8 (SIflILAR TO FLATYINGJ ON LOADING 17 IBOOY BUOY. LCAHBERJ 1. LS -- -.2345142BE-02 INTERFERENCE DRAG OF LOADING 9 ISP. ROOT FROM T. E. 1 ON LOADING 17 (BODY BUOY. ICAHBERJ 1 IS. -a34713632E-02 !-_. ." INTERFERENCE DRAG OF LOADING 10 I ELLIPTICAL C-SUB-P 1 ON LOADING 17 1BODY BUOY. LCAHBER).J..U .. -2L6583267E-02.. INTERFERENCE DRAG OF LOADING 11 (LINEAR IN ARB.REGION1 ON LOADING 17 (BODY BUOY. (CAHBER) I .IS. .. a31926119E-03 INTERFERENCE DRAG OF LOADING 12 (BODY UPYASHLOADING ) ON LOADING 17 (BODY BUOY. (U~ERl-CJf.- -451866323E-D4 INTERFERENCE DRAG OF LOADING 13 ( NACELLE BUOYANCY I ON LOADING 17 (BODY BUOY, ICAH8ERL.J. IS . e168655BbE-05 INTERFERENCE DRAG OF LOADING (NACELLEBUOYICAHBER)) ON LOADING BUOY. (CAHBERL 1 . c 14 12 (BODY IS .16865566E-05- VI INTERFERENCE DRAG OF LOADING 15 (BODY UPYASHICAflBER)) ON LOADING 17 (BODY BUOY. .LCAMBERJ .1 IS. - rr5lBB6323E-DC VI INTERFERENCEDRAG OF LOADING 16 I BODY BUOYANCYTERH 1 ON LOAOING 17 (BODYBUOY* (CAHBEP) 1-1s. . .11019121Et03.- . FORCE COEFFICIENTS OF COHPONENTAND INTERFERENCE LOADINGS 969-50C 17 LOA0 CHECK CASE 22 SPAN STAB WITH FUSELAGE AND 2 TERMS . GROSS YING AREA lC752.043141 SREFISPROG * . a.926569 CI. 1 = .938399 FOR UNIFORM OR CONSTANT LOADING CL i 9 ,908181 FOR LINEAR CHORDWISE LOADING CL 3 1.020416 FOR LINEARSPANYISE LOADING CL 4 = 1.033028 FOR QUADRATIC SPANWISE LOADING Cl 5 -917149 FOR GUAORATIC CHORDYISE LOADING C: 6 = .917731 FOR PARABOLIC CHOROYISE LOADING CL 7 * .941350 FOR CUBIC CHORDWISE LOADING CL 8 = -933996 FOR SIMILAR TO FLATYING LOADING CL F 9 ,986059 FOR Sa. ROOT FROM T. E. LOADING CLiO = -919594 FOR ELLIPTICAL C-SUB-P LOADING CLll -036338 FOR LINEAR IN AR8.REGION LOADING CLl2 9 .018252 FOR BODY UPYASH LOADING LOADING CL13 * .DO5585 FOR NACELLE BUOYANCY LOADING CL14 ,005585 FOR NACELLE BUOY(CAM8ER) LOADING CL15 * -016252 FOR BODY UPYASH (CAW8ERI LOADING CL16 = "005353 FOR 8001 BUOYINCY TERM LOADING CL17 -e005353 FORBODY BUOY. (CANBERILOADING c-n-0 1 * .Dl9612 C-M-0 2 "056583 c-R-0 3 0 -e147457 C-M-0 4 9 -.282881 C-H-0 5 = -e115619 C-M-0 6 * e171901 C-M-0 7 = -e039700 C-H-0 8 * .111885 c-n-o 9 * .175704 C-"010 * ,056104 C-"011 = "013476 C-H-012 * e001794 C-"013 9 -e001935 6-H-014 = -.OD1935 C-"015 * e001794 C-M-016 9 -.DO0475 C-"017 * -.OD0475 cc 1 lI(CL 1)(CL 11 * ,734870 co 2 2I(CL ZI(CL 2) * 1.598735 CD 3 3/(CL 3)(CL 3) = ,771391 cc 4 4/(CL 4)(CL 4) = 1.299544 CO 5 5IlCL 5J(CL 5) * 2.905185 CC 6 6/(CL 6)(CL 6) * 2.254638 CC 7 7/(CL 7)(CL 7) * 3. t5.4782 cc 8 8/(CLBI(CL 81 = e688888 CC 9 9/lCL9l(CL 9) = .955578 CD1010/(CL10)(CL10~* ,833103 CD1111/(CLll)(CLl1) * 5.926973 CDlZlZ/(CLlZ)(CLlZJ = 0.000000 CC1313/(CL13)(CL13J * 0.000000 c01514I(cL141(cL141 * 4.112798 CC1515I(CLl5)(CL15J = 0 780685 CD1616/(CL16)(CL16) 9 0.000000 CGl.717!ICLl7l(CL17) 3.845320 1 ICD 1 2tCD 2 1)IICL 1llCL 2) = a 1.955786 (CD 1 3tC.O 3 1lIlCL ZJICL3) = 1.030882 - ICD. 1.4tCD.41)IlCL 1)lCL 4) = -85666.6 .. ICD 1 5tCD 5 lI/(CL 1)lCL5) = 2.297994 J I.CD..L.Q+CD -6. l)/ICL 1)ICL 6) * 2.068573 ) ICD 1 7tCD 7 1)IICL 1)ICL 71 = 2.479876 1 (CD 1 8tCD 8 1)IlCL 1)ICL 8) = 1.380664 .). ICD 1 .9*CD 9 lII(C11)lCL 9) = 1.557544 ... .. iCD llO+CDlO1IIlCL l)(CLlD) = 1.545274 ICD 111+CD11 1)IlCL lllCL11) 2.098149 tCD 516tCD16 .51!iC% JllCL16 IC0 112tCDlZ ll/(CL 1IlCL12) = -703440 ICD 517tCD175)IICC 5!!CC1.7 (CD ll3tCD13 1)IICL lI(CL13) * ,742224 (CD 114tCD14 1)IlCL 1)lCL14) = 1.455772 (CD 115tCD15 1)IlCL lllCL15) = 1.474581 ICD 116tCD16 1)IlCL l)lCLl6) .242672 (CD 117tCD17 1IIlCL ll(CL17) = ,760576 2i.189628 lCD.23tCD 3 2)IlCL ZIICL3) = ,916458 .949892 (CD 2 4tCD 4 2)IlCL Z)(CL4) = ,486254 -.140363 ICD 2 5tCD 5 2)IICL 2)ICL5) = 4.176316 rl.97451 (CD 2 6tCD 6 2)IICL ZIICL 6) = 3.202498 2.363973 ICD 2 7tCD 7 2)IlCL ZIICL7) 4.517035 -e499868 (CD 2 8tCD 8 tJl(CL 2)ICl 8) = 1.744246 .7273D3 (CD 2 9tCD..9 2lIlCL 21lCL9) 2.109941 2.248246 ICD 210tCD10 Z)I(CL 2)(CLlD) = 2.141035 2.940951 ICD 211tCDll 2)IICL 2)ICLll) = 4.106713 Zr78739Q ICD 212tCD12 2)ICLlZl .71e483 ZlI(CL = (CD 711tCDll 7)IlCL 7llCLll) = 5,441504 ICD 213tCD13 ZIIICL Z)lCL13) -932539 (CD 712tCD12 7)IlCL 7lICL121 = !E43316 (CD 214tCDl4 2lIlCL 21ICL14) = 2.111984 '(CD 713tCD13 7)IICL 7)(CL13) ? .542912 ___ ,(CD.Z15tCDl5 2)IICL 2)(CL15) = 1,962127 ICD 216tCD16 ZJIICL 2)(CL16) = -.734548 .. (CD 217tCD17 2)IlCL Zb(CL17) = -.397008 (CD 714tCDl4 7lIlCL 7llCL14) " 1.557310- 3 4tCD 4 3)ICL4) = (CD 3)/(CL 1.626574 ICQ 715tCDl5 71/(CL 7llCL15) * 2.668586 ICD 3 5tCD 5 3)IlCL 3)ICL 5) = ,758597 IC0716tCD16 7)IlCL 71lCL16) -1.753682 (CD 3 6tCD 3)IlCL 31ICL 6) = 6 .481282 ICD 717?CDl? 71IiCL 71l.CLl7) -1.6BDl66 3 7tCD 7 3)IlCL 3)tCL 7) = ,522459 . ICD (CD 8 9tCD 9 8JI(Cl 8I(CL 9) 1.570087 (CD 3 8tCD 8 3)IICL 31(CL 81 = ,885007 (CD 81OtCD10 8)I(Cl8)ICLlDl = 1.467247 ICD 3 9tCD 9 3)IICL 31(CL 9) = 720350 IC0811tCD11 8)IICL 81.lCL11) 1.665571 ICD 31DtCD10 3)IlCL 31 ICLlOI = ,923541 ICD 812tc~1ze)/fcL B)ICL~Z) = .754155. ICD 813tCD13 8IIlCl8)(CL13) = .564739 ICD 81'1+CDl4.8lI(CLBI(CL14) f 1.045905 ICD 311tCDll3)IlCL 3)(CLll) * 1.087002 (CD 815tCD158)/(CL 8IlCL151 1.442296 (CC 312tCDlP3JIlCL 3)ICLlP) .560685 ICD 816tCD168)IICL 8)lCL161 * a326039 (CD 313tCD133)IlCL 3)(CL13) = 1.054161 (CD 817tCD178JllCL 8)ICL17) ,795207 ICD 314tCD14311lCL 31(CL14) * 1.797993 (CD 910tCD10QIIICL 91[CLlD) = 1.709106. (CD 315tCD153)/ICL 3JICL15) ,868134 (CD 911+CD1191/(CL 91(CLll) 1.723090 (CD 316tCD163JIlCL 3)(CL16) = 1.043194 LC0 912tCR129)/(CL QIICL12) .793 197 ICD 317tCD173)IlCL 31(CL17) = 1.570863 ICD 913tCD1391IlCL 9)[CL13) a378706 9 ICD 4 5+CD 5 4)IlCL4I(CL 5) = ,234655 (CD 9lCtCD14 .9)IlCL 91 lcL14) .= -789239 IC0 4 6tCD 6 4)IlCL4)ICL 6) = -a054440 IC0915tCD15 9)IlCL 91lCL15) l.ba9045 (CD 4 7tCD 7 4)I(CL4)lCL 71 -.O22089 (CD 916tCD169IIlCL 9)(CLl61 ,159638 4)(CL (CD 4 8tCD 8 4 IICL 8) = -683763 (CD 917tCD1791IlCL 911CL171 a817356 IC0 9tCD 9 4 41lCL9) = 4 IICL -352524 5 2.279159 41lCLlDl = (CDlOl1tCDl11D~/lCLlQ~lCLll~. (CD 410tCD10 4 I(CL -612361 (CD1Ol2tCDl21OlI(CLlOl~CLl2t ..721374 ICD 41ltCDll 4 IICL 4)lCLlll = .672512 5 -723195 ! ICD 412tCD12 4 4jlCL12) = (CD1D13+CO13101/ICL1R~~CLl31 IICL ,548205 (CDlOl4tCD14101/~CLlO)~C114) 1.451776 JCD-413tCD13 4 IICL 41ICL13) = ,985951 (CDlOl5tCDl51Ol/lCLlO)lCLl5l lr574337 ICD 414tCD14 4 I(CL 4)ICL14) = 1.635863 ~CD1016tCO1610~/~CL1OllCLlb~ a122538 (CD 415tCD15 4 IICL 4)(CL15) * -666304 . ICD1Q17~C~7lO~/l~LYP)ICLl7~! - r6h2552 . . (CD 416tCD16 4 IICL 41(CL16) = 1.248766 lCDll12~CD12ll)/~CC1l~~CLl2) a441247 (CD 417tCD17 4 IlCL 4):CL171 = 1.680190 lCD1113tCD1311lI1CLllloL13) 2.572ea,z. ~CDll14*CD~~11)I~CLll~~CLl4)= 5.398406 . .. ~CDl115+CD151l)I~CL1ll(CLl5~ .!. 1.750728 ~CO1116+CD1611)/~CL1l~~CLl6)= -1.728767 -.. . . .KD1117+CD1711)/~CL11)(CLl7~= -3r37005.6.- . ~C01213+CD1312)11CL12~(CL13)= 0.000000 . . (CDlZl~+.CDltl2)l~CLl2)(CL14)= ,512617 ~CD1215+CD15.12)/(CLl2)(CLl5) = .700685 .. . (CDl216+CDlbl2)l~CLl2)~CLl6)= 0.000000 ~CDl217+CDl71Z~l~CLl2~~CLll)= ,592467 - ." (CDl314+CDl413)I(CL13)(CLl4) = 4.112798 (CD1315+CD1513)I(CLl3)~CLl5~= ,490363 -. (CD1316+CD1613)I(CLl3)(CLl6) = 0.000000 .. ~CD1317+CD1713~I~CL13~ICLl71= "056414 !CD1415+CD1514lIlCL14~~CL15)9 1.002980 ~CD1416+C01614)/~CL14)(CL16)9 -e314674 (CO1417+CDl714)/~CLl4~(CLl7)- -e371088 (CD1516+CD1615~/~CL15)(CLLb)= -139263 (CDl517*COl715)I(CC15)(CLl7) = ,731729 ~CD1617+CD1716)I(CLl6)(CLl7) = 3.845320 CO YING-L IFT-ON-NACELLES 1. .002214 CO YING-L I FT-DN-NACELLES 2. .002290 CO WING-L I FT-ON-NACELLES 3. -002385 CO YING-L I FT-ON-NACELLES 4. ,001882 CO WING-L I FT-ON-NACELLES 59 .001639 CO YING-L I FT-ON-NACELLES 6- -002983 CO UING-L I FT-ON-NACELLES 79 .002105 CO YING-L I FT-ON-NACELLES 89 no01814 CD YING-L IFT-ON-NACELLES 9. ,002427 CO YING-L I.FT-ON-NACELLES 10 9 ,002254 CO YING-L I FT-ON-NACELLES 11 = -e000199 CO YING-L I FT-ON-NACELLES 12 = .000037 CO YING-L I FT-ON-NACELLES 13 = 0.000000 CO WING-L I FT-ON-NACELLES 14 -.000027 . __ CQ WING-L I.FT-ON-NACELLES 15 = .000037 CO WING-L I FT-ON-NACELLES 16 - 0.000000 CO YING-L I,FT-ON-NACELLES 17 = -.OOOD22 ...... _. ._... dRDY_zERILs. . -. ..- .. - ... - ." .. - ...... - -. -...... "... -...... -. - ... " " " - -. .. _...... - ". - "- - - ...... " " .... c-L..~" --.OOOOQ Go .noooo~ c-n-=. ,~03.9sn..x~ .=. .. -..ZZLUIL- ALPHA,^. .. - ~coooa.__ c-n-0-L 40a... .- ...... - -_ "" .. - ._: ~. -" .. . ." - .- _- -I&T, MOMENTANDORAGL INCREMENTS . ." - . - .. - ...... -- . -" ...... - -. - - - ...... - - - - ..... - . __ - .... BOPY .DUE .TO LARRY-OVER OF YING LIFT...... - - ...... -.- - . - .. . - .... -. ... __ - ..... -. .. - . " - . - . .- - -. - . - - ._. -. .- -. . - .. - - "... - . .. ". . c .... -. ..- ...- - - - ... "- - - ... - ..- - ...... TYPE w..- -t c. m ..... - ..; ...... I . UINGIIET.LOARIHG . L 0 " 0 ...I- ...... - . " ___ - .. ._ .... _...... -...... " . . . ._ 1 UNIFORfl (IR CONSTANT .118088.010218 a032205 1 ...... -2 LINEAR CHORDYISE -165429 -014306 ,006463 2 .3 LINEAR SPAWISE ,038359 ,003319 .OD8244 3 ...... " ..... 4 PUAORATIC SPANYISE .010797 0000934 .001281 4 ...... 3. _PU.ADRAUC .CH[lgRYISE-207608 a017926 -a016465 1.. ... " .. - . - - .....- "_ ." .- .... - b PARABOLIC CHOROYISE e238510-020680 ~060842 6 ." ...... 7 CUBIC CHOROYISE -.003052,023470*271410 7 ...... - - ..- - - ... 6 SIRLLAR TO FLIT YING .120023a010382 -049734 8 ...... 9 SQ,FROM ROOT T. Eo -.161934 ..014018 *0627189 . "...... - ...... LO ELLIPTICAL C-SUB-P .128894e011153-035309 10 ...... 11 . LZYEAR IN ARBeREGION ,003426 e000294.-.001259 11 . . .. - ... ._ ...... - - ...- -...... 12 Boor UPYASH LOAOING .003074 .ooom .00105o 12 ...... 13 NACELLE BUOYANCY e000086 .OOODO7 -.000035 13 .- . .. I4 NACELLE BUOY(CAH8ER) .000053 .000005 -.000022 14 .... BO DY UPYASH15 15BODY .000820(CAHBERI .000230 .002659 - .. ... 16 BODY BUOYANCYTERM .000182 .000016-e000350 16 . 17.0UQY. . BODY (CAMBER1 -e000295-.000007-*000069 17 ..." . OELTAT =.. .422 SEt.. T := tO8Z.;OlSEC...... RESTART DATA PUNCHEDI DECK IMAGE FOLLOWS. 969.-500 17 LOADCHECK CASE 22 SPAN STA. WITH FUSELAGE AN0 2 TERMS RESTART 21 .1.220. 41 .20 25 11 21 1917 22 46 45 43 42 40 39 37 35 34 ..... - ...... 3124272420161310 9 8 6 65 5 123 4 5 6 16 17 14 13 15 12 15 13 14 7 8 917 10 11 16 ." ...... 1191436.9Z6281E+Olol53.k397023106E+01 .9087199921286E+00 .7644823748055EtOO .. ~1820676OO6052E+01.163994889748lE+Ol .2016624882681E+01 a1113978786573Et01 .. . .1326744771816E+01.1227567816936E*01 .658b312236193E-01 ~1109113885135E-01 m3580BI0U4972E-IU.7023367855027E-02 e2324970154520E-01 -.1122202583347E-02 ._ . ._._ ...... - .. ._.-a3554167929359€-0201534397023106E+01 ~2427771011671E+01. .7818414627068Et00 .. .~4199359911+23E+OO ~.3202299855232E+01 .2457158862649E+01 .3554950491973E+01 ~1362013949550E+01 ~1739411186432E+01 ~1646070902093E+Ol .1247628099813EtOO r109635232721OE'Ol .4354154169237E-02 -9861148813370E-02 .2994061940195E-01 q3287426509946E-02 .1776786738630E-02 .9087199921286E+OO .7818414627068E+OO -.*.~~70819.72783~kE+01 .1772485513297E+Ol ~6531566609634E+00 .4149047863636E+OO ...... , o46.l9943bl~736E+OO.7764707795822E+OO .6672366652443E+00 .7977846344839Et00 .)710443~b1398E-01,9612965507468E-02 .5530296196860€-02 .943256075628bE-02 _.- e1k884180851k4E-01-.5245728207967E-02 -.7899126215912E-02 .7644823748055EtOO .ZQ46627129120~5+~0 .33Q!5673,019399E+OO 523638992984.7L-02 -* 8553305482~91E~OZ ~2046627129120EtQO .1586538597773€+01 ~1109931845477€~01 a722359842994CE-02 ~41kQO~7863636EtOO ~.~4Obl1k.CZZ6037E+OL .6722086987816E-01 ...... TEA2539 17 LOAOIMG VERSION OF OCTOBER 30, 1975. OPTIMUM COHBINATION OF 17 WINGLOADINGS ...... 969-5Q(L-.l? _L-CHECK-CASE 22 5M.N STA. YITH FUSEC.AGEAN0 2 TERHS _- .- " . . . - - " NuaeERoF..~NEOqy..B~nKPOINLsI_o- FLITPLATE%UdL- -1. rl ~ . NUMBEROFSEYISPAN kLEYENTS = 40.0 PRINT FLAG = 2.0 NUHBER OF SPANSTATIIINS FOR CAMBERSURFACE 22.0 SHOOTc(1NG FLAG = 1.0 - SPAU-IATloYR-PnRA~U& APEX-I-=!LO--" _ REILAFLL ILA.riz.0 .. . . BA SIC HACH NUHBER HACH BASIC = 2. ?PO0 C-L DESIGN = .loo0 ~- ...... " CRAR. 1C0.4 __ .. NUMBER oF..4tADIh&S~~-17.0000 .. .. -. PITCYIHGMOMENT CENTER AT le7.oiGo: NUHRER r)F CAMBERORuIhlATES = 12.00(r0 REF€RENCFARtA 9 Y898.OOUO NUMBEROFPOINTS DEFlNING ARBITRARY REGION = 2.0000 "0 CONSTRAN.1". ._.P100""."~€~UI~U9LO ". .. -. . - . - SPANSTATION FOR SIDE-OF-BODY = 4.9688 CAMBERBODYJROINATESNUMBER OF = 19.0000 BODYUPWASH LOADING TABLE 0.0 0.0 ~- -"" -- NAc+_LLL.Bl!.LDUNCLW~US 0.0 0.0- ." YINGUPPER SURFACE LIhITING PRESSURES * 2.0 2.0 UINGTHICKhESS PRES5bRES = -21.0 0.0 " " -______~ CAHBERSIJRFACE OPTION FLAGS = 1.0 1.0 1.0 3.0 "- 4 CONSTRAINTS4PPLED ARE ON ORDINATE- ____ "- ". - .- - CON9TRAINTLOCATIONS - " _. __" "" ". I X(!) YI I) Z( I) PLANFORHOFFINITION X X (LEADINGEDGE) Y (TRAILINGCHORDEDGE) __ - .. -. 1 59.999300 0.000000 183.861 700 243.et11noo I 2 77.328000 4.068890 166..6?000(, 243.3q8000166..6?000(, 4.068890 77.328000 2 " 3 83.104OnO 6.6L5000 16.9.133000 2413_3700(,"- "" 4 93.165000 9.510000 149.790000 2'12.0550~0 5 116 960303 n6.333000125.35U030 242.3103OU 6 "____ 77.295090168.980900 31.250(r00 246.275000 ______"" r 225.810000 47.544000 3E.6R1000 258*491CGO 8 225.810000 . 47.*45000 32.6810i#O 25B.491OUO 9 258.210003 9 66.25~000 143 445000 272.655000 -. . 1 -"- - " VALUESOF SEMISPAN LOCATIOY AT YHICHYING CAMBER SURFACE UILLBE CALCULATED 2" .-_apOOD 2" 3.00UO 4 .DO00 5.0000 6.0000 7.OOUO' 0 .oooo 10.00110 12.0000 14.0000 16.0000 19.0000 ZP.OOOO 25erlOUO 20.0000 30.0000 32.0000 36.0000 30.0006 40.4000 -__ __"" __ - - ". WING GRID SYSTEMPUTS SIDE-OF-FUSELAGE AT Y. 5.14063 AT EDGE OF'ETFHENT ROY. 3 Y .- - .. " - - ". "i UPPER YINGSURF4CE LIHITING CP TABLEI -". 0.00000 66.25Oi~O x STATIONS 0.00000 100.00000 X 207.00000 269.80000 Y STATIONS 0.00000 100.00000 .. ""- -" "" " - ARBITRARYRtiGION DEFINITION (LOADING LIMIT C-P -.137COO -.137000 "FRACTION OF I_EM.€SPW __ -.1370@0 -.137000 0.0000u 1.00000 ...... FRACTION OF LOCAL CHORO ,79943 ,60235 *******************v******************* "_ .______"" C-P LONGITUDINI_C CRA?!-ENT LIMIT e002500 eOOZ5OC .002500 m092500 ~ - ***************r*.~-~*~*'~' .. .. - .-...... -...... ...... - ...... """"_ _."-"...... . . -.. -.- - , .. ._ ___ **********.*+*S*i*+;*~*~.***~~*.*~~;.**~;.*~ SOLUTION FOR DESIGN C = .100000 L .. "...... _ ... -. .. ." ". _" _. __ __ . " MITH c CONSTRAIN~D-T~"- " .o1ooco M " " " ._ .-. "." - .- - -. . - - ..- ...... ***.j.*rj****+*~****~**~*~**j+*+.+j.******"*.0 .. - ...... "_____ .- ... H 0... - -2 3 . 4. ... . 5 ." 12 13 14 15 ...... - "...... -. " ... " ." " - ...... -...... - . WING .09005 eOJ3b74. .GO2035 ...... FUStLAGE .OOOOO .000001 e003958 WINGINOJCED ON FUSCLAGE .o0505 .003207.0007e2 __".---.-. ~ WING _I.N.JU&t,D. 0.N NACELLES .C,@OOOQ. ... .OO.O154.-~rO~~~O! ___ """ """_ """" TOTALS e004614.10000.010000 ...... - - ...... - _" - -- .- ...... A" .... __ ...... _ -...... " ...... " .. .. "- ...... - - ...... - - ..... - ...... - - I .- . LOVER UPPER' UPPER Y X X-PRIM LIFTING TrlICKNESSSURFACE SURFACE SURFACE -" C C I ~ oc ox " - ...... """_ "...... c ...... c- "" et2 L CHORD P P P P P ""...... eF75003 . 5 ~.2i... .22'2il4 - e008989 . e07 5000 e274554 e247227e274554 e075000 e008639 ___ .075000 ?29?08e_-, .272?40 . . - e006687 e07 5000 e313621e075000 297252 e004536 -075 000 ,333154-075000 ,322265 .004300 __. - 075000...... - .j=e8?--. .. 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" ...... -. 0250000 -372220 ..171685- -.004577.lo90771 ,035988' .-.Oiaoee .250000 a391754 .265020 e094943 -.00513C e030893 -e064050 -. .???!Qp!?" .-411267 ~2,36354 094688 . -e004663 e033232 -D.0614 56 e250000 e430820 e271689 -.bo3719e092279 m034Yll 057307 e250000 ,450358 e305024 .bd8 76d -.002499 ,036266 -e052503 -"" - -- - "...... " ...... - - ". . . . - " - - " .338358 .088051 -.00116b' ,371693 .Os5804 -.001090 ,037926 e508953 ,405027 e081853 -.0017bl .. e036031. 250060 .52b486 e438362 .074853 -.002828 a031564 .250000 .54d020 .471697 e066707 -e003287 ,027132 .250ooa .- ?,567553 . 505031 ..c>5955 -moo3561 . at21561 ~250000 .5870bb e538366 .035447 004628 -010207 -e025239 002256 e 250000 666619 .571701 .014lOU -.006476 -e002347 -e016448 .002062 -a013422 0 250001, 626153 e605035 -.002996 -e008883 -.01C4Zb 2 5 0000 .b45686 e638370 ,00731~ -.012084 -e012069 -.019380 .25000U 665219 e671705 eOL7c172 -e015067 -e010772 -.a27843 .25JOOO -684752 .705039 a025448 -e017772 -a009763 . ..-no35211 250000 a704285 .73b374 .028206 -e019572 -.Oil9951 -e038157 250000 -723819 .771709 .530739 -e021503 -.010381 -e041120 -.060722 250003. ,743352 .8C5043 -033562 -.0L35b3 -.oloroo , -.044462 I 250000 7b2b85 .83 63 78 068453 -.0253b3 m040428 -e048022 .250000 .78L418 ,871713 e061572 -a028363 eGZ72t4 -e054369 -.DO1544 ~2~0000 e801932 -905047 ,075813 -.030673 .015023. ,. -.060790 .. -*OG14?7 . .250003 d21485 -9303 62 .07+133 -e032473 ,007422 -.Ob6771 -.061612 "250000 ed41018 .97171r eirbLO81 -e033534 -.011107 -e073188 -.001554 , rn 250000 -85 759 1 1.0c0000 .u5a852 -.034100 -e023283 -.079135 -a001875 300000 e325939 O.DL0000 .ZL4271 .027jCG ,117985 -.lo6285 -e005351 e300000 .352987 .i)49923 -194242 -.U06747 e071G44 . -a123297 -.Ob1432 . -~16oooo'- e372220 .Ob6380 .15Y 09 5 -.010001 mO51913 -el07182 .004373 CP GRADI€NT LIMIT = moo250 .. 0 300000 -3917% e122837 ,137441 -a010424 -042884 -e094557 ,022346 . . . .. 300000 .C11287 159293 -123374 -e008849 aQ39Y58 -e083415 ,003149 CP GRADIENT LIMIT = e00250 30COOO e430dZO -195750 .lt5773 -.006297 . e.0 3 6244 -?Ob9529 a003523 CP- GRAOIENT LINIT i a00250 306000 .+SO353 -232207 a134052 -a005936 eU37602 -a066449 .000502 .300000 a469887 eZt8bb3 .A02467 -.005377 e039136 -.C63331 .OCO98d .. .. 300000 .43V420 e3C5120 .n99338 -e004326 .OCOZ64 -a059075 .OCOb57 300000 .bo 895 3 e341577 .OY7235 -.c103086 eC4130tl -a055927 -000855 .. .. 3o0009.. .- - .528+pb e378033 .093577 -e003473 G399 50 -e053628 .009523 .- .300000 m548020 .414490 e3b7979 -.003365 .037770 -.OSC209 0'21 234 300000 -5b7553 ,450947 e680299 -e001857 -033417 -.044882 e000464 .300000 - ,587386 m467403 e071495 ' -e004044 e028609 "OCZb85 000592 .._ - . - . 300000 606619 5238 60 .u54 950 006041 .016821 -e036130 .CC,2124 .30JOOO m026153 560317 .0339jb -.008163 .006638 -e027318 .O(i2033 __ .~?OOOP~...- 645686 . e596773 .U121/3 . -r010861 _.. -909b.514 ..,:a018686 .902121. . . . _. e 300000 ,665219 1633230 ,005 588 -.013360 -.013611 -.01920i) -.OC'J308 300000 -684752 ,669687 -e000030 -a 01 5602 -.020140 -.02c110 - ...300000 704285 e706143 -e (tu2 470 -.017988 -a024632 -e022362 300000 e723819 .742600 ,012714 -.022071 -e020484 -e033199 30COOO e743332 e779057 ,027685 -e024818 -.014909 -a042589 . _. 3-00003 ..,762d115 81.5513 .0-91127 -e027317 e04274.9 -.048378 I 300000 782418 .85i970 .Od7056 -a029851 ,031331 "055725 300000 ,801932 888427 aOd4092 -e030799 e022475 -.061b17 - - .300000 8214115 a924883 .08159!1 -e032693 a014950 ". -e066645 .300000 941018 ,961340 e072796 -e034480 .000&18 -e071978 .300000 .ab~J51 ,997797 .094949 -a037168 e017854 -.077095 - ..300000 .. e1161.732 1.000000 .094101 -e037300 .. .e036537 ~077_56_4. 350000 m380260 0.000000 223556 049000 ,141628 eC23669. . .218003 0274 61 474q4 -063893 .lo8Ub7 e000277 ,0777 55 e104117 e137037 -e008635 .055001 "...... "" . . - .." .- ." . . . " . - .. " e350003 e450353 .14434i - e143522 -e013864 e045629 350000 ,469687 0184565 ,125554 -e011489 e041635 .35aooo- .489420 -224789 .llb898 -e009455 a041273 e350006 e508953 e265013 ,113489 -.007209 e043324 .- .350009" .52848b . a305237 elGb456 -.OD3531 -045909 .350000 .548020 ,345461 .lo5112 me003773 e044860 .35000rl e567553 .3@5bt14 e AOC171 -e004157 ,042871 a000 194 350000 .serosb ,425908 e 093020 -.005129 ,038644 I 350000 .bo6619 466132 .0&4181 -.006029 e033344 350000 e626153 e5Gt356 ,973211 -.006567 ,027402 .350005 ,645686 ,546580 L155782 -e008256 e017400 3500G0 a665219 e586804 -037506 -a010755 *GO6166 e350000 e684752 627028 ,020193 -e014141 -e006352 a 350000 704285 ,667252 -003097 -.01795t -e019021 .350000 ,728819 e767476 -.OOb771 -.021519 -a029828 .J50@00 e743352 e747700 ,064024 -e023 771 -.025542 . _..350000 .rbzab5 .78 2479 .3rbb19 -a026479 -0021 bl2 .350000 .782418 e828147 ,077660 -e029214 -031337 350000 .bO 19s 2 ,868371 ,083500 -a032243 aG25942 350000 m321485 .308595 ,129129 -a035123 .Ob2701 .350ccI) ,841018 m948919 ,119329 -e036169 .046491 .35JOO0 .Ob0551 m98F043 -113681) -.03784C e032500 3500GO e065872 1.ocoooo .110715 -e038300 ,027354 ,400000 a434561 0.000300 .1Y8791 .0405GC .123095 -.010542 400000 ,450353 ,036223 .A99737 mGlZ173 a0964C1 -.00661b 400000 .4b9887 m081083 ,181358 -.(to5873 .07OtOO OCO6Y 3 .COO003 419420 ,125942 .162622 -.01139C .05700U .002037 400000 ,508953 ml7OdO2 i44 679 '.011984 .050+68 . a003298 CP GRADIENT LIHXT = .400000 e528486 21 5661 129202 -.011118 .046311 .GO1103 .400000 .54[1020 e260521 m123013 -.OU9846 ,045855 .OGl898 .400000 e567553 e305380 ,113279 -.007099 eC46046 e001135 .400000 e587006 .35C1239 .A11744 -.C04603 e041674 a000617 .400000 bOb619 ,395099 .LC6546 -*005141 .045435 .OGOdO2 .. .400000 ,626153 ,459958 e098348 ...005680 e040914 0000964 I 400COO -645686 -464818 ,086716 -e007471 -034329 .ooor1c .400000 ebb5219 ,529677 e075654 -.009921 ,025378 *001231. . __ ,...... COO000 .be4752 .?I74537 .Ob0461 -e012618 .01>422 .OUl204 400000 ,704285 ,619396 -041962 -a015686 .bo3437 eOC2019 e4OOOOD . .. .723r)19 604256 ,019508 -.01845J -.01O756 e001694 ...... 400000 a743352 ,709115 .JOG299 -.021829 -e024167 -.OCOb3b .400000 .762885 ,753974 004 565 -.025466 -.031648 -.000734 ,400000 82,18 -e .r .798834 -e009592 -.U29593 -.039702 000 727 " .. ,400000 .001952 1843693 eC66033 -e632933 mOlU132 -.0(475d .4nooo0 -821465 ,886553 .u4aa3 -e034636 e06028b -.OC4732 9 4opoooo b41018 e933412 e121740 -a037503 eC46970 -000170b . . . . . - -1 - .400000 860551 .97b272 .115028 -.038496 -034034 -401883 .COO000 .b70013 1.cIcocoo 110013 -e038800 ,324172 -.002160 .- . .- ...... , .4-75000 e516063 0.OLG000 elb8089 e025200 e105245 -.Oll252 *475000 .5204tlb .a34447 e197122 .OObllb ,092148 _, e r75ooo. .54802I) 088607 .lbb268 -.010878 .or1269 .47 4000 e267553 e142767 172 107 -.016002 e061099 e475OCO .587086 e196927 .A33605 -.015746 0054288 .- .r'i5ObO e606619 .251087 .141911 -a009515 ,056031" o 475000 o 626153 m305247 ,128733 -e007887 .OS2442 .. 4 0 ". - - . " - ...... - "- - . ._ ...... - - .. . -. "_ -47 5003 .6C5686 .A 5940-7 -123546 e475000 ,665219 e 413567 -116352 e467727 -0 47 5000...... 684732 107126 e475000 :704205 e321887 ,095641 e475009 723813 576047 .oa1627 .475000 . ...7433?2.". .e630207 m0631b2.. ,475 000 -664367,762865 ,047 593 4750C3 ,782418 e738527 e071663 e475000 . .a01952 .79268r .03Ul53 .005495. 4 75 000 ed21Cb5 e846847 eU3Y073 -e004491 .4 75000 .841018 .9c1007 a345352 -et12695 ,. ,. 475000 . $60551 ,955167 r092 647 -.022387 .475000 .E76721 1.ooocoo 1062685 -e027177 .550000 .5975?0 0 .OGOCOO 176762 .0961,8l m550000 e626153 -596389 eldV315 ,075439 550000 eo45686 le2228 e~78653 .Ob6602 e002053 . .55OOCO .565219 . .22bg67 163!56 a062252 ._-093167 . ._ - ._ CP GRADIEf4T LIMIT = .00250 556000 ,684752 e293905 .147518 ,059254 e003177 . CP GRADIENT ,LJMIT, ! . -00250 . .55O@OO .73428S .3 59144 136390 e055364 -001459 550000 e723819 e4253d3 126105 e049251 -55 0000- .- 143352 e491422 .117613 . .04203& . .. 55@000 e762865 .557260 .13454L .032750 .550000 ,782418 -623099 e3tJ9172 e020761 .550000 e801952 ,666938 e115693 05 3bO4 .550000 e82~485 .754777 ,081185 -025673 .550000 -841018 e820616 a054029 -001382 .550000 ... .e60551 .886454 .548860 .-.009,674 . .. 550000 ,880385 ,952293 a045291 -.024196 e550000 ad94238 1.0c0000 8049397 Q29867 625000 -679049 c.000000 a171 C18 .09 3296 625000 704285 .lo8449 195115 .GkIO310 ?b25000 .723Ul9 , e1923Ij8 .l~4300 .0723?3 e 625000 .743352 ,276327 e168901 .066220 002 50 e625103 ..762db5 e360267 e15 3665 e060600 625000 .782416 ,444206 el40201 -052424 a625000 .YO1952 -528145 .127Y78 e042931 625000 ... a21465 .612084 .1153za 0.31~11 625000 e841018 e690023 a137940 no56511 625000 .860551 e779962 .~11349 e.032387 b25000 .a 600 85 ,863901 .096604 e015018 625000 ,899618 -941840 ,091762 e005865 .625000 -911756 1. ococ0o .09160Y .001147 .. .70000~ - -.760-541 0.000000 .lb9125 .080066 7oocoo e7d2418 e129655 .210460 e067208 700000 .8019?2 245420 e197779 -079453 .ioooo3 ,8214t15 ,361185 ,181432 e070144 ,100000 e841018 476950 mlb5691 .Ob2329 ,960551 ~592715 e147346 ". ._ .*lOOOOP ." " 0 4.5 6.62 .700000 .880085 e7C6480 a131016 e028908 .700000 899 618 e824245 .148218 e04289b . .- - . .. 70wo__ -. .. ,919151 .9tO,OlO 14614U- .g 03.201 2. .100000 eV29273 1.000000 el46753 e027691 - -. 1 " e75000 0 ,797166e750000 0.000000 e021400e155210 .0~2704 -.06243+ -.u2385r 750000 ,821486 -169049 .PO6373 pa001704-e105798.lo057500~2178 750000 .BCLOIA .304831 m193639 -* 48 .OB5537 -. 02 -, 22 e7500 00 .e60551e750000 ,440612 -.016334,175541 .Ob8780 -.106761 .000497 .750000 e880085.750000 e516394 -.027170-157126 - .~ ._ 1.00 0000 ,9320731.000000 0.~00030 ,020245 .c34900 .035123 e034877 -.OS9889 e9 58218 ,384889 .9163161.000000 ,384889 e958218 a1~6719 .Ob1215-ab03359 -.045504 1.000000 ,977751 ,672452-.bo3127 ,096514 038773 -.05774~ -* 03 - 1.000000 997284 9 60016 .09em -.OZ~MI .021900-e076922 -.007513 1 .000000 1.0000901.000000 1.000000 ,021256-.025400.lo2811 -.oelsG -.iie?ii - ". ._ MINIMUM OF IC -c )= ,0094 AT 80.0000 PERCENTSEMISPAN AND 100~0000PERCENT CHORD P P IIPPER SURFACEIIPPER LIMIT ~~ MAXIMUM OF (C - LIMfTING C a00187 4T 30.~0PERCENT SEMISPAN AND 8.64 PERCEMT CHORD. m " ____ . GRA OIFNT GRADIENT GRAOIFNT 9 DELTAT = 2.707 SEC., T z3.m SEC. " .. SUMMARY OF PRFSSURELFUEL AND PRESSURE GRADIENT CONSTRAINT CYCLES MOST=TIC AL ___ MOST CRITICAL C GRADIENT DFLTA C P __- C P PLANFORMLOCATION INCREMEMTPLANFORlC LOCATION CYCLE M Y (POSITIVE 1.5 PERCENT)1 (NEGATIVt IN IS IIN PERCENT) NUMBER 0 E S4TISFACTORY)SPANWISECHOROYISE SATISFACTORY1 SPANLiISE CHORDVISE 1 .OlOOO m46140 e009442 a0.0000 100.0000 .oolen 30.~000 a.6380 ...... - SOLUTIONFORDESdGN C 9 .lOCOOP - -. . - -. . .. - - L WITH 1 CONSTRAINTS ONPRESSURE GRADIENT UITH C CONSTRAINED TO .010000 - " ...... 0 - ___ PLANFORM LOCATION OF SOLUTION PRES)URECOYSTRPBNTS f IGRADIENT AND o LEVEL) CHOROUISE __.- - -- ._" - SDANUISE "_ - (PERCtNT) (PERCENT) ..GRADIENT CONSTRAIYT AT 30.0000e.6380 AT Y 9 4.969AN0 X 0 130.R50~ 2 IS CONSTRAINED TO -4.070 AT Y 9 4.969 AND X 9 189.000, Z IS CONSTRAINEDTO -10.lb0 AT Y 9 4.969 AND X 0 243.390r 2 IS CONSTPAINEO TO -14.110 AT Y 6.625 AND X 0 189.0001 2 IS CONSThAINED TO -8.320 - A C-L I1 k. K "- '0 E ,191 2 3 4 5 *A ? 8 9 10 11 12 13 14 15 16 17 "- ~010000 ,468216 n338353 -e454679 ,251839 -e134100 .to6683 eA235bb -e042776 "172716 -e040738 -.024219 a008370 e010252 a005585 -e000389 ,022221 -e005343 -e008901 __ ". - - LAGRANGE NULTIPLIERS -.093930 .046614 -.00012e *000794 .ooom -.000231. .oooose .oo0040. ,~-.aoooos . -.cue791 ~- ." - CONFI6URATIONFORCE AYD NORENT BREAKDOUN C ~ ___ "- C C M . .' L r) 0 .... ___-"" "- UING .09Gb9 -003698 .003017 FIJSELAGE e00000 e000001 e003958 UINGINDUCED ON FUSELAGE-.O0931 .0@?806 .003025 " ..UIWG INDUCE0ON NACELLES 0.00000 0000177 0.~00000 .. ""-" """" """" ... TOTALS .IOOOO .COC~~Z .OIOOOO ~ "" ~ " _____"__ 969-500"" CHECK17 LOAD CAS! 22 S?AN STA.FUSELAGE YITH AND 2 TERMS - ...... ,...... SOLUTIONPRESSURE OISTRIBUTION LOW E R UPPER LOWER UPPER., Y- . 'X X-PRIME . TMICKNESSLIFTING SURFACE SURFACE SURFACE -. .-.____ - "e """ -...... - C .c , C, ;e OC. L DX .. - e 12 1 CHORD P P P .P.. P. ____..______,. ". e075000 ,081496 O~DOOOOO -.0921?3 -035300 . -.026767 ,065367 -e016963 .07'1000 ,008755 ,022113 -.r)A3946 .024509 -.O~J3141 .o1oeos -.013m .075000 ,lg8a ,047126 ~050334 ,e012303 .011503. =~03883_2 . .-*01093Z...... -___ 750 000 e9191S1,847956750000 -so47613e133726 ,017C36 -rll629U -.001870 I 750000 ,983738a939694 ,155842 -e054703 0034841 -0121001 -.OD1446 .750000 0941024 1.000000 a155U74 -.0555UO .033~ -.I~Ts~ 7.e~157+ .800000 e824147 0.000000 -.U13436-e063228 ,041568 ,028132 .041500 . "..800000 . .R~~Cl~.lSllllr a148856 ,024355 .093407 -.05-5448 -e012943 .800000 e860551 ,282927 ,180579 -no90464.090114 .004386 -.GO5422 .800000 e434729,880085 ,175466 -e002713-e105188.a070278-e014198 ___-.. .800000- 899618 586536 .16Ofib -.030480 -e113111,047675 -.001231 .800000,142513 ,738342 ,919151 -mu41961 a027373 -a115141 -.000222 .800000.1317'16 .938684 .890149 -.U51689 -a119317 -012429 -.GO1729 .800000 .95ze19 1.0oaouo-e001489-a123397 e006497 -.056700 ,129e04 900000 .878llO 0.000000 -1074724 ,045400 -e081259a073762-.000962 m900000,123800 -218797 ,899618 a029996 ,985293 . -e039507 -.COP156 "- - . ------. .- "______a 900006 ,919151 C! 7509 ,142244 -e068796,073448e007424 -.006536 e 93868 4 .616221 ,137765 -e019811 -044643 -.093122 -a093842-.093122 a9OOOOO-044643 -e019811 ,137765 .616221 e938684 ~ .9OVOOO 1959218 814933 .lt1253 -.OC2724 .014824 -.10642f -a003658 , .900000 e976410 1,000000e976410 .900000 mIL1135-a003391 -e118717e002417 -e055900 950_000 .. " ,905002 0~000000 -.110712 .046200 .-.018606-.09b914 I 09210C __ .Q50000 e919151 lbQl59 .OB1009 ,035582 -e014582-mol3269.067821 a950000 .404170e938664 .IZOltt .016132 ~069442 -.050685 -.Ob7386 .950000 .Q5821R ,639199 -_ ,119832 awl0747 -e034879-.076094,043738 ".- " - - .950000 I 977751 .E74220 -.039349.110815 .OlZi36 -.098678 -.007327 e9500 00 a988205e950000 1~000000 -.05680@,111397 -.OU4602 -.008170-.113998 ___" " 1.00 0000 ,9320731.000000 OIOOGOOO "143336 .nw~on -.046668 ,09666~ -.ilelaa 1.000000 .093331,384809 0958218 aU16316 e054371 -.038660 -.GO5931 1.oooooo e977751 ,672452 .0%3.851 -.~103127 ,037142 -.0029?7-.0564iI9 "" 1.000000 ,091394.960016 ,997284 -.022681 -.073208.018186 -.005630 1.000000-e005976-e076539,016239 -e025400 ,092779 1.000000 1.000000 MINIMUM OF (C -c 1. ,0136 AT so.row PERCENT SEMISPAN AND 1oo.oouo PERCENT CHORD P P UPPERSURFACE LIMIT MAXIMUM OF (C - LIMITING C 1- .Ob108 AT 7.50 PERCENTSEMISPAN AND 19.72 PERCkNTCHORD. P P .. .. GRADIE~~T GRADIENT OELTAT * 2.688 SECor T 0 25.966 SEC. - ._ "" ". .. SUMHA~~-FPRESSUPELEVELNDRESSUREGRAOIFNTCONSTRAINTCYCLES n0ST CRITICAL "-_ -. - . __ .. "" __" ___-______" .- -. --. -. -. nOST CRITICAL C GRADIENT dELTA C P C P _. ___ .. __-.____ PLANFORn LCCATI1~"""" INCREMENT PLPNFORn LUCATION CYCLE M K (POSITIVE IS ( IN PERCENT)INEGITIVE IS IIN PEIICENT) NUMBER 3 E SATISFACTORY)SPaYYISECHORDWISE SATISFACTORY) SPANYISE CHOROYISE __ ~ .~ ~ 1 -009442e46140.01060 n3.0030 ln(i.00r)F .001873 3dnbCbO 8.6380 2 m01000 ,46822 *0136U3 80.0000 100.0000 .001080 19.72027.500~ c 4 w **************************4********* SOLUTION FOR DESIGN C .100000 " WITH 2 C9NSTRAIHTS ON PRESSUREGRADIENT WITH C CONSTRATNED TO .01000C ___". n ...... 0 " ______"- PLANFORMLOCATION OF SOLUTIONDRESSURE CONSTRAINTS ( 2-CRADIEN1,AND 0 LEVEL) .. - ."G.hQIENtCO*(STRAINT AT 30.0000 8.6380 GRADIENTCCINSTRAINT AT 7.5000 19.7202 .OLpQOQ> -472958 e487378 -.425316 -189910 (;.140481 e190797,115615 -so30996 -e038986 ".lb9180 , ,-.ll295t -. .I. .ole252,004064 .ome5 -.oom 4 .003075 -a005353 -.00047b " LAGRAN6E.MULTIPLIERS -e095586,034491 eO55211i-.00@240 .PO0842 .OOU235 -e000235 .00005C oOOOO49;m000066 -, 64 " CONFIGURATION FORCE AND IOMENT BREAKDOWN _. -___ " - C C C M -__ L "-sL.-O ,. WIN6 -09063 ~003752 e003178 FmLIGE -000 QIL.0000319 """ " YING INDUCED ON FUSELAGE .00937 .om812 .002864 WINGINDUCE0 ON NACELLES 0.00000 .000165 0.000000 """_ """" """" -. TOTALS ..lOOOU 0004730 .010006 9 " " 969-5T 17 LOADCHECKCifE-ZY SPAN STATWITH FUSELAGE AND 2 TERMS "-___...... , PRESSURESOLUTION DISTRIBUTION^ , ...... I. 11., UPPER LOWER UPPFR Y . _. x X-PRIME TMICKNESSLIFTING SURFACESURFACE SURFACE . -" - ""-" C C C C DC I DX .... . st2 ,. L . C?@RD P P P P P J2-e-a ._ , .d. ;. I. .075000 .0914eb O~OOOOOO e010929 ,035300 sOZClb4 0313836 -e008219 ., __ e075000- -098755 a 13 ,036603 .[rZ4509 022133 "014469 " 95 . . " .. -. .", .. . .. .. - . " - . 750000 91 e91 51 ,847956,121752-.04?613 ,011459 -e110303 -.OD2189 750000 m93868k ,983738 -e054763,152031 ,032935 -.119096 .750000 .951021 l.O000(LII r~b"45500 032526 - 950 -7.003421 I - .800000 e825147 OmOOOOO'J -.339~+05elU07b5' .041500 ,14771~ -e015653 lBkiQLB.13U_.171793 .nooooo 104876 024355 -, " 0 000000 .dbO551 ,082355,004386,165059 ,282922 -a082705 -e003004 .800000 .d800RS-057525 -.01419U ,149961 ,434729 -.092435 -e002175 "--A!lOQppQ.-_-.949610 *36453&- .- ~~QE.".OJOSB.L".-.034533 -* 32 e919 141 ,730342 e120911 -.041961 e016571 -.lo4339 -.001101-.lo4339.800000 e016571 -.041961 e120911 ,730342 e919141 .800000-e051689 elr6033 ,890149 ,938684 m~O457L -,002521-a111461 " " BOOOOO -195t8 10 LQQUUL- "tlE@&---~&~ "1 19970 94 .900000 .078110 0.000000 e096638 .045400 -a018834-e011919 0084719 28 .9000 00 ,809618 ,2187 ,809618 .900000 99997"2lX!!iL .086861 -, 75 -. " e POOOOC ,517509,919151 ,119639 -e304904-a057493,062146 90 0000 -9386A4 e616221 .lo9115 -.U19811.lo9115e616221 -9386A4 900000 aU35318 -a078797 -.bOSU99 .9o9006.9~".RlS933"W9019-.04uz4" m77u -* 19 -. 54 .900000-e606472-e112178-.004122 -e045900 e108057 1.000000 ,976410 -"lppUo.aooooo ",mu A 046260 d73bQQ uo1n91 - 75h " I 950000 m919151 ,105961,169159 ,435402 .080256 -e025705 -e004845 e 950000 e038684 r104L30404179 ,016132 .n61429 -.J42071 -.uo5173 - .950099 "19511218,639L99 -.09%j& -a013747.&724 -~4~"-=~1~5079 ". - .950000-.039349 a091902 .675220 -917751 .OOZ680 -a089222 -.008U05 .9500 00 0988209 1.000~00 .lo01341.000~000988209.950000 -.OSbPOO -e011597-.110367-mol0233 ." "" - - -. - -- - .- " ." - ...... - 1.000000 .932073 o.oaoooo-e009574 -059574 ,034900 ,069148 -.027k20- 1.000000 .Q502181.000000 ,384089 ,016316e107535 -061623 -e045912 -.Ob2865 1.000003 ,977751 672452 .CepTp2--. -.uo7127 .04~~14~~91938~-.003105 ". . 1mOOOODO m0972q4 -e022681,104305 .960016 e324641 -.0075bl-e079665 1 .0000 00 1.000000 1.000000 e~086881.0000001.0000001.000000 -.02SkOO -e009116-e084492e024194 -_- __. ...- -. . . .- - " MINIMUH OF (C -c ). ,0161 AT 75.0000 PERCENTSEHISPAN AND 100.0000 PERCENTCHORD P P UPPERSURFACE LImIr _" - MAXIHUPI(C Of - LIIITIMG C 1. e00067 AT 55.00 PERCENTSEHISp4N AND 22.dl PERC€NT CHORD. P P - " - GRADIENT GQADIENT DFLTAT 2.698 SECer T = 26.604 SEC. SUMIARY~OF PRESSURELEVEL AN0 PRESSUREGRADIFNT CONSTRAINT CYCLES __ HOST C RITIC4L -~" MOST-CRITICAL C GRPDIENT I DELTA. C. P C - D PLANf ORM LOCATlONTNCREHEYT PLANFORMLOCATION CYCLE M K (POSITIVEIS IN PERCFNT) [HEGATIVE IS (IN PERCENT) NUMBER 0 E SATISFACTORY)SPANWISECHORDYlSE SATISFACTORY) SPANWISE CHOROYISE "- -~ - ___-~ A .Ol@OO mS6140 -309442 80.0000 1oo.cuoo ,001873 30.000~ 8.6380 2 .01000 .46@22 a013603 80.0000 .001080100.0000 ~9.72027.5t~0 . 3 .01000e47296"-.016050 75.000~100.0cr)b lrOU672 >5.9000 22.8567 _" _" ...... SOLUTION FOR DESIGN C = .100000 "_ L WITH 3 CONSTRAINTS ONPRESSURE GRADIENT YITH C CONSTRAINEDTO .010000 ~~ ti ...... 0 PLANFORMLOCATION OF SOLUTIONPRESSURECK?TRAINTS f 3 GRADIENT AND 0 LEVEL) " -. .-. -. .. - SPANWISECHORDYISE "" ... (PERCENT)(PERCENT) - - " GRADIENT CONSTRAINT 41 30.0000 8.6380 GRADIENT CONLTRAINT AT 7.5000 19.7202 GRADIENT CONSTRAINT AT 55.000b 22.8067 " ' .4$4800 ,213185 .0130OO -.1-454*1 e044701 -e073990 -.I55339 -~041060 e128217 -.02775?.134221 -e028291 . ., . &" . .. . -016580 e016252 ,005585 -moo4352 e000490 -a005353 .000810 i .2; ..!?,>*.. .. LAGRANPEMULTIPLIERS -.000694' 000233 -.000172 .oooa 75 .OOO" " -e104929 -.147483.281509 .395908 .001247 -a005317-e141623 .. A ND HOBENT BRFAKDOWN FORCF HOBENT CONFIGURATION AND " ., L D 0 WING .D9051 203912 .003217 " FIJSELAGE .OOOOO .OOOOOl a003958 YING INDUCE0 ON FUSELAGE e00946 .Ob0824 ,002825 " YINGINDUCED @I4 NACELLES 0.00000 .000212 0.000000 -""" """" """" TOTALS .1OOOU .00'1948 .010000 I ,. 969-500 17 LOADCHECK CASE 22 SPAN STA. WITHFUSELAGE AND 2 TERMS ...... SOLUTIONPRESSURE DISTRIBUTION ... . " " LOYER UPPER UPPER A/ . . !. ! ' X THZCNNESSLIFTINGY-PRIrlE'SURFACE 1 SURFACE :5URFACE _.__* 2.i 5 7:: 4 ... ., _. -r DC , ox . ':L.y?]-, . - - -""" .c c-, c c .. . 612 L CHORD P P P P P 075000 .oelqeb o.oooooo e074924 ,035300 e056762-.OlBl62-e004951 e074000 ,098755 .022113 ,076137 a024509 .043200 -r035537 -.(r055?9 "CIA* 11 0901 .n~-I 7k no- I os . n1 >ana .n76~2~ -.05&757 -.QO4467 1 .fnrOOO ..938b84.98?73A e12359b 4.994 IUS mUA0 I10 'i 10*048.".. -. ... boom="""- __ 750000 ,941024 1.000000 -123579a017579 -.055500 -e106002 -e001899 .800000 e824147 0.000000 ,109456 m041500 e389578 -a019878 -bo13494 .800000 .841018 e133076e131115 e024355,085518 -e041559 -.DO5596 . - 800000 -860551 ,282922 ,1344a4 .004366 .Ob7067 -a067417 -r004057 800000 .881)005 e434729 ,127827 -mol4198 .0464~ -.oe1368 -.00307o - .800000 899 b 18 e586536 ,118248 -.030480 .OZbkOb -e091842 -.LO2015 .800000 .73e342,919151 e105646 -e041961 e009039 -e096807 5.000837 .800000 -938684 .890149 .. e108641 -.C51b89 .00\)876 -.lOT765 -.003845 .800000 e9S2819 1.000000 e1140Q5 -e056700 -.OOlkO3 -a115497 -.001236 . 900000 e878110.900000 0.000000 .083804 .041400 -.005502.07n302 -.om43 "" 900000"" ,899618 - ,218797 ell2A44 -029996 -079815 -a033029 -.004509 900000 a919151 e417509 ,111909 e007424 ,658281 -.053628-.OJS549 900000 .938&84 eblb2Zl ,104697 -.019611 m028109 -e076588 -.LO4ZD6 e 900060 *95U218 e814933 ,092433 -e042724 -.092016.000417 -moo5752 e900000 e976410 1~000000 ,101239 -a055900 -.007531-.lo8769-.002287 -. m950000"____ m905092 0~000000 .Ob8169 .os9 .u70635 a002665 -.0179b0 -. . 950000 e919151 ,169159 ,097539 ,035482 .076045 -a021493 -.0052lZ e950000 ,938684 ,404179 e100089 ,016132 -059424 -.0406bb -.OD5646 950000 .958218 ,639199 ,093180 -e010747 .030412 -e062768 -e034781 'e950060 -977751 ,876220 .08POb8 -e039349 e000733 -e087275 -.008685 .9 50000 ,988205 l.O(r00~0,988205.950000 ,090474 -e056800 -.01SO63-.lo5537-.006582 , ___" -. 1.000000 e932073 0,OOOOUY.056750 e034900 ,053373 -e003375 -e021102 1.0 00000 .9S8218 .088Bll .OX6316 e052261.OX6316 .088Bll .9S8218 1.000000 ~384884 -e036550 -a003088 ___-1.000000 ,6724521977751 ,019809 -.003127 . e030421 -e049318 -e002550 ..... 1.000000 -997284 .010340.960016 -e022681 .011059 -a066681 -e004785 1.600000 1.000000 1.00bO~O .077822 -e025400 -.Ob9061.008761 -a003234 ______. - MINIMUH OF (C -c )- e0215 AT 80.0000 PERCENTSEHISPAN AN0 100.0000 PERCENTCHORD P P UPPERSURFACE LIMIT " HAXIHlJH OF (C - LIHIITYNG C I= e00073 AT 15.00 PERCfhTSEMISPIN AND 10.40 PERZENTCHORD. __-___--___P P - G RIOIEN T GRADIENT GRIOIENT DELTAT 0 2.727 5EC.n T . 31.39L SEC...... ". . .- " SUtlHARY OF PYkiSURELEVEL IN0 PRCSSUREGRADIENT CONSTRAINT CYCLES -...... " ... " ...... HOST CRITICIL M OST CRITICAL C GRADIENT C CRITICAL MOST "...... - ...... DELTA. C ...PL.j;N.~-~R"-. LocArI.oN- . IHCRE~ENTP...... " . -c -. - P *.PLANFbaMLOCATION CYCLE M K (POSITIVEIS ( IN PERCENT 1 .(NEGATIVEIS (IN PERCENT) . . " . - ...NUH.BE 2 -.O t SATISFACTORY) SPANdISECHORDWISE. .SATISFACTORY) SPANhISE CHOROHISE-,...... - ...... SOLUTION FJR OETIGN C - .100000 L - - ". "" "___ YITH 2 CONSTRAINT? ON PRESSUREGRADIENT YITH C CONSTRIINEDTU .010000 n ~ ~ ...... 0 - .. .. - ." "- - GRADIENTCONSTRAINT AT 7.5-0 ____ -. " .. - .. -- - 19.7202 " GRADIENTCONSTRAINT AT 55.OtiOO 22.8067 CFO ON NICELLES O.OOCO0 .UP0195 0.0000~0 """" """" .. . """_ - ___ " TOTALS GO4831 ,010C00 .. , .. . ___-__ -. 969-500 17 LOAD CHECK"K<~T~KN STA. YITHFUSELAGEAND z TERMS ...... ___" " i' . .800000 e952819 1~000000 ell6870-.056700;.: -rU0001(1 -.llb685 ,-rOOl541 . I .900000 - a678110 0.000000 -.023012 .O454OO e024894 -017906-mi55482 .900000 899618 .218797 el09703 , e020906 .r)18189 -e031503 'e007375 900000 e019151 e12C382e417509 ,007424 e062518-e057865 -.OO6056 "",938684.900000 -e019811,114578e616221 a033049-e981529-e004189 900000e958218 ,824933 e101682 -.OIL724 .005041 -a096641-e004050 900000 e976410 1.0000C10 ,103289-a055900 -.006505 -e109795 -.003141 " _" e 950000 m905092e950000 0.000000 -e049868 .046200 .Oil816 -061684 -st166083 950040 e919151 ,169159 e078~75 .0354RZ a066363 -.011811-.G10807 - -. -. -.950000- ._ e9386R4 .404173.lo2306 . e016132 .Ob0577-e041819 -e006656 950000 ,958218 ,099634-639199 -e010747 -.001022-.Ob5993e033639 .9 50000 ,917751.950000 -e039359,642731a874220 e003094 -e089636 -.007520 _" .98a205a950030 ._ 1.oooo~ .09382_6-.GO7880 -.OS6800-.lO7213 "013387 " 1.000090 ,932073 0.00000U -.~6b477 .034900 -.o~am a058238 -e080495 1.000000 m95R21R e384880 ,090808 -.0049jL-.037548,053259.016316 - 1.000000 e672452,977751 e090328 -.003127 -035b81 -e054647 -e003176 7 1;oooooo ,930016-997281 .088812 -a071917-e022681,016895 -.I205205 1..0000001.000000 z.o_ooooo m3895U7 . -.025400 .01+604-rJ74902 -.005141 MINIHUH OF [C -c 19 a0201 AT 80.0000 PEQCENT SEMISPAN AND 100~0000PERCENT CHORD P P - __ - -- -. ".. - ""- UP PER SURFACEUPPER LIT -HAXIMUM OF (C - LIMITING C 1. .nO000 AT 15.00 PERCENTSEHISPAN AND 13.20 PERCENT CHORD. P P G RADIEN T GRADIENT GRADIENT " DELTA1 0 2.722 SEC., T 0 3C.ll3.fEC~.- SUMNARY OF PRESSURELEVEL AND PRESSUREGRADIENT COHSlRAINT CYCLES ___-___ " HOSTCRITICAL MOST CRITICAL C GRADIENT ___ !E.CTd _c ______" P C P PLANPORMLOCATIONINCRCHENT PLANFORM LOCATION CYCLE M K lP9SITIVE IS (IN PERCENT)(NEGATIVE IS (IN PERCENT) E SATlSFACfORYJSPAkYISE SATISFACTORYJ SPANYISE CdORDUISE - NUMBER 0 "" __"____"___ CHORDUISE 1 eC1000 .013603.46140 PO.CGOD l~C.0fi~O .ci01080 19.72027.5UOO 100~0000 -" ~ r - -___75.0000 53.0000 2 2.8007 _. 3 .01000 a4R314 .Ot0115 eo.0000 ~uo.oooo .000uo1 1SeOOOO 13.2033 ...... ___ SOLUTIONFORDESIGN C .10VOOO ". . .___"_._____ c dITY 3 CONSTRAINTS ON DRESSUREGRAOIENT WTTH C CONSTRAINED TO .010000 M - ...... 0 ""_ ." "-. - - - "" - DLANFORMLCICATION OF SOLUTION PPESSURE CONSTRAINTS ( 3 GRADIENT A= 0 LEVEL) __ . - "SfWUX€CHORDWI S E -_. (PEPCENT)(PERCENT) - "- " ."_ " GRADIFNTCCINSIRAINTAT 7.5GOO 19.72D2 ____- - GRADIFNTCONSTRAINT AT 55.3FOn 22.8067 GRADIENTC3YSTRAINT AT 15.0000 13.2033 969-50017 LOADCHECK CASE 22 SPANSTA. dITH FUSELAGE AN0 2 TERMS -. .___.______" ...... SOLUTIONPRESSURE OISTRIBUTION .. *******P********************** . .i. . " " + " >' . ,.,.:. .1 . 'iouER * UPPER UPPER X X-PRIME LIFTING THICKNES'S SURFACE SURFACE'SURFACE I -...... " SUMHARY OF PRESSURELEVEL AN0 PRESSURE GRIOIENT CONSTRAINT CYCLES ' MOSTCRITICAL MOST ( IlTlCIL C GRIDIENT -. .. . -.- .- - - " ." .- - " 4 HOST ClITICAL C GRADIENT DELTA C P C P INCREkENT ___~. ~ PLANFORHLOCATIOM "- "_ " " LOCATIONPLPNFPRR CYCLE H K IPOSITIVF IS f1g PERCENT)(NEGATIVE IS (IN PERCENT) NUHRFR 0 F SATISFACT'lRY)SPANWISECtiOROYISE SATISFACTORY) SYINYISECHORDVISE - ". ." ". 1 .010- a46140 ,013603 80.0000 100.0000 .001080 7.5000 19;?2r 2 7'i.0000 10C.OCUO 55.0030 22.BOb7 3 e01000 e020115 100.~00 ."~____"___,68314 8O.OOOG .00~001 1!.203315.OOLtO " 4 .01030 a48502 .018510 7.5000 100eUJO3 -.001l116 35.JOUO ~R.4565 LARGESTVILUtS OF WING UPPcRSURFACE _. __" ." -" - """ ._ . LONGITUDIN4LPRESSURE GRAOIFNT DUE TOBODY BUOYANCY AND UPHISH LOADINGS, _____- . 4NO TO WING THICKNESSPRETSUR_ES _. -~___ .. " ..- PLANFORMLOCATION __ -" - - -. .- LARGEV"- (IN DERCENT) ~ " . .. GRAOIENTSSPANYISE CHOROUISE 19. 00 10.401 _____. e003106 19.00 ~. - - 15.00 13.20 - "_" .- _" - CAMEFRSURFACF CORRESPONDING TO OPTION 4 __"_"____969-500 17 LO40 CHtCK CASE 22 SPAN STA. UITH FUSELAGE AN0 2 TERMS SPANUISEOISTRIRUTION OF SECTIONORAGt LIFT, AN0PITCHING MOMEN7 ". - -- -. .- - - ."- "- - -_ Y SECTION SFCTION SECTION -" C C C CHORD 0 L ~.Ef 2 "" ___-H. .~ 0~0000000 183.9917000 0~0000000 -.OS51104mO609469 " . .. .- .. - ."*0250004.. ..- "- -, !~?*?~~~~~-.OI-~.OO_ORO~.-,0635613 -.0597203 - ~______.0500000172.0073528 0~0000000 -e0668628.Ob08513 .0750GOO 166.0731792 -.0776302.0777107 .3073%1 ,0 769 02 7 de0774657,0769027 ' . __" . _*ruoooqc" 16 0~?30600 -* !@?.a 3.5 2". -- - .1 25 00 00 154alQ51R5Q.1250000 .3044110 -.07R8723.07579d5 .1 500 00 0 148.25nh941.1500000 .004066Q -.080>416-0750135 ".. .17?4.?PI). ~142_.3~~~~.".oos~712"" -0753591-.08404b3 .2 000 00 0 136.3933123 .0047023 a0762417 -e0892392a0762417136.3933123.0047023 .2000000 a25C000n -.lL25623e5704822126.6106675,0047331 .- - .. - - .3oooooo_ ~13~3324744 ~053.322 ,0344334 -.A262182 ____ - "...... 3 50 000 0 103.2602613.3500000e0055635 -.15662@0.@90SllA .4000000 " . . 4 7 5OOOG. .5sOOOOO .I5!5.33¶l4 - ." ." .. .bZSOOUO.-- - .7000000 7500009 "" -_.8~00000~ .QOOOOOO .9500000 1.0000000 ~ ~. . SPANWISEINTEGRATION 9Y TRbPAZilID RIJLF = 311.572109 FOR KODT,L!IAONO(KOPTI, KVAk 18 0 2 Y Y Y Fl VI Y FiY) Y F(Y) I.!! . .- "- _. -. . "- ".-. - - . . 0.00000 44.26773 42,a3s42 ,7974541.40910 1.1961839.979iP 1.5949038.55347 1 .993 63 37.12099 2.39236 35.69184 2.7910835.69184 2.39236 37.12099 1.99363 34.7636029.99861 3.98726 32.83537 3.18981 ~s.le41L3-Lr_4_1Pez___ri*>.!?Z !L..L?*.W-fi 6.3196222523le5 15,188626.77197 18.46383 7.57560 9.9t1815 11.91341 11.16433 9.63819 11.961787.36478 12.759246.58732 14.35414>e03241 15.15150 4.25495 15.94904 3.47749 ____ ~ "" ~ - _.______- - - - -.- - " SPANYISE INTEGRATION BY TRAPAZOIORULE = 8584.760860 FOR KOPTrLOAONOIKOPT)~ KVnR = 18 0 3 Y F(Y) Y F(Y) Y F(Y) Y F(Y) Y Fly) ,398 73 l'l35.130021714,71362,39873,797451598.383111.190181486.13239..-. 1.59490 _, o.oOo60!_1959a11 ?IL" . "_ ____ "." .~ "" 1.99363 1377.96822 2,302361273.907552.791081173.99455 3.183811078.16126 3.98726899.92852 4.79471 752.39516 5.58217 618~0411~ 496.9L05Q340.912996.379627.57560 8.7719723~.b9410 11.16433 74.61840 11.9b~7854.23998 12.7592443.39282 "9.96E11_..141,929?1? -~ " . L4".2b?32513.. - 15.1515918.10461 15194904 12.09296 SPANYISEINTEGRATION 8Y~-TRAP4ZOXO RULE m 25.932097 KOPT,LOADNO(.KOPTI,FUR KVAR m 18 0 4 .. - " ___ Y . Fly) ,Y F(Y) Y *(Y) Y F(Y) Y Fly) 0.00000 2.69708 2.72287,39873 ,79145 2.951J72.94922 1.59490 3.1068b 1.19618 _"___ 1.99363 2.01.349 .2.67737 2.39236 2.7910R 2.~59207 3.18981 2.50342 3.98726 2>.3?31 4 .7847 1 2.31599 5.5R217 2.250235.5R2172.315994.78471 b.37962 2.21098 7.57580 2.04>54 8.77197la90548 9 .968 15 1.677259.96815 1.1395811.91178 1.37150 11.1b433 ,5238014.3541+ .88b37 12.75924 15 .151 59 e33313 15.94904 .2951b15.94904 e33313 15.15159 "_ __" - . . SPANUIStINTEGRATION EV TRAPIZOIO PIJLE 1.099969 FORKOPT,LOAONOIKOPTJ, KVAR 18 0 5 Y F(Y) Y Y F(Y) Y F(Y1 Y FtY1 Y F(YI . 0.00000 0.00000 .39873 0.00000 ,19745 0.00000 1.19618. -29556 a.59~0 ,20569 16374 2.39236 ,17728 2.79108 16690 3.18981 15440 3.98726 15440 3.18981 16690 2.79108 "l.&5m3- ,17728 2.39236 16374 4.78471 e14626 elA0146.37962 .13631 5.58217 7.57580 "e09387 e058228.77397 9.96815 .01413 11.16433 -e03301 11.96178 -.GO604 14.35414-e0063712.75924 -e00535 15.15159 ,00533 15.94904 - ,009Q2 - ___ -. . SPANUISEINTEGRATION BY TRAPAZOID RULE 9 -1205.475049 FORKOPTt LOAONO(KOPT)# KVAR 18 0 b .- Y F(Y) - Y"- F(Y) Y F y "" "x . 0~00000-107.99616 ,39873 -109.99460 ,79745-1li.6;L1.19618 -124.08284 1.59490-115.12475 1.99363-108,68349 2.39230 -102.60252 -98.672262.79108-96.2~4503.18981 3.98726-92.29871 4.78471 -94&59h 5.-94&59h-99.66148 4.78471 7.57580m .-~95~uws~-." e. 3 796? -m'1141 8.71197 -97.6b969 .- 9.96815 -91.52644 11.16433 -78.99073 11.9617~-67.1~68~ 12.75924-53.14722 14.35414-32.49718 15.15159 -21.10084 15.94904 -18.91b89 ._ " . " -. X CP c 9 .100020 C2""- = 70~m4 K= 4~59'14 L 0 L E s 2-.___" .- ~ " ~. - REF M ----- .920569 -- = -.057282 C . .OOOW3 """I- - .- .. . -n - .- "_ . PR OG 'L 0 "" " CONFIGURATIONFrlRCE ANdMOMENT BREPKOOHN " C C M L 0 0 " ". YIYG ,003835.n9041 .OU~~RS FUSELAGE .OOOOO .003Q58.000001 WINGINOUCEO ON FUSELAGE e00961 00833 002750 ".. - WINGINOUCEO ONNACELLES O.OOOC0 .000193 0.00QOU0 """_ """" """" ,004862 .MI9993TOTALS .lOOOZ ,004862 " .. LO YER UPPFR LOYER ." 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' . 7000 0.000002.58+202.473612.39403 2.24904 1.97211 1.53291 ,86225 ,45458 2.60034 2.52261 2.32170 - __ I I -" PUNCHEO OROINATE>.YAVEBEFNR€QUE$EDA"!LY!JR9WZSE AN0 S.PANUISE LOCkUW4P 0 F OROUATES 4RE PUNCHLD FIRST. . AN IRAGE OFTHE PUNCHED DECK FOLLnb'S. . 069-500 17 LO.~"lPALlU.XIH€lLiELAGLA"L".LLtI~- " 0.000 5.000 LO.OOO 2o.ono 30.000 +u.ooo 50.000 60.000 70.000 80.~~3 90~000109~000 "QQ.QQ~.QO"LQQO 7.500 ~O~QQO~UA.~AUQ~~ 30.000 35.005 4OeOOO 47.500 55.000 62.500 70.000 75.000 80.000 QO.006 9~~000100~000 ".UQ -O_.pM. _.Q~QQ~~.-OLQQL".!~OQO0.000 0.000 0- 0.000 0.000 0.000 0.000 o.ooo O.OUO n.ooo o.noo 0.000 0.000 0.000 0.000 -QdM"- -QdM"- "_ "___ 0.000 0.000 0.000 0.000 0.000 o.(rou O.OJ0 0.000 0.000 0.000 0.000 0.000 0.000 ,161 -182 -a624 -dm198 -I&h-4.783 -5.668 -0.255 -6.b75 -7.132 -8.500 0.006 e004 -e102 "799 -1.887 -3.009 -*.GO1 -5.808 -5.397 -5.777 - _.-6.161 -6e85L. 0.000 -e005 -e098 -e707 -1.644 -2.632 -3.523 "4.251 -4.779 -5.085 -5,343 -5.742 -.-~.OOO -e039 -.le$. -e879 -1.870 -&936 -3.920 -4.754 -5.393 -5.807 -6.132 -6. 500 0.000 "032 -.175 -a838 -1.799 -2.833 -3.801 -4.640 -5.284 -5.702 -5.974 -b.2R2 - "____ OeOOL1 -e017 -e134 -.749 -1.657 -2.649 -3.592 -4.407 -5.036 -5.434 .- 0.000 ,456 ,862 1.533 1.972 2.2492.394 ZeC74 2.584 2.bOO UPCITEC UINC OEFIWITIPh .. - ...... WING CAPeEPSURFACE PEIC INTO BASIC GECPFTRY ~ ." **** - **** **** **+* **v* - -yaG"- - **** .-*3*7-- ..". . . -. . . .- .. - PFRCENT HALF-Tk 'ICKNFS2 CAPPER CHORD IlPPCP LCUFP 17) ---.0,CI-- 0.6090 -- O.OGOO c.coc0 2.5 .'700 .570C a0032 5 -0 .?I 40 ,7140 ,0064 -. 19.C .Q??O . .P72C -.1E3? 20.0 1.0500 1.0500 -1.2795 I .14r0 -3.C217 ... . 3G.6 1.1650 1.1560 40 10 1.2000 l.ZOGO " -4.E184 X 2TSU"- 50.6 1 .?300 1.2300 -6.4069 1.2lr5O 60.0 1.2490 1.2490 -Tat992 1,2270 .. 70 e0 1.1700 1 e170C -P.&474 1.1270 "- 80.0 93'0 .937c -9.29c4 .a830 90.0 .546C e5460 -5.et5e -5070 300.0 - 0. ocoo 0-.0004 - -mi-Y8ff .- " . ." P€RCENT CAMeER HALF-THICKNESS -. -.". c)rMtr;- f UPPEW- " -.tnyR . .- 0.G c.cco0 G.ooac ~.COO~ 2.5 .a735 .570~ .97oc 53- . .1471 -727C .12lC 10.0 .2?RS .QOZC .902c a9110 20.0 .121e 1.0980 l.CS@O "c-- -.2601 r.z?cc 1.zzcc 40-0 -.1@7C 1.2840 1.zwo 50.G -1.3FZF 1.31501.315c .. 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"...... - ...... __- 'PCAAFFPP aPFAKVOTNTS -. I X Y 7 CHCK'D AUX. CHOPC XLE XTF AUX XTF O.OOL0 166.C7CC 1t6.07011 0.0000 166.0700 1C6.0700 0 .ocoo 1Wilt30 lEC.1330.- O.OOCO 149,7500149.7900 0.0c00 12C.3500 125.35CO "0;OCtJ - 77.245G O.@CO@ I6 0.ocoo '0 0.0000 - TT" 19 .o __ - "" ...... '7 .4 .7 r !2 !4 " - -&RE K"". z-"" 0.0cc00 1o.ocooo 23.50000 8.55CGO 57.500CO 7.10000 . p9.0COOO 5 64000 117.00COF 4.17COC . .73coo .. - .. 126.00G00 . 119.ROC00 1.28000 108.00c00 -.14CGO ... -. . - 104.0ot00 -1 .bCOOO 107.00000 -3.04L3C .. ". -- . - ...... "- 107.0ooao -4 e500GO 106~00000 -5.90000 102 -7.46tOO .oucoo ...... 94.00COC -P.BXOO . 79 .OOOOO -1C.25L00 59.00000 ?3.00GOO 8.OOCCO -14.6CQ00 0.00000 -15.7COCC ...... -, - ... ...... -...... -- . - . "- "" ". ". HODIlrJNTdLTAIL PLINFCRM X Y 2 CPORO 1 261.000@ 7.6000 -14.0COO Z3.CGOO 1 260.3B@9 2at.ooca ."" - . 2 277.OCOi- 11.0006 -14.0000 9.LQOO 2 263.?333 2Pt.COCO 3 2PC.CCGC2CC.277P ...... - . - ... - 4 269.2222 2@6.(iOCC "3- 272.16t7 ZPCDCCCE e 275.1331 2FC.COOC 1 278.C556 2RC.COCO V!NC r)CVNYAW AT TAIL SHIFTED PER C-B INTSCTN h) 0 0 XPCT 0 .oo 5 .oo 1c.ac 2C.OG 30.00 40.00 50.00 6O.CC 7C.00 eo.oo -. ." ." .... " ...... "" __ .- .- ... QO.OC 100.00 - 1/812 .- .. ., .. -...... -. ... -.... ._ -" U.OOOG o.ooonn -7.13200 e00400 -.lC7OC -.79900 -1.8a7co -6.9 ?ROO .- -.00500 -.09POO -1.644Ob -5.74200 ...... - ...... -.03900 -.195CC -1eP7tCO -E.50000 ..... 17% -.OS7CO -.I7500 -.s~eco P700 -6.2...... " .zoo0 -.01700 -.1?4OC -.749CC -..Q?POO ...... -. . - ...... _ 2500 -.625CG -, ..- ...... 3000 -.56zca .HOC, o.oomn .~P~OCI .c4ooc -.43200 -*.072')r, -6.4C900 - . - - -- .- ...... 04000 O.OEOOO 13700 el7600 -.105CO -4.94000 -5.43100 .- " ...... ,4750 0.nCOOC .?llOC -4.99oor .. - ...... 55cc o.oooon .532OC ,556OC -3.15LCO ...... e6250 0.000n0 ,71900 lmOE7CC -.Q170? ."" ...... ". ,7000 0.00000 .RIZOC 1.53300 9.57300 ...... a7500 0.00006 .01200 "_ "_ .. - - ...... CHCRII C .OC 5 000 . "" . .. 1o.ar 20.00 36.Ot 4o.oc 50.uc 6G.OC _. . .. 7GeDO 00.00 90 -00 .lGG .OO 20.00 30.00 40.06 6C.CO 7C.CP eo.co 90.~0 100.00. ." ..."_ -~ .?07 -I bF4 -1.145. SO7 -.CF4 -1.145 4.46t 3.753 3.47t 1.9?F 3.456 3.258 3.438 3.321 3.3ee 2.4PP 2.412 2.430 1.761 1.693 1.tR7 1.298 1.255 1.247 .9P3 ,958 a951 .6OS .; cc3 'eEC2 ,412 -414 ,417 .294 e299 e304 .ne ..nn.- - .27f .l$S .. - a173 ,179 ,134 -137 -153 .top- - -.111 117 -090 ,092 a096 " -064 .Of 5 ,667 .0ZT - ;045' - .c47 .026 ,027 ,028 -...... 021 .ole .GI6 ._ ...... " - -.It4 -.It4 .74t .689 e719 .497 1345 ,253 .IS1 .I20 .ce3 .Cbl .c45 .c35 .C?I e034 .a34 .O?f .e34 ...... - ... -, - -...... oze .CZF .c2e *02F .CZB .bZR 0023 .C14 eC14 -.011 ~900 .01C .00q .... -...... ooo .OOP .cca .. ...... -. .. - ..... N w0 ,797 ,776 -7501.0c0 ,776 ,797 -- ,634 .t6C ,696 .to1 e5h7 .500 - -. ..- ." " " ." - - ." ,723 ,533 TNCPE~ENTAL FltSELAGE UPYbSH ON TAIL PEP CEGREE ALPI-b - . .. XPCT 0.00 10.00 7O.CO 30.00 4C.CO 50.00 6C.CO 7c.cc 8C.CO Tim-.. 100.00 11812 - 0.000 -.Q97 .07R 1.czt 3.321 C.C25 5.899 5.5441.1t3 2.465 4.143 .3sn - "_ - " .1CO -.P97 .OfR 1.526 3.321-4.62TT.899 5.544 4.143 -T;7;Fc lelt3 .351 .ZCOlsC4t 1.369 l.Z?P 1.t7E le9271.7392.034 1.970 1.b03 -746 0486 ,300 1.074 1.1161,106 1.086 1.L34 e972 .eel ,714 .le9 ,468 0356 400 .750 7r-.m ,672 ,633 .589 ,541 .47C ,401 336 .v .500 a530 508 .4?5 .4r9 a432 ,403 0 ?74 .?44 .304 02t4 2?6 .66G .?PC ea72 .q.4 .??6 e31P e 299 .ZEC ,261 ,242 ,219 0194 _. .-..- 0 297 285 .nY" .249 247 .20e e19t .1€4 a171 .eo0 .2?S 226 217 ,200 -199 190 .IC1 ,173 ~164 -156 a147 90 0 e190 .1P4 .17e e172 a166 -159 a166 e172 .17e .1P4 e190 900 . -152 ,157 .I42 ,196 -130 .. 1.000 .7 3*4- 137,149 a141 -144 em- .129 125 e120 0136 2130420000 20R65000 .i443t4 2CC.234617 0.coaooo 2 14029 6d00 2.917145 2140296d00 26.734131 .0443t4 206.979940 .071716 215.m 2.96P253 27.67- oC41510 ZC7.M .C@PthC 2 16004W O O 3.0233C9 2P.716025 3.0233C9 216004WOO .C44?22 ~CP.~C~~SC .a19467 I t4 0 ul _" _" NACELLE PRESSURE FIELO Y1012 XDPER CENT CHORD AN0 PRESSURE COEFFICIENT ,100 83.104151.692,100 '"1.702 232.424 232.146 233.e68 234.551233.912 2?6.035 23t.757237.479238.201 23P.023 2?9.64'240.367. -24!.090 241.A12 242.534 243.25t 243.?74 -~ ". C.00094.60C 44.14992.69893.247 92.79C 02.790 95mC51 95.502 96.84596.40495.953 97.906 07.751 98.208 [email protected] i I I ...... "G.COOOC ...... omaoo- " ...... -. .. "...... " ...... - ..- -...... -. ._ ...... e50 240.997 7C5.130 "..- " . -- . - ". . -~ C.Cb0 Ima" N L 0 - --mm-" 0.00 5.0 10.00 20.N7.00 4a .a0 50.0u tU.OU 7@.@@ 80.00 00.00 100.00 0.000 04 ,00236 .02R9? .0703P s15C .09?b7 eOP615 00734 09650 .LC144 elGI42 .c~34 cc71p 0 ,02345 .Ob137 .I75 .lo114 009504 09508009504 .lo114 .I75 .1a343 010242 10590 .OFT64 o06602 a03683 .oltra--- .01?0l a03303 .200 e0llbO -02679 ,223 .I r'17 lUY36011606 01lC90 e11027 .lotee .CnOC sOt546 .03594 -01136 .07141 " 250 llq6Z 11 601 .12 11741 2110 .11c14 .00904 .Ol?lR ,275 .lZEZO12901 el2591 ,17477 1264C , .I 343 .0$563 .e6790 .03m e01446 .OOPOE 001354 I , .3co -425 -15741 .1s511 ,15452 ,15627 ~15255 el3509 .11e47el3509 ~15255 ,15627 ,15452 .1s511 -15741 -425 .OPt15 eOE46C e04093 .we116 ,07191 ,451, .15@?5 .I9935 ,159e4 ,16246 e15763 -14430 el2362 .c971e a06994 mO4CtO 1. "_ e03560 *02'60 "___ - .475 .16000 e16397 .le574 el6715 a16235 e14534 .1z7eC .lCltC a01627 e05696 e04470 .ow46 " 500 el6544 ,16674a16598 .16991 ,16790 .~mo.1o5se .13C@5 .0PtS2. a06438 e05761 e04514 -___-.- . ____ "" . .525 a16677 .17lt5 e172P4.1L404 e16774 .17353 e07195.0?94F e11173 .I 3444 -06049 e05434 - e540 -17777.08011 .09$57.1?727 .1172e n13914 e17100a15022 a16705 ,17124 - .ObPV ""*??" "-1- - ".___ "_ - " . .. . .17197 ,17327 ,17336 .16a45 .15ES? e14155 ,12217 e10324 .wn12e10324 ,12217 ,575e14155 .15ES? .16a4501 hqC?,17336 ,17327 .17197 .or771 .07041 .or771 " e6CO *I6730 el7141 e1729t -17359 m1651Sa12723 el4451 .1552e m11073 e09715 aOP715 eOP136 . -, . ___ "" - - -. - - e625 elh7'JC ,16824 e10724elt9lt.11o12 e13356 .iae9r-14736 .15ee4 a16626 a09703 aOCt54 " __.__ ,650 elbl7C q15949,164.re-16485 ,16701 -16650 e1saq5 el4019 e12w1 e11013 .10@?3 e09707 " _I__ "" .- 675 e15613 675 e15PbR.It052.16310 ,16325 ,16063 a12859,13905 a14847 .155C4 .ll*Sa e10906 "- " 7UO ,157'2 .1lrqqo m16200elt4Y3 e16597 .16510 oICC72 .15425 a14'40 e13519 ,12467 a10767 " ___". .. . e725 e16641alCS11 ,14424.I5203 .Itt73 e16765-1t359 .lob91 r1682e e13555 e750 e15336 ,15332 .1532*e15141 .15317 .I4872 ,14491 .UP73 -13367 e775 *136?4 ,13164 -13871 .13954 -14077 e14085 e139tT el3106 e13353 a12913 .17 ~0 1 ,120~9.17~01 1 .8bQ ,12111 .l?JOQ a12546 ,12755 e12924 e12571 m12924 .12soe -12.71 -12216 ,11754 __.11010 - " .e25 eliI~02 .tlO41 ,11180 -11452 ,1166~ all791 .1te~7 e11778 .me4 all535 ellt6C a1093C .a50 econ3p .owon .0995@ .lot78 .lo220 e10290 e10503 alC594 elOC10 ;iom - -.7 F9Pl ,10551 .e75 eobrlf *OR440 .OR56R .c~eo7 e09029 .OS?24 .CCkC@ .C917C aOPi47, aO9bt7 .09PPC .lOOd4 - TLBLE OF FLbT PLLTF CP bl 1 DFG &i@LC Cf ITTACK XPc f 0 .bo e.. 10.tO 20.00 30.00 4C.00 50.CC t6.m 70.00 Po .a0 QO.00 100.00 .. .., .. 0 360 .ne' 8 e0432P,03983 .C19?5 03141 . 2007 aOZC17 . 14J4 ,018401931 .O174P .01hOO " 0125 e06168 o0464Z 03426 ,02142 a02035 mtu em5 .U eC1961 .QIrn"" .O17BP e01650 .425 a06F56 .o557e002218 .02e21 -04335 .02243 e02173eC2195 .020es 002c11 001961 e018S5 " e450 ,0791 4 ,65946 ,03116,04552 e02304 e02300 .a2275 ,021~ .OPlC5 .02CC2 .O?OlC .aloelr ,475 e07914 -0631'1 ,34987 e03293 .Ut408 .a2388 .0?320 oCt212 .O2150 a02174 e@2@52 e01964 -500 . ,07454 ,06357 .05m e03604 e02598 .02419 .E2257.C71?3 .OZP2@ e02155 .020?4 .0?001 ____- .. - .. .. 0525 .OF139 .ohm ,03967,05577 .0280? ,02408 .am3 me2337 .c21er-02267 .O216l e02153 - . .. 0550 eO7PR1 ,09778-06827 eC4Gbl .02578 .C24C6 .a24w .C2404 e02304 .022C@ e02266 .022Q5 _. "...... I -09 498 -08937 -06'347 .a5361 -04626 .C3975 -01489 .C3975 -04626 .a5361e725 -06'347el07'19 -08937 -09498 ,07966 m032C2 .0310@ e03123 -75 0 .09504 -750 e09078 e08653 e06659..07733 .am1 .05a13 e03793~ .a4377 e03455 ,0326'J .031?5' " ". - .774 .09?99 ,09084 e08624e07028 -07843 ectn3a e03777.a54e3 e04304 .a4873 e03572 e03367 ',* ,824 .OR279 .OM 41 ,07415 e08053 e07848 .am3 .0t3@SeC51t5:' eC4726 eO46tZ .ob?on .owe .e50 .07P1?.0?541 e07687 .07759 005541-07303.C6042 eOt511 .0t549 e04091 ..0475q .a4421 __- "" "~ ___ "" -07467 ,07450 e07384 ,01289 eC7t69 .Of791 .Ct415.Of791 e875eC7t69 ,01289 .t?74F4e07384 ,07450 -07467 mq22 -05578 ,0*09q ,04433 " 900 .070?3 e07069 . e07C93a07116,070P4 .07c5t? .at et5 .C6639 .059@l006319 969-500 17 LOAD CHECKCASF ?7 SPAN STI. WITH FLSELAG1ND 2 TERMS PICHNUHEEP = z.7oac HOR?ZI?NTAL TAILCONTRIBUTJON EXCLUDED F0-S F0-S .I IYTFPFEPENCE ORLF COFFFICIENTS __"" TNTERFERENCEDRAG COEFFICIEPTS CC = .10294943F-02 CD slSP37749E-02 - - NACFLL€PRE*WRFS ON FLbT SURFACE FLATYING PRFSSURES ON KLCELLE _" CO = ,676rniP .us e001684 .00617 e00422 .0016a9 .COIJ~~.oc14a7 -06 .co7ma ~30530 e002167 e00335 .OC2281 aOOllE7 e07 .C07@@? .00442 e002772 C024P oc31as ,002973 .oa .PO1951 .003S9a00161 .003504 .CC4055 .OC39C6 e09 ,004427 ,304362 .a0074 e005132 aO049t5 .??!?6R ~ - .lo e005430 .001a1 .00534@ -.OOU14 eTC633.6 .OCC152 .ll eOOh5'iQ .00093 e0064SF -e00101 aCC16h7.0014CI .12 .007R16 .o(ron6 ,0076qa -.oolee .co= .00€!90s -.000A1 .iJrl90.6 3 -.GO215 .ClO7@9 .CICC71 __ -_ CONFITUSATICN STPEAWUISE LIFT OISTPT6uTC.- BlfIC LIFT DISTRTBUTION INCREWJT PEP bEGPft ALPHA X Wll v-+c NLC TAIL suv w-e-c Tin su t .0140" ,000t7 OmOOOCO OmGOCOO 4.154 a00067 e00047 a00067 -" 4.154 "7.COGrn O.OOOC0 .COO47 e" e" m- .mTn--"-;m-rv;mr-.0011T" 12.462 -04224 .002e? 0.00000 o.ooooo ,002e5 .c0200 c .000CO .o02co 16 .615 ,0563) 16.615 ,00529 0.OOOOC U.06000.00301 e00429 0.00000 . m00301 2U .769 e07040 ,00579e070402U.769 0.0000c 0.000CV"- :6U5TTp -0040t a.ocaac .c0w6 24.923 .on449 e00732 0.0000C O.0OOCO e00732 mOGSl2 0.000CO ,00512 29 .077 ,0985729.077 eOOP89 0.00000 OmOOCCG .00t21.aoee9 C.00000 .GO621 33.231 . ,11265 . 1043 0.00000 o.oooot- .0107.ocne C.OOCCO .oa72e 37 .385 .12673 a01195 .12673 37.385 e011q5 0.00000 0.0000G .0~e33 o.ooooo .ooe33 41.53 9 a14081 ,01354a14081 41.539 0.00000 0.OOOGO 000535001344 O.OOOG0 rC0935 45.692 .l';sea45.692 -01496 0.00000 0.0000c e01037-01496 a.ooaoo .u1az7 49.846 .IOPI)' .0144@ 0.00006 O.OOOC0 ,01134a01646 O.OOCO0 001135 . 5s.000 .IVOS mol795 0.00000 O~OCOCO e01234,91795 0.0OOCO e01235 5@ . 1545@. a1-01973971? O.OOOOC"~.OCOO~~"- "927;01332 0.00000 ~ITZ!" 62.308 e02026 .?lltl O.OGOOO O.OCOO0 -01412e02026 0.00000 .O1412 66 .462 e22529 ~02110e22529 66.462 0.00000 0.00000 e01479.02110 0.00000 -01475 70.616 .7-9-7 e02175 OmCOOOO O.OC-60 -01332e02175 O.OO6CC 0015P2 74 e769 74 .7536+ e02275 0.OOOOC OeOCOCO -02224 eOl5tC C.00000 eG15tt 78 e923 .2E754 e02907.2E754 e923 78 0.UOJOO 0.00600 002307 ~01605 0e000CO oOlbC5 83.077 *m62 .UFiW"O~- o.ocoaD-~5"".Pleae 0.00000 oblbce e7 e 231~29570 eO282P O.OCU00 0.0GOCO.ole37 .02e2e O.OOOCO .ole37 91 .3e5 .3097?91.3e5 ,03305 0.00000 0.00000 009305 m02054 0.00000 eCZO44 95.539 .ow7 o.ooao(1 0.0000C . 99.692 :K e04625 b.OUOOC 0.0000C e04625 ;ozeeo O.OOOCO .ozaeo 103 .e46 ,35202 .0*410,35202 103.e46 0.00006 O.C000(i - 005410 0034tt C.00CO0 .034tt 10P .000 3 6 bI0 0 6 Z3T"T.DoVvT) 0.mmC" "ZtFzm .04T3@ 0.000C'o r0413F 112.154 a3ROlP ,07104 0.00000 0.00000 e07104 .04e91 o.ooooo .CS~PI 116e30e e39426 e08109 0.00000 0~00000.o5esl .oelo9 o.ooooo .om1 120.+62 .4097' oq171 o.ooo(1c o.oGoco 2 124.616 .1ozeoe47243 o.ooooo o.ucoao .102eo .oeoo~ o.ocooo .ococl 128.769 043691 e1146P 0.OOOOC O.OCOOGC.00000 ,04250 - ,11468 .O925C 132.923 I4903S -.xZ-e3T.- moo0 0 .oocm- .lZ891 0- 137.077 04f467 .143ee o.ooooc O.OOOCC .14m :z;i :::",:: .E7- 141 a231 ,47075 .1+1)95,47075 a231 141 OrOOOOt 0.00000 ,13742~16095 OeOCCCO e13752 145.385 es9m .19r0.Olj~r o.ac(rcc 18043 .1>495 u1.00000 .1.4SF 149.539 ,50697 -70177 0.00000 OeOC000 e20177 -17353 0.0OG00, 017353 153.693 .570QQ e27466 OICOOOO 0.0OOC0 a19265 ,22566 0.00000 e192t5 157.856 ,53907 24898 --V;JT~.ToOaP- a24C9C 162 .000 -54915162.000 .27538 0.0000c O.OGO0O e21538 :::E :::::x: ::::E 166.154 e5hqlS .qozn1 o.ooooo o.oooco -30281 eZ5C94 G.00dC0 ,25654 170.30e ,57732170.30e 3 m!-? 0.OOQOC a.ooooo e27513e33127 0.CO000 .27971 174 .462 ,59140174.462 *'6145 0.0000C OeOGOCC ,36145 ,36394 0.00000 e30354 17P.616 -.6654p ,39269 OmOOOOL 0.00000 ,39269 e3251C O.OOOC0 03291t "" - m.no .fmh ,~~--~u---"~.*ZSIO n'54L4 O.CCCC0 .394C4 l8C.923 .4satl,63364 o.oaooc O.OCOOO e45621 .?ec32 O.OOCCO .t80:2 191 .077 064772191.077 .CUqflO 0.00000 0.00000 .4eeeo .46e32 o.oaGco .408x 195.231 bbl m@ 57107 0.0000c 0.OCCCC e43CBI o*ooocu 043614 199.385 ,675PP .5?275 0.0050C 0.000c0 -55275 -4644C O.OOOC0 -46440 293.539 . .6wf .m15 - o.ooooo o.oroco .5e41s .WM c.occca .49355 201.693 e70404 --.r"7C I! *" 211.847 .71e12 :,"::,; ::::E X:8% ,64411 .Y~I O.OOOQO .ma1 21C.UW e67336,73220 0.00000 O.QC000 -67336 .sew o.occco .mn 220.154 076679 ,701 36 224.30~ .'I6097 .77V6 .005@3 0.00u00 .?3479 oCW44 9.00000 .e4944 228.462 .774b? .7361R -01608.7361R .774b? 228.462 OoOOOOO ~71246 .6em O.OOOOO e60373 232.61 6 .7A959 .7R312 e02t49.7R312 .7A959 232.616 0.00000 .'e91161 e72214 0.00000 -72214 236.710 aQ0261 .n1m -03791 Oe00000 -7t47t.e5314 OmOOOCO a7647C . 240.924 .rj4511 .0466€ .rj4511 240.924 eQl6hQ 0.0c000 ,89279 .BC457 0.000CO .e0997 245.077 .R7449,93077 e05234 OeOGOCC a92684 .M~OC- o.ooooo el23305 249.231 .P44P7 .e9646 e05371.e9646 .P44P7 249.231 0.oooao ,88744-9501P o.oaaao *E8144 253.385 .P~EIQF .Ql¶86 91047 ,97111 o.coaoo ,52047 257.539 -67301 .9?1156 -67301 257.539 .9 e eo9 .94101 0.000CO e55101 261.693 -94231 m99932 ,97670 m99932 -94231 261.693 .~e7oq .05?01:x x:::::: 0.00000 0.00000 a97hIO 2651047 .OOllP .947?5 .05701 0 .OGUUO 1.00486 .95481 0 .oaaco ,994P3 270.001 .94941.91526 ,05701 OmOCOCO 1.003921.00642 O.OOOGd 1.00352 274 .154 .929*b .94921 -05701.94921 .929*b 274.154 O.COOUO 1.0041t1.00322 O.OOOC0 1.00416 278 .308 .94342 ,94675 e05701 ,94675 .94342 278.308 O.CU0007U3761.CO24b 1.0029t o.oocoa 282 .462 ,95740 .94*53 e03701.94*53 ,95740 282.462 OmOCOOC 1.00244 r.oc1eC O.OOOCO 1.001~9 286 .616 e9715P ,94441 e05701,94441 e9715P 286.616 0.OGOOC 1.00142 1.00109 0.00000 1eC0109 290.7711 .9P5ibt 05701.9c360 0 .OOQOO 1.00061 1.00C4t 0.000CO 1.0004t 294 i924 .Q9974 .94100 .05701.94100 .Q9974 294i924 o.oaaac 1.00001 l.OOCO1 OeCOCOC 1oC00Gl 295 .OOG 1-0OOOP .94PQQ - - ,05701 0.c0000 1.00000 1.OCCOC~ c.00000 1.000cc 969-50017 LOA@ CHCCC CAfF 22 SPIN STA. YITH FUSELICE AND 2 TERMS PACNLUCBFR m 2.7000 HORIZONTAL TAIL ALPHA. 2aCC0 TYTERFERENCF DRAG COEFFICIENTS FLAT UIVG PQEFSUPFf ON CIMqFPEb SURFACECAPBERED UIKC PCESSURCS CN FLAT SURFACE Dm 9 E-02 CD a1574 E-C2 8 CMXBAR WITH WAC = e000631 -f -09S714 -CL)t -.102881) FOR CL 0. CMIBIR a01047e N N 0 "" FIGUR~EAMYISCLIFT OmRIBUTTCN BASIC LIFT OISTRIBUTIOH INCRSMEHT PEI OEERFF ALPHL Y XI1 U-R-T NAC TAIL SUP u-8-c Ti11 sur 4.154 .0140P - .OOOhP 0.OGOOC 0.00000 e00068 eOO04t 0.00000 m0004t P, 308 ,00164 .OZFlc. o.ooo[lr - -0mcI- "XQl-64"- mrn 0 .mw 00112 12 .4h 2 ,04224 12.4h2 .O@ZR6 0.00000 O.OC0OC ,00286 e00195 0.00000 eo0195 o .ooaoc o .woo -.ao2s5 O.OCOO0 OOOZSII O.OOO0G 14;%1)75T .UU391 o maaa eCt397 24.923 .00736.On449 0.09000 O.OC000 ,00736 moo502 O.OOOCO ,00502 29.077 .09R57 .OOPQ4 o.ocooc 0.OCOCO "..OOBW ~ .aocae~..~.~ 6.ooooo~ ~~~ a006CE~~ ". -.~ 41.644. - - 3 3.23 1 ,11265 .0104Tp,11265 33.231 K"- ~.~ ~ ~.~ o.oooc(i-~...~... 13 O.0OOCO a00713 3'1.3 85 .1?67S .01202.1?67S 3'1.385 0.00000 0.000CO ,01202 .ooelrr O.OCOCO eooe15 4 1.53 9 .I4041 .01357.I4041 41.539 0.00000 O.OCOC0 e01352 .0051C 0.00000 e0091C 45.692 .I -190 01 504 c.m?m"T.~u---- ,015OL .a 017 o .oooao 4 9.84 6 .lC8Q' .01657.lC8Q' 49.846 0.00000 0.000cc ,01115,01657 C.00000 eC1115 54.000 .OlpO4.! 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Sommer, Simon C. andShort, Barbara J.: Free Flight Measurements of Turbulent BoundaryLayer Skin Friction in the Presence of SevereAerodynamic Heating at Mach Numbers from 2.8 to 7.9. NACA TN 3391, 1955. 2. Craidon,Charlotte: Description of a Digital ComputerProgram for Airplane Configuration Plots. NASA TM X-2074, 1970, 3. Harris, Roy V,, Jr.: An Analysisand Correlation of Aircraft Wave Drag. NASA TM X-947, 1964. 4. Carlson,Harry W. and Eliddleton, Wilbur D.: A Numerical Method for the Design of Camber Surfacesof Supersonic Wings with ArbitraryPlanforms. NASA TN D-2341, 1964, 5. Sorrells,Russell and Miller, David S,: Numerical Method €or Cesign of Minimum CragSupsrsonic Wing Camber with Constraints on Pitching Moment andSurface Deformation. NASA TN D-7097, 1972. 25 1 APPENDIX A INSEFiACTIVE GRAPHICS The cathoderay tuke (CRT) displayana program coding for the design and analysissystem are based on the NASA-LRC CRT and associatedsoftware. However, all displaypcrtions of the system codingare subroutined or overlaid from the basicprograms, so that tne systemcould be readilyconverted to other CRT arrangements. The basicinput parameter required toactivate the graghics routines is the executivecard CRT (punched in columns 1-3) , which may he placed at the beginning of the datadeck, or anywhere else in the inputthat an executivecard may be read, If the CRT card doesnot appear in thedata deck, no graFhic displays will be generated. The CRT card is actuallyan on-off device, Successive readings of the CRT card either turn on the graphics, cr turn them off and placean end-of-file mark on the hard-copy file, dependingupon the Freviousstatus of the graphicsroutines, However, the usual mode of graphics operation is to place a CRT card at the beginning of the datadeck, if graphics are desired, BASIC CRT OPERATION Several types of video displays are generated by the design and analysissystem, using the NASA-LRC software. These include: 0 Menus A list of disFlay choices with CorresFonding function keys 0 Edittables A list of numhers with variable names 0 Plots Displaysof x-y plots When a display is complete, one of two systemxressages will appear atthe top of thevideo screen, If the display is a menu, the message AWAITING OPERATOR ACTION will apFear, To continue prccessing,the user must press a definedfunction key, selected frcm the nenu. The secondsystem message is FLOT FRAME COMPLETE. When this messageappears, the graphicssoftware is programmed to 253 allow (a) editing of the displayvariables, (bj resumption of programexecution, or (c) hard copyplot generation- If (b) is selected,the user presses function key 3 (NEXT FRAME) - Editing andhard copy options are discussed on pages 254 and 255. Menus Menus consist of a set of display choices, tcgetherwith defined function keys, Some menu linesdisplay sets of functionkeys. For instance, a menu line may say FK KEY 6 DISPLAYS WING THICKNESS. Pressingkey 6 will bring UF a seccnd menu, with the message FN KEYS 1 THRU 20 IDENTIFY AIRFOIX NUMBER, whichwould require the user to select one of the inputairfoils. I% should be noted that the upperkey number (20) is the maximum number of airfcils allowed in the input. For a particular case, however, theuser may haveinput o~ly7 airfoils. If theuser now presses an undefinedfunction key (8 thru 54) the message AWAITING OPESATOR ACTION will appearand another function button must be chosen. Edit Tables If the displaycontains an edit table, the user maynow usethe consolekeyboard to typein a new value for anyvariable in the table. The variable name used on the display is first typed, follcwed by anequal sign and the new value,followed by the consclekeys FETURN and E0M. (e.g,, CONSTR (3) = -1.0 will change the third value of the array CBNSTR to -1-0)- The new value will be displayed at the top left of the video screer,, Tne LRC software allows the definition of onlyone edit format per display in theusing program, It canhappen that there are both fixed andfloating point numbers on the screen to be edited. If this happens, the edit formatcan be changed by the typed-in message F'0RMAT = XXX RETURN E0M (where XXX is thedesired format). This format remaim in effect so long as thedisplay is up, i.e., untilkey 3 is pressed. In case of doubt,the display will identify the current format if the message FQRMAT = RETURN EOM is typed. Special Usage of Key 55 Function key 55 is used in two ways. If the statement "RESUMES EXECUTION" appears on the menu line andkey 55 is selected,the currentgraphic program will beterminated and execution will continue at the next executable statement encountered- If the statement *IDISPLAYS PROGRAM (ZIPTION MENUII appears on the menu line, 254 and key 55 is selected,the current menu will be erasedand the previous menu redisplayed. Hard-Copy Plots Each time the systemmessage "PLOT FRANE COMPLETE" appears on the display screen, the user has the option of generatingVarian hard- COFY Flots of thecurrent display, assuming the run terminates normallyand the job controlcards specify the correct post- processor.Selecting key 6 (RECORD PLOT) or key 8 (RECORD PICTURE) followedby key 3 (NEXT FRAME) will savethe display informationand continue processing. GkAPHICS USAGE The principaluses ot thegraFhics routines ir? thedesign and analysissystem are to display the configuration,edit input geometry,and to displayand/or alter the basicprogram calculations. There is no Frovision in the system to alter the inputdata stream on-line, so theintended usage of the graph'icsand the inpt data card set UF mustbe carefully coordinated. Limited capability to redirect the systemcalculation sequence is availableand these opticns are displayedon the CRT screen when ercountered. Geometry Configurationgeorcetry may be displayedeither from the PLOT module, or, ir. simplified form, from thegecxretry module. The PLCT displaydraws a picture of the configurationon the screen (as instructed by theinput view cards) , Dut has no edit caFability. All editing of geometrymust be performed inthe geometrymodule. When thegeometry module is enteredfrom the executive to read or changeconfiguration geometry (executive cards GEOM, GEQM NEW, FSUP or WGUP) , the CET program DISGEQM is used to displayand/or edit the configurationgeometry. The first Kenu generatedgives the user the oFtion of executing or bypassing the videodisplays: FN KEY 1 DXSFLAY AND EDITS GEOMETRY FN KEY 55 RESUNE EXECUTIOK When key 1 is selected, the programoption menu appears: LPN KEY 1 DISPLAYS COKFIGURATION PLANFORM 255 FN KEY 2 DISPLAYS FUSELAGEAREA VS X FN KEY 3 DISPLAYS WING CAMBER (5 vs X) FN KEY 4 DISPLAYS WING CAMBER (33 vs Y) FN KEY 5 DISPLAYS WING CAMBFJ? (5/C vs Y) FN KEY 6 DISPLAYS WING THICKNESS (SIC vs X/C) FN KEY 7 DISPLAYSFUSELAGE SECTIONS (NON-CIRCULAR) FN KEY 8 ECITSCONFIGURATION CODES FN KEY 9 EDITS PERCENT CHORD ARRAY FN KEY 10 EDITS X,Y,% AND CHORD (AILRFQILS 1-10) FN KEY 11 EDITS X,Y,% ANC CHORD (AIRE'0ILS 11-20) PN KEY 12 ECITIDISPLAY WING T. E. (TBBRD) FN KEY 13 EDITiDISPLAY WING T.E. (T550RD + %LE) FN KEY 14 ELITIDISPLAY WING THICKNESS (ZiC VS X/C) PN KEY 15 EDITSFUSELAGE X =RAY FN KEY 16 EDITS FUSELAGE 55 ARRAY FN KEY 17 EDITSFUSELAGE AREA AERAY FN KEY 18 EDITSX,Y,E AND G OF NACELLES FN KEY 19 EDITS NACELLE X ARMY Ffi KEY 20 ECITS NACELLE R ARRAY FN KEY 21 EDITS X,Y,B AND CHCRD OF FIN AIRFOILS FN KEY 22 EDITS X,Y,B AND CHORD OF CANARD AIREOILS FN XEY 23 EDITS CAMBER ARRAY FN KEY 55 RESUMESEXECUTION The table below describes the tunction of each key. "-KEY """_FUNCTION 1 Aview plan of configurationthe geometry is displayed. 2. A Flot of fuselage area versusstaticn is displayed. 3. Given 'an airfoil number 1 through 20 (1 beingmost inboard] a side viewplot of camber (cambervalue + E of leadingedge) versus station at the Y of the specified airfoil is displayed. 4. Given a percentchord number 1 through 21 (1 at leading edge), a rear viewplot of camber (camber value + E of leadingedge) versus Y atthe percent chord specified, is displayed. 5. Same as key 4 but camber valueversus Y 6. Givenairfoilan number 1 through 20 (1 being most inboard), a side viewplot of airfoilhalf thickness (upperand lower) versuspercent chord at the specified airfoil, is displayed.The array cf thicknesses (THK) is displayed below the plot and may beedited by the 256 user. THK (1) representsthe half thickness at the leading edge. 7. Giventhe tuselage segment number 1 through 4, andthe sectior. number 1 through 30 withir. the segment, the Y and E coordinates defiRing the fuselage half -section are displayed. Tne horizontal X axis is positioned vertically at thefuselage centerline Z value (EFUS). 8. Thekasic geometry input parameters 50 through XBARIN are displayed on thescreen and nay be editedby the usIr. Theprogram defined format is 14.. If it is necessary to modifyvariables REFA, CBAR or XBARIN the user must firstchange the format to floatingpoint, such as F8.4. 9. Thepercent chord array (XAF) is disFlayea on the screen and may he edited by the user, 1011 1 Fourarrays, XLED, YLED, ELED and CLED representingthe X, Y and E coordinatesof the input airfoil locations of the wingleading edge and the airfoil chord lengths are displayed on thescreen and may be edited by theuser. Key 10 displayscoordinates of first 10 airfoilsand key I1 the last 10 airfoils. 12/13Keys 12and 13 provide a specialcaFability to remove ttspikeslt or irregularities in the King camber surface, A plot of camberline E values (from array WZ0RD) or E + e versus Y alongthe wing trailingedge is displayed. The correspondingtable of E or Z + ELE values is displayedin a tableunder the plot, which may be edited. When the NEXT FRAME key is depressed,the following menu appears: FN KEYS 1 THRU 21 DISPLAY PERCENT CHORE LINES TRAILING EDGE MAY BE ECITED FN KEY 33 TWISTS WING TO MATCH EDITED TOE, FN KEY 3U 5ESTARTS WITH ORIGINAL CAMBER CEFINITION FN KEY 44 SAVES NEW CAMBER DEFINITION FN KEY 55 DISPLAYS PRQGRAM OPTION MENU 1-2 1 A plot of B or !3 + BLE versus Y at the percent chord selected is displayed. 21 The E or B + ELE array is displayed below the plot and may be edited. 33 If the trailing edge has been edited, the remainder of the camber surfacedefinition is altered, by 257 linear twist, to agreewith the trailing edge change.The trailing edge is redisplayed. 34 Restart option, If the change to thetrailing edge was made incorrectly,the original camber surface may be recalledand the editing redone. (The restart option is availableuntil key 44 is depressed) -44 The wingcamber surface, WEeIiD, whichwas altered in a scratcharray until ncw, is permanently changed to match thesurface displayed under key 33. 55 Return to redisplaycoxrpleto option menu. 14 Given an airfoil number 1 through 20, a side viewplot of airfoil thicknessversus percent chord is disFlayed. Thethickness array of the specifiedairfoil is also displayed beltzw the plct and may be edited. 15 Given a fuselagesegment number 1 through 4, thearray of fuselage X valuesfor the segment are displayedand may beedited. 16 Given a fuselagesegment number 1 through 4, the array of fuselage e valuesfor the segment are displayedand may he edited. 17 Given a fuselagesegment number 1 through 4, thearray (A) fuselageof area values forthesegment are displayed and may be edited. 18 Four arrays, X, Y, E and D, representingthe coordinates of thenacelle origins are displayedand may beedited, 19 Given a nacelle number 1 through 9, thearray of nacelle X coordinates are displayedand may be edited. 20 Given a nacelle number 1 through 9, thearray (R) of nacelle radii values are displayed and may be edited. 21 Given a fin number 1 through 6, thevariables XL, YL, E L, CL, XU, YU, HJ and CU, representing the X,Y,P, and chordlengths of tne lower andupper fin airfoils are displayed and may he edited. 22 Given a canard number 1 or 2, the variables XI, YI, BI, CI, XO, YO, EO and CO, representing the X,Y,E andchord lengthsof the inboardand outboard canard airfoils are displayed and may be edited. 23 Givenan airfoil number 1 through 20, thearray (C) of camber valuesfor the airfoil are displayedand may be edited. SkinFriction Module/Near-Field Wave CragModule When the skinfriction proqram executes,the force coefficient surrmary from theprogram may be seen,-orbypassed, according to the menu below: FN KEY 1 DISPLAYS SKIN FRICTION RESULTS 25 8 FN KEY 55 RESUMES EXECUTION Sirilarily,function keys 1 and 55 display or bypass the summary results from the near field program when it executes, Far-Field Wave Drag Module When the far-f ield wave dragprogram executes, the menu choice of display or bypass first comes up. If display (FN key 1) is selected, the displayprogram (DIS080) will give the userthe opticn ofgenerating displays as follows: FN KEY 1 EEASES SCREEN FN KEY 2 DISPLAYS GRID FN KEY 3 CISPLAYS BODY AREA VS X FN KEY 4 DISPLAYS OPTIMUM BODY AREA VS X FN KEY 5 CISPLAYS CONFIG AKEA VS X FN KEY 6 DISPLAYSkESTFtAINED CONFIG AREA VS X FN KEY 7 INTERRUPT PROGRAM TO ALIDW HARE COPY PLOTGENERATION FN KEY 8 DISPLAYS FAR FIELD WAVE DRAG SUMMATION FN KEY 55 RESUMESEXECUTIDN The user's options at thispoint are differerrtfrom the other disFlays. Here theuser constructs the plct to include as many curves as desired, with or without a grid, , and may or may not generate a .hardcopy plot. To v,iew theconfiguration area plot, the user needonly select, key 5 (xollcwed by key 1 to remove the plot).If the user wants a hardcopy plot cf all curves with a grid,he selects keys 2,3,4,5,6 and 7 followed by keys 6 or 8 and 3 (NEXT FRAME). He may thenresume executicn, display the drag summation or build a new display after erasing the current display with key 1. If the user selects key 8, the menu is erasedand the wave drag prcgramdrag summary is printed(illustrated by typicalvalues): 70 CHARACTER TITLE ARRAY FOR CURRENTCASE CASE= 14 MACH= 2.700 NX.= 50 NTHETA= 36 WXNG VOLUME CHECK EXAC T V0LUNE EXACT = 11432.023 EQUIVALENT BQDY VOLUME = 11429.954 ENTIRE AIKCKAFT D/Q = 20.27199 CDW = .00263 0PT. CDW* = .002481 DRAG CF THANSFERREC Al?@ DISTRIBUTIONS 259 OPTIMUM EQ- B0CY CDW* = ,00089225 AVERAGE EQ. B0CY CDW* = -00104445 POTENTIAL CDW* CHANGE = -.00015220 At this point, the system message PLOT FRAMECBMPLETE will appear. To get a hardcopy plot of the display, press key 6 or 8. To continue, the user selects key 3 which erases the screen and re-displays the functionkey menu. NOTE: There is oneinstance when the wave dragdisplay subroutine will not be called. That is when the restraintpoints exceed allcwable storage of 33, which causes the optimization calculations to be omitted- Wing CesignModule The graphicscapability of the wing design prcgram consists of: 0 display of lvbucketggplot, drag-due-tc-lift factor (KE) versus Cmo- Q KE versus Cmo for camber surfaceccmtraint solutions, if requested 0 Editing of thedesign solution variakles (Cmo, CL) and constraint or restart codes 0 Continuation to next input case or return to executive Thedesign camber surface, which is automaticallystored in common block CAMBER, can be viewed in the geometry module, but not in the wingdesign rodule. The initialdisplay to appearin the wingdesign module is the byFass or display menu: FN KEY 1 DISPLAYS BUCKET PLOT FN KEY 55 RESUMES EXECUTION When key 1 is sElected, the optimum drag-due-to-liftversus Cmo 81t=ucketg1 plot is displayed.Additional symbols are also plotted, givingthe flat wing (.+I, uniformload (x) , andthree term ( A ) solutions. (Theuniform load andthroe term solutions will be plotted only if those solutionshave been calculated). 2 60 After the bucket plot is generated, the NEXT FRAME keybeings up the set of currentdesign inputs: 70 CHARACTER TITLE OF CURRENT CASE CMO = -0200 CLDEIN = ,1000 RESTART = 2.0000 CfZNSTR(1) = 1.0000 C0NSTR(2) = 1.0000 CeNSTR(3) = 1 .OOOO CaNSTR(4) = 1.0000 Theuser may edit any of the variables on the display. If editing is performed, the wing design case may then be re-executed when the NEXT FRAME key is againdepressed, which generates the menu: FN KEY 1 EXECUTES NEXT CASE FN KEY 55 CALCULATES EEITEDDESIGN POINT If key 55 is selected, the programreturns to the wing optimization overlay, andrecalculates the wingdesign for the edited designinputs. If key 1 is selected, the programcontinues to the next statemer,t in the normal execution Frocess. When the wing design case is completed,and key 1 is selected, a final cption menu is displayed: FN KEY I TEEMINATES WING DESIGN PROGRAM EXECUTION FN KEY 55 kEADS NEXT DATA CASE The Furpose of this choice is to permit the user tc abort a series of wing design input cases once the desired wing design has been obtained. LiftAnalysis Module GraFhicsoptions provided in the analysismodule consist of: 261 0 Displayand editing of wing twist array 0 Editing of configurationangle of attack,and canard and horizontal tail setting (if used) 0 Editing of Mach numbzr,and inputs SYNM,WHUP, and ANYBIZID 0 Display of wingpressure coefficients andfuselage upwash 0 Display of forcecoefficient summary The initial menu seen is: FNKEY 1 DISPLAYS WING TWIST (DEG) VEXSUS SPAN FK KEY 2 EDITS WING ThIST ARRAY FN KEY 3 EDITS CANARCANGLES OF ATTACK FN KEY 4 EDITS SY", WHUP and ANYB0D FNKEY 55 RESUMES EXECUTION The user selects the function key associated withthe task desired, noting the following conditions: 1. Iffunction key 1 is selectedand nc twist array was input, no plot will be generated,and the user will be required to select anotherfunction Key. 2. If functiorrkey 2 is selected, thevariable TWXSTN (the current number of twist anglesin the array ATWIST) and the ATWIST array are displayed. If Entries axe added or deleted in ATWIST, a correspondingchange must be made in TW-ISTN. 3. If functionkey 3 is selected, thevariable ALPN (the current number of canardangles of attack ir, array TCA) andthe TCA array are displayed, If entries are added or deletedin TCA, a corres,mndingchange must be made to ALPN. When key 55 is selected,the analysis module ccntinues execution, halting with the menu, FN 'KEY 1 DISPLAYS UPWASH VERSUS PERCENT CHORD FN KEY 2 DISPLAYS UPWASH VEFtSUS PERCENT SEMI-SPAN FN KEY 3 DISPLAYS WING PRESSURE VEESUS FEIiCENT CHORD FN KEY 4 DISPLAYS WING PRESSURE VERSUS FERCENT SEMI-SPAN FN KEY 55 RESUMES EXECUTION whichFrovides the displayoptions indicated. Selection ot' keys 1 through 4 brings UF cne of the following secondary menus: FN KEYS 1 THRU 21 IDENTIFYSEMI-SPAN PERCENT FN KEY 55 DISPLAYS PROGRAM OPTION MENU 262 .. ... " . ._". . . . ._ .. .. .- . . .".- EN KEYS 1 THRU 11 IDENTIFY PERCENT CHBRC FN KEY 55 DISPLAYS PROGRAM OPTION MENU FN KEYS 1 THRU 41 IDENTIFY SEMI-SPAN PERCENT FN KEY 55 DISPLAYS PROGRAM OPTION MENU FN KEYS 1 THkU 20 IDENTIFY PERCENT CHOEC FN KEY 55 DISPLAYS PROGRAM OPTION MENU If no fuselage was input,function keys 1 or 2 will produce no respcnse. Key 55 returns to the primary menu, Upcn resumption of theanalysis calculations, program FINISH is entered which halts with the menu: FN KEY 1 DISPLAY DRAG DUE TO LIFT PROGRAK RESULTS ANC EDIT NEXT HORIZONTAL TAILANGLE FN KEY 55 RESUMES EXECUTION If key 1 is selected, the drag summary table is printed (illustrated with typical values) : 70 CHARACTER TITLE FOIi CURRENT CASE PAACH NUMBER = 2.70 CONFIGUEATION ALPHA = 0.00 CARAkDALPHA = 0.00 HORIZONTAL TAIL ALPHA = 0.00 CL CC (OFF) CM (OFF) Ci) (ON) CP (ON) 00 -000551 - OOSO 1 .000645 -00598 .01 -00045 1 .00677 -000525 -00474 .02 -000480 .0055k -000533 .00351. . . . . 18 -018392 -.01424 .018121 -01627 -19 -020603 -.01548 -020311 .01751 20 ,022942 -.01671 -022639 -01874 NEXT HCFIZONTALTAIL ALPHA (TFALP) = 1.50 Ir! the table, the (otf) and (on) reter to nacelles.Canard alpha and hcrizontal tail alphas are not printedif I?O canard or horizcntal tail is present. 263 It is possibie to trim theconfiguration by the properselection of hcrizontal tail angle. Ifthere will be anotherhorizontal tail angle, its value is indicated as shown,The value may be edited by tyFing THALP = XXX EETUEN E0M. Key 3 (NEXT FRAME) will then resumeexecution. A broader editing capability for altering thecalculation seqaence is enabledby the next menus to appear, The Frimary menu sets up the choice: FN KEY 7 ALLOWS USER TO VIEW AND EDIT MACH NUMBER, CONFIGURATION ALPHA AND CANARD ALPHA, FOR CUFGEN'I AND NEXT EXECUTIONCYClE FN KEY 55 RESUMES EXECUTXON Selectionof key 1 displaysthe current Mach nurrtber, configuration alpha andcanard -alpha, In addition, it displays the next parameter in the cycle to change,which may I=E editea by the user (tyFical values are shown): CURI;ENT MACH KUMBER = 2.14 CUkl?F.NT CONFIGURATION ALPHA = 7: 70 CUERENT CAKAEC ALPHA = .50 NEXT CANAED IiLPiIA (CAN) = .95 - or - NEXT COKFIGURATION ALPHA (C08) = 1-87 - or - NEXT MACH KUMBEE (XMCH) = 2.30 The Frogranexecutior? sequence is canardalpha loop, configuration alFha 100~ ana Mach number loop, in that order, When the individual loops are complete, tne words CURRENT and NEXT are replaced with LAST. The user has the option of editir,g the variables CAN, CQN and XMCH when theyaFpear on the screen. 264 NASA-Langley, 1976 CR-2716