NDARC NASA Design and Analysis of Rotorcraft - Input, Appendix 2

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NASA/TP – 2015 - 21 8751

NDARC

NASA Design and Analysis of Rotorcraft

Input

Appendix 2 Release 1.10 March 2016 Wayne Johnson NASA Ames Research Center, Moffett Field, CA

March 20 1 6

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NASA/TP – 2015 - 21 8751

NDARC

NASA Design and Analysis of Rotorcraft

Input

Wayne Johnson NASA Ames Research Center, Moffett Field, CA

National Aeronautics and

Space Administration

Ames Research Center

Moffett Field, California 94035 - 1000

March 201 6

Available from: NASA Cente r for AeroSpace Information Nati onal Technical Information Service 7115 Standard Drive 5285 Port Royal Road Hanover, MD 21076 - 1320 Springfield, VA 22161 (301) 621 - 0390 (703) 487 - 4650 13 21 25 31 32 37 38 44 45 46 47 48 51 54 59 60 67 76 77 Contents . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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OffParam Performance PerfParam MapEngine MapAero FltCond Mission MissParam MissSeg FltState FltAircraft Data Structures and Input Input Based on Configuration Parameters and Constants Job Design Cases Size SizeParam OffDesign 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16. 17. 18. 19. 20.

ii 94 95 96 102 105 106 108 110 113 116 119 123 125 127 129 145 147 154 156 157 161 164 165 167 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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FltFuse FltGear FltRotor FltWing FltTail FltTank FltProp FltEngn FltJet FltChrg Solution Cost CostCTM Emissions Aircraft XAircraft Systems WFltCont WDeIce Fuselage AFuse WFuse LandingGear AGear Contents 21. 22. 23. 24. 25. 26. 27. 28. 29. 30. 31. 32. 33. 34. 35. 36. 37. 38. 39. 40. 41. 42. 43. 44.

iii 168 169 185 188 191 193 195 197 199 208 211 212 214 218 220 221 224 226 231 232 241 242 243 249 250 251 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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WGear Rotor PRotorInd PRotorPro PRotorTab DRotor IRotor WRotor Wing AWing WWing WWingTR Tail ATail Wtail FuelTank WTank Propulsion WDrive EngineGroup DEngSys WEngSys JetGroup DJetSys WJetSys ChargeGroup Contents 45. 46. 47. 48. 49. 50. 51. 52. 53. 54. 55. 56. 57. 58. 59. 60. 61. 62. 63. 64. 65. 66. 67. 68. 69. 70.

iv 256 257 258 262 265 267 271 273 276 279 281 283 285 287 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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DChrgSys WChrgSys EngineModel EngineParam EngineTable RecipModel CompressorModel MotorModel JetModel FuelCellModel SolarCellModel BatteryModel Location Weight Contents 71. 72. 73. 74. 75. 76. 77. 78. 79. 80. 81. 82. 83. 84.

Overview Chapter 1 Data Structures and Input 1–1 The NDARC code performs design and analysis tasks. The design task involves sizing the rotorcraft to satisfy specified design conditions and missions. The analysis tasks can include off-design mission performance analysis, flight performance calculation for point operating conditions, and generation of subsystem or component performance maps. Figure 1-1 illustrates the tasks. The principal tasks (sizing, mission analysis, flight performance analysis) are shown in the figure as boxes with heavy borders. Heavy arrows show control of subordinate tasks. The aircraft description (figure 1-1) consists of all the information, input and derived, that defines the aircraft. The aircraft consists of a set of components, including fuselage, rotors, wings, tails, and propulsion. This information can be the result of the sizing task; can come entirely from input, for a fixed model; or can come from the sizing task in a previous case or previous job. The aircraft description information is available to all tasks and all solutions (indicated by light arrows). The sizing task determines the dimensions, power, and weight of a rotorcraft that can perform a specified set of design conditions and missions. The aircraft size is characterized by parameters such as design gross weight, weight empty, rotor radius, and engine power available. The relations between dimensions, power, and weight generally require an iterative solution. From the design flight conditions and missions, the task can determine the total engine power or the rotor radius (or both power and radius can be fixed), as well as the design gross weight, maximum takeoff weight, drive system torque limit, and fuel tank capacity. For each propulsion group, the engine power or the rotor radius can be sized. Missions are defined for the sizing task, and for the mission performance analysis. A mission consists of a number of mission segments, for which time, distance, and fuel burn are evaluated. For the sizing task, certain missions are designated to be used for design gross weight calculations; for transmission sizing; and for fuel tank sizing. The mission parameters include mission takeoff gross weight and useful load. For specified takeoff fuel weight with adjustable segments, the mission time or distance is adjusted so the fuel required for the mission (burned plus reserve) equals the takeoff fuel weight. The mission iteration is on fuel weight or energy. Flight conditions are specified for the sizing task, and for the flight performance analysis. For the sizing task, certain flight conditions are designated to be used for design gross weight calculations; for transmission sizing; for maximum takeoff weight calculations; and for antitorque or auxiliary thrust rotor sizing. The flight condition parameters include gross weight and useful load. For flight conditions and mission takeoff, the gross weight can be maximized, such that the power required equals the power available. A flight state is defined for each mission segment and each flight condition. The aircraft performance can be analyzed for the specified state, or a maximum effort performance can be identified. The maximum effort is specified in terms of a quantity such as best endurance or best range, and a variable such as speed, rate of climb, or altitude. The aircraft must be trimmed, by solving for the controls and motion that produce equilibrium in the specified flight state. Different trim solution definitions are required for various flight states. Evaluating the rotor hub forces may require solution of the blade flap equations of motion.

Engine Airframe Flight

ANALYZE

Mission Analysis Performance Map Analysis Aerodynamics Map Performance Mission max takeoff GW adjust & fuel wt iteration each segment Outline of NDARC tasks.

Aircraft Description previous case Flight State fixed model or previous job or Figure 1-1 max effort / trim aircraft / flap equations design missions max GW Flight Condition Sizing Task size iteration

DESIGN

design conditions Data Structures and Input COMPREHENSIVE ANALYSIS STRUCTURAL DESIGN LAYOUT DESIGN COMPREHENSIVE ANALYSIS engine design FILES airframe geometry performance performance aerodynamics INTERFACE NDARC Interfaces.

Figure 1-2 input PRINT design and performance

NDARC

additional output additional cases aircraft solution description Data Structures and Input perf=1,&END _ mission=1,TASK _ size=0,TASK _ Illustration of NDARC input (primary input).

data=0,&END _ Figure 1-3a input=0,INIT _ (one mission, mission segment parameters as arrays) (one condition) (one condition) &JOB INIT &DEFN action=’ident’,created=’time-date’,title=’standard input’,&END !######################################################################## &DEFN action=’read file’,file=’engine.list’,&END &DEFN action=’read file’,file=’helicopter.list’,&END !======================================================================== &DEFN quant=’Cases’,&END &VALUE title=’Helicopter’,TASK &DEFN quant=’Size’,&END &VALUE nFltCond=0,nMission=0,&END !======================================================================== &DEFN quant=’OffDesign’,&END &VALUE title=’mission analysis’,nMission=1,&END &DEFN quant=’OffMission’,&END &VALUE &END !======================================================================== &DEFN quant=’Performance’,&END &VALUE title=’performance analysis’,nFltCond=2,&END &DEFN quant=’PerfCondition’,&END &VALUE &END &DEFN quant=’PerfCondition’,&END &VALUE &END !======================================================================== &DEFN action=’endofcase’,&END !######################################################################## &DEFN action=’endofjob’,&END Data Structures and Input WMTO=’input’, _ SDGW=’input’,SET _ design=’helicopter.design’,&END _ Illustration of NDARC input (secondary input file).

rotor=’radius+Vtip+sigma’,’radius+Vtip+sigma’, _ &END &END &END tank=’input’,SET _ Figure 1-3b &END perf=’none’,SET (other parameters in other structures) _ DGW=1,SET _ (Aircraft parameters) (geometry) (Rotor 1 parameters) (technology factors) title=’Helicopter’, SIZE FIX &DEFN action=’ident’,created=’time-date’,title=’Helicopter’,&END !######################################################################## ! default helicopter &DEFN action=’configuration’,&END &VALUE config=’helicopter’,rotate=1,&END !======================================================================== &DEFN quant=’Cases’,&END &VALUE title=’Helicopter’,FILE &DEFN quant=’Size’,&END &VALUE &END &DEFN quant=’Solution’,&END &VALUE &END !======================================================================== &DEFN quant=’Aircraft’,&END &VALUE &DEFN quant=’Geometry’,&END &VALUE &DEFN quant=’Rotor 1’,&END &VALUE !======================================================================== !======================================================================== &DEFN quant=’TechFactors’,&END &VALUE !######################################################################## &DEFN action=’endoffile’,&END Data Structures and Input .

QUANT=’Job’ namelist has the following DEFN = 44 = 41 = 48 = 45 = 49 = 43 = 42 = 47 = 6 = 46 = 40 = 40 = 40 = 5 unit number (and default) nuin nufile nufile nufile nuout nudesign nuperf nuaero nuengine nugeom nuacd nusoln nusketch nuerror input take effect for the next case. The namelists are read to define the action and contents of the subsequent DEFN QUANT=’Job’ file logical name standard input FILE FILE FILE standard output DESIGNn PERFn AEROn ENGINEn GEOMETRYn AIRCRAFTn SOLUTIONn SKETCHn ERRORn namelist, then JOB Primary Input Secondary Input File Aircraft Description Solution Output Design Performance Airframe Aerodynamics Engine Performance Geometry Aircraft Description Solution Sketch Errors Table 1-1. Input and output files. INPUT OUTPUT The primary input starts with a Input NDARC Input and Output Data Structures and Input 1–2 Figure 1-2 illustrates the input and output environment of NDARC. Table 1-1 lists the possible input and output files. A job reads input from one or more files. The primary input is obtained from standard input (perhaps redirected to a file). The primary input can direct the code to read other files, identified by file name or logical name. The input data are read in namelist format. Unit numbers are part of the job input. Output file names are part of the case input. Input files names are defined in the input itself. 1-2.1 Figure 1-3 illustrates NDARC input. information. The job parameters include initialization control, error action, and input/output unit numbers. Job parameters can be read during case input using The initialization takes place before case input, so changed initialization parameters in parameters.

, deletes deletes all (ignore ACTION=’delete’ rotor=’radius+Vtip+sigma’ .

_ SET , or ACTION=’delete all’ character string. A solution file (text or namelist with .

DEFN charge=’none’ PARENT and subsequent structures (all if number absent).

_ ACTION n .

QUANT=’PerfCondition’ SIZE , or jet=’none’ _ WMTO=’input’ (or EOF) is encountered in a secondary input file, the file is closed QUANT=’PerfCondition n’ _ deletes structure , or SIZE SET , , = last-case-input (default). A QUANT=’SizeCondition’ ACTION=’end’ input perf=’none’ _ _ are written when a file is created by NDARC, and read and stored for each input file. is absent. Note that each mission, with the mission segment parameters as arrays, is SDGW=’input’ _ INIT SIZE ACTION=’delete last’ ; value is 1 if absent; input variable can be TITLE SET QUANT=’OffMission n’ , PARENT , and namelist. The code searches for the keyword in the ds=’input’ _ CREATED DEFN limit _ (all if number absent); ; and each condition is input with n SET QUANT=’EngineParam’ , variables.

QUANT=’SizeMission n’ , ; the parent number is 1 if variable in the structure are set for no size iteration: namelists to define action and contents. When tank=’input’ Size _ are compared with the version and modification of the code, and input continues only if they match.

deletes structure identifies the file; probably there is only one identification per file, and only the last occurrence is stored. The identification consists string includes the structure number; if absent, the number is 1, or the next condition or mission. Parent structures may be required: ACTION SET , DEFN MODIFICATION QUANT=’OffMission’ , or QUANT DGW=1 _ : character string of code modification (length = 32).

QUANT=’EngineParam’ VERSION . MODIFICATION QUANT=’SizeCondition n’ , FIX identifies the data structure to be read (namelist format), initialized, or copied. Table 1-3 describes the options. The input corresponds to the data ACTION=’ident’ , : character string of creation time and date (length = 20).

: code version number as character string (length = 6). , input variables in the and : integer; engine model number for : integer; for copy action.

: character string (length = 32; case independent). : character string (length = 32, case independent); corresponds to data structure in input; string includes structure ACTION=’delete one’ : character string of title identifying input file (length = 80).

TITLE , : file name or logical name (length = 256).

QUANT QUANT PARENT TITLE VERSION ACTION SOURCE FILE CREATED MODIFICATION VERSION QUANT=’SizeMission’ a) b) number (1 or next condition/mission if absent). c) d) ENGINEMODEL e) f) g) h) i) namelist with CREATED wing=’area+span’ _ ACTION=’nosize’ DEFN Data Structures and Input Table 1-2 describes the options for the binary) can be written by an NDARC job and then read by a subsequent job, restoring the solution to the state that existed when the file was created. Then additional output and additional cases can be obtained. An aircraft description file can be written by an NDARC job and then read by a subsequent job, restoring the aircraft model (but not the solution). A secondary input file has and the code returns to primary input. A of the If present, The parameter structures of the analysis. The engine model number for input with A case inherits input for flight conditions and missions from the previous case if this input as specified by structure number); For SET FILE FILE (text) (binary) SOURCE=SRCnumber FILE FILE FILE ACTION=’endofjob’ namelist namelist VALUE VALUE function open and read secondary input file, name = open and read secondary input file, name = load aircraft description file, name = read complete solution file, name = read complete solution file, name = stop case input, execute case stop job input, execute case, exit code read read copy input from source (same structure), set structure variables to default values delete all conditions or missions delete one condition or mission delete last conditions or missions set input based on aircraft configuration set input for no size iteration identify file Secondary: close file, return to primary input Primary: same as ’text’ QUANT blank ’text’ not ’structure’ ’structure’ ’structure’ ’structure’ ’structure’ ’structure’ ’structure’ desc , quit , open , (or EOF) (or EOF) keyword — file aircraft solution solution end+case end+job — list copy init del+all del+one del+last config nosize ident end end options.

ACTION Table 1-2. ACTION Primary Input Only blank ’open file’ ’load aircraft’ ’read solution’ ’read solution’ ’end of case’ ’end of job’ Primary or Secondary Input blank ’read namelist’ ’copy input’ ’initialize’ ’delete all’ ’delete one’ ’delete last’ ’configuration’ ’nosize’ ’identification’ ’end’ ’end’ Data Structures and Input PARENT EngineModel number n maximum nFltCond nMission nMission nFltCond nRotor nWing nTail nTank nPropulsion nEngineGroup nJetGroup nChargeGroup nEngineModel nspeed nEngineTable nRecipModel nCompressorModel nMotorModel nJetModel nFuelCellModel nSolarCellModel nBatteryModel WRotor , DRotor , IRotor , as array as array PRotorTab , FltState FltState WChrgSys + + WEngSys , WWingTR , , WGear WDeIce , WJetSys , , PRotorPro WFuse , MissSeg MissSeg , , , WWing DEngSys DChrgSys EngineParam FltState FltState AGear , , , , , + + WDrive WTail WTank , DJetSys xxx , AFuse WFltCont _ , , , , PRotorInd AWing CostCTM , , ATail , FltCond MissParam MissParam FltCond , TECH Location data structures read Job Cases SizeParam one one OffParam one PerfParam one MapEngine MapAero Solution Cost Emissions Aircraft Systems Fuselage LandingGear Rotor Wing Tail FuelTank Propulsion EngineGroup JetGroup ChargeGroup EngineModel EngineParam EngineTable RecipModel CompressorModel MotorModel JetModel FuelCellModel SolarCellModel BatteryModel all all options.

QUANT Table 1-3. QUANT ’Job’ ’Cases’ ’Size’ ’SizeCondition n’ ’SizeMission n’ ’OffDesign’ ’OffMission n’ ’Performance’ ’PerfCondition n’ ’MapEngine’ ’MapAero’ ’Solution’ ’Cost’ ’Emissions’ ’Aircraft’ ’Systems’ ’Fuselage’ ’LandingGear’ ’Rotor n’ ’Wing n’ ’Tail n’ ’FuelTank n’ ’Propulsion n’ ’EngineGroup n’ ’JetGroup n’ ’ChargeGroup n’ ’EngineModel n’ ’EngineParamN n’ ’EngineTable n’ ’RecipModel n’ ’CompressorModel n’ ’MotorModel n’ ’JetModel n’ ’FuelCellModel n’ ’SolarCellModel n’ ’BatteryModel n’ ’TechFactors’ ’Geometry’ Data Structures and Input : , in binary or text format: identifying the file as text or binary: theDesign(kcase) QUANT theDesign(kcase) ADIMEN nrotor=m,nwing=m,ntail=m,ntank=m,npropulsion=m,nenginegroup=m,njetgroup=m,nchargegroup=m, ADIMEN nrotor=m,nwing=m,ntail=m,ntank=m,npropulsion=m,nenginegroup=m,njetgroup=m,nchargegroup=m, SDIMEN nsizecond=m,nsizemiss=m,nperfcond=m,noffmiss=m,&END _ _ _ nenginemodel=m,nenginetable=m,nrecipmodel=m,ncompressormodel=m,nmotormodel=m,njetmodel=m, nfuelcellmodel=m,nsolarcellmodel=m,nbatterymodel=m,&END nenginemodel=m,nenginetable=m,nrecipmodel=m,ncompressormodel=m,nmotormodel=m,njetmodel=m, nfuelcellmodel=m,nsolarcellmodel=m,nbatterymodel=m,&END Formats &DEFN action=’IDENT’,created=’time-date’,title=’xxx’,version=’n.n’,modification=’xxx’,&END &DEFN quant=’STRUCTURE n’,&END &VALUE param=value,&END &DEFN action=’NAMELIST’,quant=’STRUCTURE n’,&END &VALUE param=value,&END &DEFN action=’COPY’,quant=’STRUCTURE n’,source=#,&END &DEFN action=’IDENT’,created=’time-date’,title=’xxx’,version=’n.n’,modification=’xxx’,&END &VALUE &VALUE theStructure%xxx,&END &VALUE theStructure%xxx,&END &VALUE theStructure%xxx,&END &DEFN action=’AIRCRAFT’,file=’aircraft.acd’,&END &DEFN action=’IDENT’,created=’time-date’,title=’xxx’,version=’n.n’,modification=’xxx’,&END &VALUE &VALUE &VALUE theStructure%xxx,&END &VALUE theStructure%xxx,&END &VALUE theStructure%xxx,&END &DEFN action=’SOLUTION,quant=’TEXT’,file=’aircraft.soln’&END Data Structures and Input 1-2.2 Namelist input has the following format (see also figure 1-3). An aircraft description file is written in a separate file by NDARC, from This aircraft description file is read by identifying it in the primary input: A solution file is written in a separate file by NDARC, from This solution file is read by identifying it in the primary input, with Location ChargeGroup , or ).

JetGroup rotor(1)%SL , _ loc variables), and all geometry (all xxx namelist. All mission segments are defined in _ namelist.

EngineGroup , ). TECH rotor%SL OffMission _ or loc Propulsion PerfCondition , or Tail , variable that is the array (for example, Wing SizeMission , Location SizeCondition Rotor variables (long character string) to record information about the input.

notes input. Note that it is the variables) is input in a separate variables) is input in a separate Geometry or FltState variables are arrays. Each variable gets one more dimension, with the first array index always segment number.

variables. There are also , and FltState title variables, which are read as elements of the data structure (for example, and FltState TechFactors MissSeg , and Location FltCond MissSeg MissParam Conventions Software Tool Data Structures Data Structures and Input 1-2.3 Each flight condition ( Each mission ( this namelist, so Geometry input includes Variables can appear in more than one namelist. Specifically there are separate namelists for all technology factors (all variables), with corresponding options for output. A variable that is a scalar in the input becomes an array in the Case is not important in character string input. Character string input consists of keywords; the code searches for the keywords in the string. Default values are specified in the dictionary (blank implies a default of zero); all elements of arrays have the same default value. Tasks, aircraft, and components have 1–3 All information about data structures is contained in a dictionary file. This information includes the parameter name, dimension, type, default value, description, identification as input, and formats for write of the parameter. A software tool was created to manage the data, including construction of the module of data structures. The software tool reads this dictionary file and creates subroutines for the input process: namelist read, copy, print of input, initialization, set to default. This software tool is a program that manipulates character strings, to produce compilable module and subroutines for NDARC. 1–4 Table 1-4 outlines the data structures used for NDARC. The following chapters describe the contents of each structure. Note that a ”+” sign in the column between the type and description identifies input variables. Input variables can be changed by the analysis, so may not be the same at the end of a case as at the beginning. All variables, input and other, are initialized to zero or blank. If default values exist (only for input variables), they supersede that initialization.

FltAircraft FltFuse FltGear FltRotor(nrotormax) FltWing(nwingmax) FltTail(ntailmax) FltTank(ntankmax) FltProp(npropmax) FltEngn(nengmax) FltJet(njetmax) FltChrg(nchrgmax) FltState(nfltmax) auxtank(nauxtankmax) engine jet charger _ _ _ _ [Location]loc Weight WTank Weight WDrive [Location]loc DEngSys Weight WEngSys [Location]loc DJetSys Weight WJetSys [Location]loc DChrgSys Weight WChrgSys [EngineParam]Param [EngineParam]ParamN(nspeedmax) FuelTank(ntankmax) Propulsion(npropmax) EngineGroup(nengmax) JetGroup(njetmax) ChargeGroup(nchrgmax) EngineModel(nengmax) EngineTable(nengmax) RecipModel(nengmax) CompressorModel(nengmax) MotorModel(nengmax) JetModel(njetmax) FuelCellModel(nchrgmax) SolarCellModel(nchrgmax) BatteryModel(ntankmax) fuselage gear rotor pylon pivot nac wing tail _ _ _ _ _ _ _ _ [Location]loc AFuse Weight WFuse [Location]loc AGear Weight WGear [Location]loc [Location]loc [Location]loc [Location]loc PRotorInd PRotorPro PRotorTab IRotor DRotor Weight WRotor [Location]loc AWing Weight WWing WWingTR [Location]loc ATail Weight WTail Fuselage LandingGear Rotor(nrotormax) Wing(nwingmax) Tail(ntailmax) cg _ MissParam MissSeg(nsegmax) FltState(nsegmax) MissParam MissSeg(nsegmax) FltState(nsegmax) SizeParam FltCond(nfltmax) FltState(nfltmax) Mission(nmissmax) OffParam Mission(nmissmax) PerfParam FltCond(nfltmax) FltState(nfltmax) CostCTM [Location]loc Weight XAircraft Weight WFltCont WDeIce Cases Size OffDesign Performance MapEngine MapAero Solution Cost Emissions Aircraft Systems Table 1-4. NDARC data structures. Design Data Structures and Input (direction Rotor%rotate nChargeGroup=0 var , _ trim ’coll’,’latcyc’,’lngcyc’,’pedal’,’pitch’,’roll’ ’coll’,’lngcyc’,’pitch’,’pedal’ ’coll’,’lngcyc’,’pitch’ ’coll’,’pedal’ ’coll’ ’coll’ ’coll’,’latcyc’,’lngcyc’ ’coll’ ’coll’ ’coll’,’latcyc’,’lngcyc’,’pedal’,’prop’,’roll’ , the analysis defines a number of input parameters nJetGroup=0 , is: (each rotor, only for multicopter). The convention is that (number of rotors, only for multicopter), multicopter _ nEngineModel=1 , ACTION=’configuration’ multicopter=#,#,&END _ ACTION=’configuration’ Aircraft%nRotor Rotor%ang input. With nEngineGroup=1 , ; compound state not active tandem=#,#,ang _ trim _ nPropulsion=1 quant QUANT=’Aircraft’ ; _ (rotorcraft configuration), trim ’force x’,’force y’,’force z’,’moment x’,’moment y’,’moment z’ ’force x’,’force z’,’moment y’,’moment z’ ’force x’,’force z’,’moment y’ ’force z’,’moment z’ ’force z’ ’CTs rotor 1’ ’CTs rotor 1’,’betac 1’,’betas 1’ ’P margin 1’ ’force x’ ’force x’,’force y’,’force z’,’moment x’,’moment y’,’moment z’ in the trim=IDENT nTail=2 _ , . Note that default values are defined for all input quantities.

config mtrim 6 4 3 2 1 1 3 1 1 6 Aircraft%config nWing=0 (each rotor, only for tandem helicopter), and config.f90 _ trim _ FltAircraft%STATE tandem control=’coll’,’latcyc’,’lngcyc’,’pedal’,’tailinc’,’elevator’,’rudder’ _ _ input IDENT ’free’ ’long’ ’symm’ ’hover’ ’thrust’ ’rotor’ ’windtunnel’ ’power’ ’ground’ ’comp’ IDENT , (except multicopter), , selected by Rotor%overlap control=1 _ /σ trim=9 _ ncontrol=7 T nRotor=2 C nstate nstate namelist contains only the parameters All Configurations 6-variable longitudinal symmetric 3-variable hover thrust and torque hover thrust hover rotor wind tunnel full power ground run compound VALUE &DEFN action=’configuration’,&END &VALUE config=’aaaa’,nRotor=#,rotate=#,#,overlap Chapter 2 Input Based on Configuration The rotorcraft configuration is identified by the variable in order to facilitate modelling of conventional configurations. The input required to execute The of rotation, each rotor), the first rotor is the main rotor for the helicopter or compound configuration; the front rotor for the tandem configuration; the right rotor for the tiltrotor configuration. This capability has been implemented for rotorcraft, helicopter, tandem, coaxial, tiltrotor, compound, multicopter, and airplane configurations. The analysis creates the following input, through the code in the file 2–1 a) Components: b) Aircraft Aircraft controls: Control states: Trim states: (no connection to aircraft controls) (no spinner drag) diam=0 _ spin=0.

_ INPUT , (input quantity is tip speed) cant=0 (no connection of tail incidence to aircraft controls) DRotor%DoQ gear=1 _ _ , WTank%nplumb=2 , (all controls, all states) (no connection to aircraft controls) INPUT incid=0 INPUT int=1 _ , _ , Dspin=1 _ (not tilting) diam=0.

_ incid=0 (helicopter) (helicopter) _ T INPUT , incid=0 _ xmsn=0 _ rotor=2 _ INPUT Vtail=1 Htail=1 , _ _ cant=0.

WTank%ntank KIND DRotor%SET (all controls, all states, all panels) INPUT _ , , , T DN=0 , _ (no fixed wing flight controls, no conversion controls) aileron=0 (rotor reference) _ incid=0.

(not tilt); _ lngcyc=0 xmsn=1 WRotor%KIND INPUT Wgear%nLG=3 _ T _ incid=0.

, _ , ) T WTail%MODEL WTail%MODEL CVfc=0 kPropulsion=1 , , INPUT , _ /σ Mauxtanksize=1 , KIND T wt=1 INPUT , Idrag=0. (rotor), (1 for TPP tilt, 2 for hub moment, 3 for lift offset) _ , C tail=2 tail=1 TailVolRef=1 aileron=0. _ _ gear , _ _ MODEL (fixed gear), lngcyc=0.

T , _ , T config=1 flaperon=0 latcyc=0 latcyc=1 (all controls, all states) KIND KIND _ , _ _ _ LG=0 STATE _ (helicopter), , (internal tank), kPropulsion=1 TailVol=2 FWfc=0 _ _ (1 for thrust and TPP command) KIND INPUT (tail control connection to aircraft controls), INPUT , , KIND , incid=0.

(1 for thrust, 2 for flaperon=0.

latcyc=0. _ (fixed shaft) _ _ (not tilt) direction=’main’ KIND T T T nGear=1 , place=1 , , MODEL aeroaxes=1 coll=0 flap=0 cont=1 _ _ _ _ control=1 coll=2 lngcyc=1 tilt=0 _ _ _ _ WRotor%MODEL SET Idrag=0.

coll=0. flap=0. cont=0.

_ _ _ INPUT T KIND KIND KIND KIND INPUT T INPUT T Input Based on Configuration c) Systems: d) Landing Gear: e) Fuel Tank: f) Rotor First rotor is primary: Second and other rotors are dependent: Configuration: Drag: Weight: Control: g) Wing Control: Drag: h) Tail First tail is horizontal tail: Second tail is vertical tail: Configuration: Control: i) Propulsion: ) ’pedal’ (standard position) geom=0 _ SET , WDrive%fPower=0.15 direction=’x’ , (no spinner drag) , (elevator direct control) spin=0.

_ fPsize=1.

mainRotor=1 (rotor collective connection to aircraft control , r (sized), − cont(6,1)=1.

_ DRotor%CD T WDrive%fTorque=0.03 , , ), power=0 _ coll(4,1)= _ spin=0.

geom=’tailrotor’ T ’tailinc’ _ _ , SET ), SET , geom=’standard’ coll=1 spec _ _ _ (rotor control connection to aircraft controls) N SET WDrive%fShaft=0.1 , , DRotor%DoQ fArea=0.

, INPUT , lngcyc=1 _ (no connection to aircraft controls) 1 1 fArea=1. lngcyc(3,1)=-1.

_ Dspin=1 , − − _ T (tail rotor) = , = (gear ratio from fThrust=1.

r INPUT r , r yaw=0 , _ − 0) 0) fDGW=1. < < gear=1 , (all controls, all states) config=2 _ WDrive%ndriveshaft=1 INPUT _ latcyc=1 , , DRotor%SET twin=’none’ _ , _ (not tilt) (incidence connection to aircraft control INPUT latcyc(2,1)= yaw=0.

, _ (thrust and NFP command); (rudder direct control) _ incid=0 INPUT T T _ spin=0.

, , , config=’tail+antiQ’ _ (tail incidence connection to aircraft controls) config=’main’ (no tail) Idrag=0.

Rotor(1)%rotate Rotor(1)%rotate , (sponson) Swet coll=1 INPUT , WRotor%MODEL _ control=2 , ; if ( ; if ( incid(5,1)=1.

, incid=0. _ _ incid=1 _ nTail=0 _ T drag=1 cont(7,1)=1.

kPropulsion=1 T = 1 coll(1,1)=1. = 1 WDrive%ngearbox=2 _ _ PRotorInd%MODEL place=2 , _ T r r INPUT T KIND amp=0 Sspin=1 _ _ INPUT MODEL SET amp=0. Helicopter Tandem _ INPUT T rotation: control: control: direction=’tail’ rotation: control: Input Based on Configuration j) Engine Group Configuration: Drag: k) Engine Group, Jet Group, Charge Group Control: 2–2 a) Rotor First rotor is main rotor: Second rotor is tail rotor: Performance: Drag: b) Tail Control: Horizontal tail: Vertical tail: c) Propulsion: 2–3 a) Components: b) Fuel Tank: twin=2 twin=2 _ _ int int _ _ WDrive%fPower=0.6 WDrive%fPower=0.6 (no spinner drag) (no spinner drag) , , IRotor%MODEL IRotor%MODEL , , spin=0. spin=0.

_ _ twin=0.85 twin=0.85 _ _ DRotor%CD DRotor%CD r WDrive%fTorque=0.6 WDrive%fTorque=0.6 , , ; fRadius=1.

; − , tandem r _ PRotorInd%Kf PRotorInd%Kf , spin=0. spin=0.

, _ _ fRadius=1.

overlap , − latcyc(4,1)= latcyc(4,1)= twin=1.

_ _ twin=1.

lngcyc(3,1)=-1. lngcyc(3,1)=-1.

_ geom=’coaxial’ _ = 1 _ _ T T _ , , (rotor control connection to aircraft controls) T (rotor control connection to aircraft controls) T h r r , , , WDrive%fShaft=0.1 WDrive%fShaft=0.1 r r r r ) , SET , − − DRotor%DoQ DRotor%DoQ − − , , − − , h = = geom=’tandem’ − _ r r lngcyc=1 lngcyc=1 ; _ ; _ 1 1 1 1 √ PRotorInd%Kh SET PRotorInd%Kh h Dspin=1 Dspin=1 − − , fArea=.5 − − , _ , _ , − = latcyc(2,1)= = latcyc(2,1)= = = _ _ INPUT INPUT h latcyc(2,1)= latcyc(2,1)= r (rotor control connection to aircraft controls) r (rotor control connection to aircraft controls) r r , _ , _ T T rotate=-Rotor(1)%rotate , , T T 0) 0) 0) 0) − , , , r r rotate=-Rotor(1)%rotate < < < < fDGW=.5 , fDGW=.5 , , WDrive%ndriveshaft=1 WDrive%ndriveshaft=0 )(cos latcyc=1 latcyc=1 , latcyc=1 latcyc=1 , DRotor%SET DRotor%SET twin=’tandem’ twin=’coaxial’ _ _ _ _ , , /π _ _ (elevator direct control) coll(3,1)=-1. coll(3,1)= 1. coll(4,1)= coll(4,1)= _ _ _ _ = (2 (rudder direct control) otherRotor=1 INPUT T INPUT T INPUT T INPUT T spin=0. spin=0.

, , , , , , , , m _ _ otherRotor=1 otherRotor=2 otherRotor=2 Rotor(1)%rotate Rotor(1)%rotate Rotor(1)%rotate Rotor(1)%rotate Swet Swet coll=1 coll=1 coll=1 coll=1 , , , from _ _ _ _ , if ( , if ( , if ( , if ( cont(6,1)=1.

_ T cont(7,1)=1.

config=’main+tandem’ config=’main+coaxial’ m/ = 1 coll(1,1)=1. = 1 coll(1,1)=1. WDrive%ngearbox=2 = 1 coll(1,1)=1. = 1 coll(1,1)=1. WDrive%ngearbox=1 _ PRotorInd%MODEL PRotorInd%MODEL _ _ _ _ T r r r r − INPUT T INPUT T INPUT T INPUT T Sspin=1 Sspin=1 1 _ _ SET SET Coaxial fArea= rotation: control: control: rotation: control: control: rotation: control: control: rotation: control: control: Input Based on Configuration c) Rotor Configuration: First rotor is front rotor: Second rotor is aft rotor: Performance: Drag: d) Propulsion: 2–4 a) Rotor Configuration: First rotor is lower rotor: Second rotor is upper rotor: Performance: Drag: b) Tail Horizontal tail: Vertical tail: c) Propulsion: (no hub drag) hub=0.

_ WingForRotor=1 , twin=2 DRotor%CDV _ fRadius=1.

, (wing tip weight move) int _ hub=0.

_ ) fWmove=1.

, ’tilt’ (from clearance), IRotor%MODEL , DRotor%DoQV , Wmove=1 _ ext=0 twin=1.

_ ) _ TRgeom=1 SET _ Vhub=1 SET ’coll’ _ , KIND lngcyc(4,1)=1.

) (input quantity is gear ratio) lngcyc(4,1)=-1.

, _ _ T T , , PRotorInd%Kf ’latcyc’ , gear=2 _ DRotor%SET , cruise(1)=1 _ twin=1.

INPUT geom=’tiltrotor’ _ RotorOnWing(2)=2 cruise=2 ; _ _ , lngcyc(3,1)=-1. hub=0.

lngcyc(3,1)=-1.

kgear _ _ (rotor control connection to aircraft controls) (rotor control connection to aircraft controls) (incidence connection to aircraft control _ , T SET T , (helicopter mode reference), kcont ; , , (wing control connection to aircraft controls) lngcyc=1 lngcyc=1 _ _ ref=90.

conv(1)=1 PRotorInd%Kh conv=1 fArea=1. _ DRotor%CD _ , _ , , (fixed wing flight controls, conversion control) incid(5,2)=1.

coll(2,1)=-1. coll(2,1)=1. _ aileron=1 _ _ _ incid T RotorOnWing(1)=1 kgear INPUT T INPUT T (collective connection to aircraft control , (tiltrotor) (engine at each nacelle) kcont , , , , , , , nVincid=1 hub=0.

_ rotate=-Rotor(1)%rotate CVfc=1 fDGW=.5 control=’coll’,’latcyc’,’lngcyc’,’pedal’,’tilt’,’flap’,’flaperon’,’elevator’,’aileron’,’rudder’ , INPUT _ _ , , coll=1 coll=1 _ _ twin=’tiltrotor’ hover=1 _ Idrag=90.

_ (tilting) (state 1 for helicopter mode, state 2 for airplane mode) hover(1)=1 coll(1,2)=1.

coll(1,1)=1. coll(1,1)=1. incid(5,1)=1.

_ IDENT MODEL (retractable) _ _ _ fFuelWing(1)=1. _ , , T INPUT T INPUT T T kcont , DRotor%DoQ nEngineGroup=2 flaperon=1 (independent), kgear , nRotorOnWing=2 _ , otherRotor=1 , LG=1 otherRotor=2 rotor=1 _ _ (tiltrotor), (airplane mode aileron connection to aircraft control (wing), control=2 FWfc=1 _ _ incid=1 INPUT _ , ncontrol=10 KIND (shaft control = incidence), Dhub=1 _ nWing=1 flaperon=3 config=’main+tiltrotor’ fDGW=1.

nstate _ PRotorInd%MODEL place=3 MODEL INPUT tilt=1 aeroaxes=2 flap=1 _ _ _ KIND WRotor%KIND SET Tiltrotor aileron(2,2)=-1.

_ KIND helicopter mode control: helicopter mode control: helicopter mode control: helicopter mode control: control: DRotor%SET INPUT T Input Based on Configuration 2–5 a) Components: b) Aircraft Aircraft controls: Control states: Control state in conversion: Drive state in conversion: c) Systems: d) Landing Gear: e) Fuel Tank: f) Rotor Configuration: First rotor is right rotor: Second rotor is left rotor: Airplane mode control state: Tilt: Performance: Weight: Drag: g) Wing Configuration: Control: (tiltrotor) DEngSys%CDV=0.

, ) ’tilt’ Vtail=2 _ WDrive%fPower=0.6 , DEngSys%DoQV=0.

, Wtail%MODEL , Vdrag=1 WDrive%fTorque=0.6 _ ) Htail=2 ; _ ) ’lngcyc’ ’pedal’ DEngSys%SET geom=’standard’ ; _ Wtail%MODEL (rotor control connection to aircraft controls) SET WDrive%fShaft=0.1 , , (nacelle incidence connection to aircraft control (no engine nacelle drag) lngcyc=1 _ fArea=1.

, − Idrag=90. DEngSys%CD=0.

= (wing reference); , , INPUT incid(5,2)=1.

r _ , (flaperon direct control) T 0) , (tiltrotor) (aileron direct control) fDGW=1. < drag=0 control=’coll’,’latcyc’,’lngcyc’,’pedal’,’tailinc’,’elevator’,’rudder’,’prop’,’aileron’,’flap’ (rudder direct control) , _ _ WDrive%ndriveshaft=1 (elevator direct control) latcyc=1 , (tiltrotor) (flap direct control) _ TailVolRef=1 geom=1 , _ IDENT MODEL , , incid(5,1)=1.

DEngSys%DoQ=0.

; compound state active INPUT _ nWing=1 wing=3 SET , , flaperon(7,2)=1. , _ T , _ aileron(9,2)=1. config=’main’ ; cont(10,2)=1.

T _ nVincid=1 Rotor(1)%rotate cont(8,2)=1. _ , RotorForEngine=2 T flap(6,2)=1. TailVol=1 _ T _ , _ nVincid=1 Swet=0.

coll=1 T , (airplane mode elevator connection to aircraft control (airplane mode rudder connection to aircraft control RotorForEngine=1 T , drag=1 trim=10 _ , ; if ( _ _ , ncontrol=10 incid=1 nRotor=3 _ KIND fPsize=0.5 = 1 WDrive%ngearbox=2 r nstate INPUT Swet=1 _ INPUT WWing%MODEL SET flap(6,1)=1. flaperon(7,1)=1. aileron(9,1)=1. cont(3,2)=1. cont(8,1)=1. cont(4,2)=1. cont(10,1)=1. Compound _ _ _ _ _ _ _ T T T T T T T DEngSys%SET rotation: control: Input Based on Configuration Weight: h) Tail Configuration: Horizontal tail control: Vertical tail control: i) Propulsion: j) Engine Group Configuration: First engine group: Second engine group: Control: Drag: 2–6 a) Components: b) Aircraft Aircraft controls: Trim states: c) Rotor First rotor is main rotor: ) ) ’pedal’ ’prop’ WDrive%fPower=0.15 (no spinner drag) , (elevator direct control) spin=0.

_ mainRotor=1 (rotor collective connection to aircraft control , r (rotor collective connection to aircraft control − cont(6,1)=1.

_ DRotor%CD T WDrive%fTorque=0.03 , , ), (rotor collective connection to aircraft controls) coll(4,1)= coll(8,1)=1.

_ _ r spin=0.

geom=’tailrotor’ T T ’tailinc’ geom=’standard’ _ _ , , _ SET SET , , coll=1 coll=1 _ _ coll(4,1)= geom=’multicopter’ _ coll=1 _ T _ WDrive%fShaft=0.1 , ) DRotor%DoQ , fArea=0. SET a , INPUT INPUT fArea=0.

(wing control connection to aircraft controls) , , , INPUT cos( multicopter _ lngcyc(3,1)=-1.

_ Dspin=1 − _ (auxiliary thrust) fArea=1.

T ang (tail rotor) aileron=1 , , = fThrust=1. _ r = fThrust=1.

, r , coll(3,1)= a − _ ; 0) T , < INPUT − ) config=3 config=2 _ , a _ WDrive%ndriveshaft=1 = DRotor%SET , twin=’none’ (parametric) r , _ sin( (incidence connection to aircraft control 0) − latcyc(2,1)= fDGW=1/nRotor _ < (thrust and NFP command); (thrust and NFP command); (rudder direct control) , T flaperon=1 wing=2 spin=0.

, _ config=’tail+antiQ’ _ (thrust and NFP command); _ (tail incidence connection to aircraft controls) (no tail) config=’prop+auxT’ Rotor(1)%rotate rotate (aileron direct control) coll(2,1)= Swet _ WRotor%MODEL (flap direct control) , WRotor%MODEL INPUT , T control=2 control=2 ; if ( incid(5,1)=1. ; if ( , , _ _ , _ incid=1 nTail=0 _ T cont(7,1)=1. control=2 fDGW=1. config=’main’ coll(1,1)=1. = 1 = 1 _ WDrive%ngearbox=3 _ PRotorInd%MODEL _ T r r T KIND KIND Sspin=1 flap=1 _ _ nVincid=1 INPUT KIND WWing%MODEL SET aileron(9,1)=1. flap(10,1)=1. Multicopter coll(1,1)=1.

_ _ _ control: direction=’tail’ rotation: control: direction=’prop’ control: INPUT T T rotation: T Input Based on Configuration Second rotor is tail rotor: Third rotor is propeller: Performance: Drag: d) Wing Configuration: Control: Weight: e) Tail Control: Horizontal tail: Vertical tail: f) Propulsion: 2–7 a) Components: b) Rotor Configuration: Control: ) (elevator direct control) ’coll’ cont(6,1)=1.

_ T WDrive%fPower=0.6 , (wing reference) (no spinner drag) kRef=1 , (longitudinal stick), spin=0. Ref=2 _ WDrive%fTorque=0.6 _ ; KIND (rotor collective connection to aircraft control cont(3,1)=1.

_ DRotor%CD T , ), spin=0.

WDrive%fShaft=0.1 ’tailinc’ _ coll(1,1)=1.

, _ T , coll=1 _ WDrive%fShaft=0.1 geom=’standard’ , DRotor%DoQ _ , SET INPUT (aileron direct control) (geometry scaled with wing span); , (rudder direct control) Dspin=1 _ (auxiliary thrust) = longitudinal stick kScale=1 fDGW=0.

, , WDrive%ndriveshaft=nRotor-1 , lngcyc cont(7,1)=1.

(wing reference) config=3 aileron(8,1)=1. _ control=’coll’,’latcyc’,’lngcyc’,’pedal’,’tailinc’,’elevator’,’rudder’,’aileron’,’flap’ _ _ T (wing control connection to aircraft controls) _ WDrive%ndriveshaft=1 T , DRotor%SET scale=2 twin=’none’ (parametric) _ , _ (incidence connection to aircraft control fThrust=1.

(fixed wing flight controls) , IDENT (pedal), , (wing span reference) KIND nWing=1 , aileron=1 wing=2 spin=0.

, _ = lateral stick, (thrust and NFP command); _ _ TailVolRef=1 (tail incidence connection to aircraft controls) , (lateral stick), FWfc=1 Swet _ WRotor%MODEL , latcyc geom=2 INPUT Lscale=2 , (flap direct control) incid(5,1)=1.

_ _ , ncontrol=9 _ incid=1 nRotor=1 _ T control=2 cont(4,1)=1.

fDGW=1.

WDrive%ngearbox=nRotor WDrive%ngearbox=1 _ TailVol=1 _ PRotorInd%MODEL _ T KIND MODEL Sspin=1 flap=1 INPUT _ config=’prop+auxT’ _ KIND nVincid=1 INPUT WWing%MODEL KIND = propeller, SET Airplane aileron(2,1)=1. flap(9,1)=1.

_ _ coll direction=’prop’ INPUT T T Input Based on Configuration Performance: Drag: c) Propulsion: 2–8 a) Components: b) Solution: c) Aircraft Geometry: Aircraft controls: d) Systems: e) Rotor Propeller: Control: f) Wing Configuration: Control: Weight: g) Tail: Control: Horizontal tail: Vertical tail: h) Propulsion: 36 5 4 8 20 10 6 40 5 4000 100 19 20 21 22 23 24 25 26 27 28 29 tank Bmargin rotorL rotorfL CLs rotorV rotorX rotorfX CXs XoQ CTs _ _ _ _ _ _ _ _ _ _ _ QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT _ _ _ _ _ _ _ _ _ _ _ mpsimax npanelmax nauxtankmax ngearmax nratemax nengtmax nengkmax nengrmax nspeedmax nrowmax naeromax TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM 20 21 41 20 200 4 20 16 20 16 51 40 4 5 6 7 8 0 1 2 3 4 5 none end range rangelow range100 climb _ _ _ _ _ _ liftoff rotation transition climb brake _ _ _ _ _ QUANT QUANT QUANT QUANT QUANT QUANT takeoff takeoff takeoff takeoff takeoff _ _ _ _ _ _ _ _ _ _ _ nsegmax nfltmax ndesignmax ncontmax nsweepmax qsweepmax ntrimstatemax mtrimmax nvelmax ntablemax nrmax mrmax SET SET SET SET SET MAX MAX MAX MAX MAX MAX Value 10 40 8 4 8 4 4 10 8 6 4 20 Value 0 1 2 3 4 5 6 7 8 9 10 error file ident list copy init delete delone dellast config nosize _ _ _ _ _ _ _ _ _ _ _ Chapter 3 Parameters and Constants Parameters ncasemax nfilemax nrotormax npropmax nengmax njetmax nchrgmax nstatemax nwingmax ntailmax ntankmax nmissmax Constants ACTION ACTION ACTION ACTION ACTION ACTION ACTION ACTION ACTION ACTION ACTION 30 31 32 33 34 35 36 37 38 39 40 41 0 -1 -2 -3 -4 -5 -6 -7 -8 -9 0 -1 -2 -3 0 1 2 3 4 5 6 7 found found Tmargs Tmargt betac betas hubMx hubMy hubQ wingL wingfL CL Lmargin tailL _ _ _ _ _ _ _ _ _ _ _ _ _ _ none not alpha beta not pitch roll ROC side speed turn pullup Vtip Nspec rotorcraft helicopter tandem coaxial tiltrotor compound multicopter airplane _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ VAR VAR VAR VAR QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT VAR VAR VAR VAR VAR VAR VAR VAR VAR VAR _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM AERO AERO AERO AERO RCCONFIG RCCONFIG RCCONFIG RCCONFIG RCCONFIG RCCONFIG RCCONFIG RCCONFIG 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 0 -1 -2 -3 -4 -5 -6 -7 -8 -9 -10 -11 -12 -13 -14 -15 -16 -17 -18 angle power PoV alt Pmargin Qmargin PQmargin Jmargin PJmargin QJmargin PQJmargin Bmargin Lmargin Tmargs Tmargt _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ none vel ROC side alt turn pullup xaccF yaccF zaccF xaccI yaccI zaccI xaccG yaccG zaccG pitch roll Vtip _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT VAR VAR VAR VAR VAR VAR VAR VAR VAR VAR VAR VAR VAR VAR VAR VAR VAR VAR VAR _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ MAX MAX MAX MAX MAX MAX MAX MAX MAX MAX MAX MAX MAX MAX MAX MAX MAX MAX MAX MAX MAX MAX MAX MAX MAX MAX MAX MAX MAX MAX MAX MAX MAX MAX 11 12 13 14 15 1 2 3 1 2 3 4 5 6 1 2 3 1 2 1 2 1 2 3 1 2 3 4 5 1 2 3 4 5 radius area _ _ chrg none _ _ none radius thrust radius DL ratio scale not area WL not span ratio engine rotor none _ _ _ _ _ _ _ _ _ _ _ _ _ desc soln endfile endcase endjob _ _ _ jet none _ _ _ _ _ _ _ newcase onecase endofjob init size miss perf maps out _ _ _ _ _ _ _ _ _ perf perf perf jet jet charge charge rotor rotor rotor rotor rotor rotor rotor rotor wing wing wing wing wing _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ Parameters and Constants ACTION ACTION ACTION ACTION ACTION STATE STATE STATE STATE STATE STATE STATE STATE STATE SIZE SIZE SIZE SIZE SIZE SIZE SIZE SIZE SIZE SIZE SET SET SET SET SET SET SET SET SET SET 1 2 3 4 5 6 7 0 1 2 3 4 5 0 1 2 3 4 0 1 2 3 4 5 6 7 8 9 0 1 2 3 4 5 twin _ main tail prop tandem coaxial tiltrotor not _ _ _ _ _ _ _ none sidebyside coaxial tandem multirotor _ _ _ _ _ standard tiltrotor coaxial tandem tailrotor multicopter free span bratio edge station radius _ _ _ _ _ _ _ _ _ _ _ _ twin twin twin twin twin none mu muz alpha muTPP muzTPP alphaTPP CTs Mx Mat _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ geom geom geom geom geom geom panel panel panel panel panel panel _ _ _ _ _ _ _ _ _ _ _ _ ROTORCONFIG ROTORCONFIG ROTORCONFIG ROTORCONFIG ROTORCONFIG ROTORCONFIG ROTORCONFIG SET SET SET SET SET SET MODEL MODEL MODEL MODEL MODEL tablevar tablevar tablevar tablevar tablevar tablevar tablevar tablevar tablevar tablevar SET SET SET SET SET SET -19 1 2 3 4 5 6 7 8 9 10 1 2 3 -1 -2 -3 1 2 3 4 5 6 7 8 9 10 11 12 13 1 2 3 4 dtem temp dtemp temp _ _ dtemp temp table dtemp temp _ _ _ _ _ _ _ input dens notair std std std polar polar polar trop trop trop hot hot hot hot Nspec _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ input ref speed conv TAS CAS Mach _ _ _ _ general hover vert right left rear Vfwd Vmag climb takeoff _ _ _ _ _ _ _ _ _ _ _ _ _ VAR vel vel vel vel vel vel vel vel vel vel vel2 vel2 vel2 atmos atmos atmos atmos atmos atmos atmos atmos atmos atmos atmos atmos atmos atmos atmos atmos Vtip Vtip Vtip Vtip _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ MAX SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET 6 7 8 9 10 1 2 3 4 1 2 3 4 5 1 2 3 4 5 1 2 3 4 1 2 3 4 5 6 7 8 9 10 11 span _ input fDGW fSDGW maxfuel perf input fDGW fWMTO maxfuel perf _ _ _ _ _ _ _ _ _ _ radius width hub panel not input Ratio Pav Preq input miss misspower fmiss DGW SDGW WMTO fDGW fSDGW fWMTO input maxP maxQ maxPQ maxJ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ wing wing wing wing wing tank tank tank tank SDGW SDGW SDGW SDGW SDGW WMTO WMTO WMTO WMTO WMTO limit limit limit limit GW GW GW GW GW GW GW GW GW GW GW _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ Parameters and Constants SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET 6 7 8 9 10 11 12 13 1 2 3 4 5 1 2 3 4 5 6 7 8 9 1 2 3 1 2 RPTEM table recip comp compreact motor gen motorgen motorcell fuelcell solarcell _ _ _ _ _ _ _ _ _ _ _ RPJEM react simple width hub adjust area Sratio chord cratio taper _ _ _ _ _ _ _ _ _ _ _ area vol span AR chord engine engine engine engine engine engine engine engine engine jet jet jet charge charge _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ panel panel panel panel panel panel panel panel tail tail tail tail tail _ _ _ _ _ _ _ _ _ _ _ _ _ SET SET SET SET SET SET SET SET SET SET SET SET SET MODEL MODEL MODEL MODEL MODEL MODEL MODEL MODEL MODEL MODEL MODEL MODEL MODEL MODEL 5 6 7 8 9 10 11 0 1 2 0 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 found _ not forcex forcey forcez momentx momenty momentz nz nx ny power Pmargin Qmargin powerEG PEGmarg thrust Jmargin charge Cmargin _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ default extend retract _ _ _ hover cruise man OEI xmsn mu Mat _ _ _ _ _ _ _ LG LG LG _ _ _ QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT QUANT _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ Vtip Vtip Vtip Vtip Vtip Vtip Vtip _ _ _ _ _ _ _ SET SET SET SET SET SET SET STATE STATE STATE TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM TRIM 12 13 14 15 16 17 18 1 2 3 4 5 1 2 3 4 1 2 3 4 5 6 7 8 9 0 1 2 3 taxi dist time hold climb spiral fuel burn takeoff _ _ _ _ _ _ _ _ _ none start groundrun enginefail _ _ _ _ maxPJ maxQJ maxPQJ source fsource payfuel paymiss none input delta scale pay fuel payfuel miss paymiss _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ MissSeg MissSeg MissSeg MissSeg MissSeg MissSeg MissSeg MissSeg MissSeg _ _ _ _ _ _ _ _ _ GW GW GW GW GW GW GW UL UL UL UL UL pay pay pay pay takeoff takeoff takeoff takeoff _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ Parameters and Constants SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET SET KIND KIND KIND KIND KIND KIND KIND KIND KIND SET SET SET SET 1 0 1 0 2 Default ) to analyze case #1 design in subsequent cases input=2 Version n.n, modification "xxx" _ INIT : : if always start-last-case, then case starts from default use input data if default, all input variables set to default values if last-case-input, then case inherits input at beginning of previous case if last-case-solution, then case inherits input at end of previous case if default, all other variables set to default values _ _ INIT INIT number n.n modification string for headers ( input parameters (0 default, 1 last case input, 2 last case solution) other parameters (0 default, 1 start of last case, 2 end of last case) action on error (0 none, 1 exit) action on version mismatch in input (0 none, 1 exit) status keyword for write (0 unknown, 1 replace, 2 new, 3 old) NDARC Version (set by main program) Initialization Errors File open Description + + + + + + + + Type c*6 c*32 c*64 int int int int int status _ input data error version _ _ _ _ Chapter 4 Common: Job Variable version modification versionout INIT INIT ACT ACT OPEN 5 6 40 41 42 43 44 45 46 47 48 49 for secondary file) nufile ) ) ) ) ) ) AEROn ncasemax for primary file, ) ) ENGINEn AIRCRAFTn nuin ) can be changed as appropriate for computer OS GEOMETRYn ) SKETCHn PERFn status _ SOLUTIONn DESIGNn ERRORn OPEN standard input secondary file input standard output design ( performance ( airframe aerodynamics ( engine performance ( geometry output ( aircraft description ( solution ( sketch output ( errors ( default input/output unit numbers usually acceptable default input output current case number number of cases (maximum case state job state design output state (1 file open) performance output state (1 file open) geometry output state (1 file open) errors output state (1 file open) nuout or nuerror scratch structure file input status (0 for primary file, 1 for secondary file, 2 for aircraft or solution file) unit number for input ( Input/output unit numbers Analysis Input + + + + + + + + + + + + + + + int int int int int int int int int int int int int int int int int int int int int FltState int int state state _ state _ state _ _ state input design perf geom error _ state _ _ _ _ _ _ Common: Job nuin nufile nuout nudesign nuperf nuaero nuengine nugeom nuacd nusoln nusketch nuerror kcase ncase case job out out out out nuinit fscratch kind nread theDesign(kcase) copied to theLastCaseInput theInput + data) ; first is standard input) theInput + nfilemax incremented and action=IDENT kcase ) ) ) theDesign(ncase) + used to print only what changed from last case Aircraft Aircraft Aircraft Job used for input (not changed by analysis) system data = all data structure parameters = input (can be changed by analysis) or other (generated by analysis) theInput theLastCaseInput after case input concluded, number of identifications (maximum title creation date case start case end case job case start case end case job number of rotors ( number of wings ( number of tails ( Input file identification (stored from Design Input Input from last case CPU time Clock time Case dimensions int c*80 c*20 Design Design Design real real real real int int real real int int int start(8,ncasemax) end(8,ncasemax) start(ncasemax) end(ncasemax) _ _ _ _ case(ncasemax) job _ _ case case case case case(ncasemax) job _ _ _ _ _ _ case case title(nfilemax) created(nfilemax) _ _ case _ _ _ Common: Job ninputfile input input theDesign(ncasemax) theInput theLastCaseInput CPUtime CPUtime CPUtime CPUtime DateTime DateTime ElapsedTime ElapsedTime nrotor nwing ntail ) ) System output) ) Aircraft ) ) ) ) ) ) ) ) ) ) ) ) ) ) ) Aircraft Propulsion Aircraft ) ) EngineParam Aircraft ) Aircraft ) Aircraft ) Aircraft ) Aircraft Aircraft Aircraft FuelTank Aircraft Aircraft 2 Aircraft Aircraft or kg/m Wing Aircraft Aircraft Aircraft Aircraft /π or m/sec 180 π/ 2 (slug/ft ρ (ft/sec g π number of fuel tank systems ( number of propulsion groups ( number of engine groups ( number of jet groups ( number of charge groups ( number of engine models ( number of engine tables ( number of reciprocating engine models ( number of compressor models ( number of motor models ( number of jet models ( number of fuel cell models ( number of solar cell models ( number of bettery models ( number of controls ( number of control states ( number of wing panels ( number of aux tank sizes ( number of drive system states ( number of trim states ( number of trim variables ( number of other fixed useful load categories ( number of engine ratings (for π 2 π/ degree/radian = radian/degree = gravity SLS density Job constants Case constants int int int int int int int int int int int int int int int int int int int int int int int real real real real real real real case _ case case case case _ case case case(ntankmax) case case _ _ _ _ _ case _ _ _ case case _ case _ _ _ case case _ _ case _ sls case(nwingmax) case control trim case(ntrimstatemax) _ case _ case(npropmax) _ case _ _ _ _ _ _ Common: Job ntank npropulsion nenginegroup njetgroup nchargegroup nenginemodel nenginetable nrecipmodel ncompressormodel nmotormodel njetmodel nfuelcellmodel nsolarcellmodel nbatterymodel ncontrol nstate npanel mauxtanksize ngear nstate mtrim nwoful nrate pi twopi halfpi degrad raddeg gravity density ) ) Dscale Dscale _ _ Units Units ; depending on ; depending on 3 2 / / 2 1 /lb /lb 2 2 =ft =ft k k from from 2 2 ; m ; m 3 2 / / ) 2 1 /kg /kg 2 2 =m =m D/q k k (ft/sec or m/sec) ) s c ; liter from m from from 2 2 or m (ft (ft 2 from input) 3 2 / / 2 1 or m from ft kW kW = = SLS speed of sound power (hp from ft-lb/sec; kW from m-N/sec) speed (knots from ft/sec or m/sec) range (nm from ft or m) weight (lb from lb; kg from N) mass (slug from lb; kg from kg) volume (gal from ft D/q D/q velocity (knots from input) altitude (ft or m from input) payload (lb or kg from input) time (minutes from input) distance (nm from input) drag (ft rate of climb (ft/sec or m/sec from input) power (hp from hp or kW) weight (lb from lb or kg) force (lb from lb or N) length (ft from ft or m) area (ft fuel (gal from gal or liter) slug (slug/lb or kg/kg) inches (in/ft or m/m) power (kW from hp or kW) meter (m from ft or m) weight (kg from lb or kg) volume (liter from gal or liter) energyflow (MJ/hr from hp or kW) Conversion factors Conversion factors for scaled Conversion factors for mission and flight condition input Conversion factors for weight equations Conversion factors for energy real real real real real real real real real real real real real real real real real real real real real real real real real real real real real hp lb frc ft ft2 gal slug in kW m sls _ vel alt pay time dist drag ROC _ _ _ _ _ _ _ _ _ _ kg L dE _ _ _ _ _ _ _ _ _ _ Common: Job csound powerconv knotsconv nmconv weightconv massconv volumeconv DoQconv23 DoQconv12 uconv uconv uconv uconv uconv uconv uconv wtconv wtconv wtconv wtconv wtconv wtconv wtconv wtconv wtconv wtconv Econv Econv Econv ) 99) /.

ln(1 ) 99) /.

ln(1 vel _ ROC _ ) uconv ) uconv Cases dist or kg/m _ ) uconv pay _ alt time drag _ _ _ Cases from lb/ft uconv uconv uconv uconv ); , m disk loading (lb/ft ton (from lb or kg) range for fuel=1%GW (nm from 1/(lb/hp-hr) or 1/(kg/kW-hr), times endurance for fuel=1%GW (min from hr, times write fuel energy for burn weight analysis units (from mission units, temperature (from speed (knots, mph, kph, ft/sec, m/sec); rate of climb (ft/min, ft/sec, m/sec); distance (nm, mile, km); time (min, hr); drag (ft payload (lb, kg); altitude (ft, m); length area velocity temperature weight power fuel flow mass flow sfc thrust sfc specific range fuel efficiency productivity force moment dynamic pressure density Conversion factors Output Units for output real real real real int int int c*10 c*10 c*10 c*10 c*10 c*10 c*10 c*10 c*10 c*10 c*10 c*10 c*10 c*10 c*10 c*10 c*10 c*10 c*10 c*10 c*10 c*10 c*10 c*10 case _ temp _ Common: Job DLconv tonconv rangeconv endconv WRITEenergy Uwrite Uwrite Ukts UROC Udist Utime UDoQ Upay Ualt Ulen Uarea Uvel Utemp Uwt Upwr Ufuelflow Umassflow Usfc Utsfc Uspecrange Ufueleff Uproductivity Ufrc Umom Uque Udens Default Cases Size Aircraft for Design Conditions and Missions Mission Analysis Flight Performance Analysis Map of Engine Performance Map of Airframe Aerodynamics Solution Procedures Cost Emissions Aircraft Systems Fuselage Landing Gear Rotors Wings Tails Fuel Tank Systems Propulsion Groups Engine Groups Jet Groups Charge Groups Engine Models Engine Tables Reciprocating Engine Models Compressor Models Motor Models Jet Models Fuel Cell Models Description Type Cases Size OffDesign Performance MapEngine MapAero Solution Cost Emissions Aircraft Systems Fuselage LandingGear Rotor Wing Tail FuelTank Propulsion EngineGroup JetGroup ChargeGroup EngineModel EngineTable RecipModel CompressorModel MotorModel JetModel FuelCellModel Chapter 5 Structure: Design Variable Cases Size OffDesign Performance MapEngine MapAero Solution Cost Emissions Aircraft Systems Fuselage LandingGear Rotor(nrotormax) Wing(nwingmax) Tail(ntailmax) FuelTank(ntankmax) Propulsion(npropmax) EngineGroup(nengmax) JetGroup(njetmax) ChargeGroup(nchrgmax) EngineModel(nengmax) EngineTable(nengmax) RecipModel(nengmax) CompressorModel(nengmax) MotorModel(nengmax) JetModel(njetmax) FuelCellModel(nchrgmax) Solar Cell Models Battery Models SolarCellModel BatteryModel Structure: Design SolarCellModel(nchrgmax) BatteryModel(ntankmax) 1 1 1 0 0 2 1 1 Default Turn off all tasks to just initialize and check the model, including geometry and weights title subtitle subtitle subtitle notes identification time-date identification size aircraft for design conditions mission analysis flight performance analysis map of engine performance map of airframe aerodynamics selection (0 none, 1 all, 2 first case) TechFactors (0 for none) Geometry (0 for none) Case Description Case Tasks (0 for none) Write Input Parameters Description + + + + + + + + + + + + + + + + + Type c*100 c*100 c*100 c*100 c*1000 c*32 c*20 int int int int int int int int TechFactors Geometry _ _ engine aero _ _ input input input _ _ _ Size Mission Perf Map Map _ _ _ _ _ Chapter 6 Structure: Cases Variable title subtitle1 subtitle2 subtitle3 notes ident timedate TASK TASK TASK TASK TASK WRITE WRITE WRITE 0 0 0 0 0 0 0 1 2 5 0 1 0 1 0 1 0 0 34 ’ ’ ’ ’ ’ ’ ’ ’ ’ ’ ’ ’ ’ ’ ’ ’ ’ ’ ) ) ) ) AEROn ) ) ENGINEn AIRCRAFTn ) ) ) PERFn GEOMETRYn SOLUTIONn DESIGNn SKETCHn ERRORn design file performance file geometry file aircraft description file solution file (1 text, 2 binary) sketch file errors file design file ( performance file ( geometry file ( aircraft description file ( solution file ( sketch file ( engine performance file ( airframe aerodynamics file ( errors file ( page control (0 none, 1 form feed, 2 extended Fortran) design (1 first case only, 2 all cases) weight statement, max level (1 to 5) weight statement, style (0 omit zero lines, 1 all lines) weight statement, component (0 for none) fuel energy for burn weight (0 for none) flight state, component loads (0 for none) design, performance, or geometry (1 single file of all cases) sketch component forces (0 none) sketch flight condition (0 none, 1 design, 2 performance) flight condition number selection (0 for none) file name or logical name (blank for default logical name) formats Output + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + int int int int int int int c*256 c*256 c*256 c*256 c*256 c*256 c*256 c*256 c*256 int int int int int int int int int int int load cond _ _ level long comp _ _ _ page design wt wt wt energy flight files sketch sketch _ _ _ _ _ _ _ _ _ _ design perf geometry aircraft solution sketch error design perf geometry aircraft solution sketch engine aero error _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ Structure: Cases OUT OUT OUT OUT OUT OUT OUT FILE FILE FILE FILE FILE FILE FILE FILE FILE WRITE WRITE WRITE WRITE WRITE WRITE WRITE WRITE WRITE WRITE ksketch 0 1 0 0 0 0 0 0 0 0 0 0. 0.

) error=1 _ OUT ) produces airframe aerodynamics file ) or separate file ( aero _ = case number in default name) n Map error=0 ) _ _ 2 ) produces engine performance file OUT , 2 m TASK g engine ; 1 ft _ Map _ or m TASK velocity units (0 knots; 1 mile/hr, 2 km/hr, 3 ft/sec, 4 m/sec) altitude units (0 ft or m; 1 ft, 2 m) payload units (0 lb or kg; 1 lb, 2 kg) time units (0 minutes; 1 hours) distance units (0 nm; 1 miles; 2 km) temperature (0 F or C; 1 F, 2 C) drag units (0 ft rate of climb units (0 ft/min; 1 ft/sec, 2 m/sec) option single file of all cases for design, performance, or geometry (form feed between cases) same information as in standard output, in tab-delimited form flight condition required to use Euler angles, control and incidence, component forces override default units (0 no, 1 yes) selected files are generated for each case ( size and analysis tasks can produce design and performance files aircraft or solution file can be read by subsequent case or job geometry file has information for graphics and other analyses sketch file has information to check geometry and solution (DXF format) engine map task ( airframe aerodynamics map task ( error messages to standard output ( specification (0 standard, 1 input) input gravitational acceleration reference density (0. for air at SLS) reference speed of sound (0. for air at SLS) analysis units (1 English, 2 SI) units for input of missions and flight conditions Gravity Environment Units + + + + + + + + + + + + + + + + + + int real real real int int int int int int int int int int ref ref _ _ miss vel alt pay time dist temp drag ROC grav _ _ _ _ _ _ _ _ _ _ Structure: Cases SET grav density csound Units Units Units Units Units Units Units Units Units Units xxx _ Units structure Aircraft scaled with gross weight (0 analysis default, 1 English, 2 SI) D/q English: ft-slug-sec-F; weights in lb, power in hp (internal units) SI: m-kg-sec-C; weights in kg, power in kW (internal units) pounds converted to slugs using reference gravitational acceleration speed in knots, time in minutes, distance in nm, ROC in ft/min input SizeCondition SizeMission OffMission PerfCondition Analysis units: must be same for all cases in job Weight in the design description is actually mass Default units for flight condition and mission: override with case input of other structures recorded in there must be input for systems, fuselage, landing gear, fuel tank there must be input for all structures used units for parameters Cases Size OffDesign Performance MapEngine MapAero Solution SizeCondition SizeMission OffMission PerfCondition Input for case Last input + + int int int int int int int int int int int int int int int int Dscale _ Structure: Cases Units inCases inSize inSizeCondition(nfltmax) inSizeMission(nmissmax) inOffDesign inOffMission(nmissmax) inPerformance inPerfCondition(nfltmax) inMapEngine inMapAero inSolution lastSizeCondition lastSizeMission lastOffMission lastPerfCondition Default conditions conditions missions Size Aircraft for Design Conditions and Missions Parameters Sizing Flight Conditions Design Missions Description Type SizeParam FltCond FltState Mission Chapter 7 Structure: Size Variable SizeParam FltCond(nfltmax) FltState(nfltmax) Mission(nmissmax) 0 0 0 ’jet’ ’miss’ ’ratio’ ’none’ Default ’engine’ ’f(DGW)’ ’f(DGW)’ ’WL+aspect’ ’DL+Vtip+CWs’ nMission or nFltCond ): size input specifies how fixed aircraft determined ): at least one Size=0 _ Size=1 _ that have zero power margin are not used to size engine or rotor that have zero torque margin are not used to size transmission = power from maximum of power required for all designated conditions and missions Cases%TASK = radius from maximum of power required for all designated conditions and missions = power required not used to size engine/rotor flight conditions and missions (max GW, max effort, or trim) Cases%TASK : ’engine’ ’rotor’ ’none’ perf _ size task ( no size task ( SIZE title notes quantity sized from performance jet group sized from performance charge group sized from performance parameter iteration (0 not required) rotor parameters wing parameters design gross weight (0 calculated, 1 fixed) weight empty (0 calculated, 1 fixed) fuel tank capacity structural design gross weight maximum takeoff weight drive system torque limit Size Aircraft for Design Conditions and Missions Sizing Method Description + + + + + + + + + + + + + + + + Type c*100 c*1000 c*16 c*16 c*16 int c*32 c*16 int int c*16 c*16 c*16 c*16 ds(npropmax) _ perf(npropmax) jet(njetmax) charge(nchrgmax) param rotor(nrotormax) wing(nwingmax) tank(ntankmax) SDGW WMTO limit _ _ _ _ DGW WE _ _ _ _ _ _ _ _ Chapter 8 Structure: SizeParam Variable title notes SIZE SIZE SIZE SIZE SET SET FIX FIX SET SET SET SET ), ’panel’ or ’hub’ or ’width’ or ’sigma’ , ’sigma’ , for one or more main rotors ’radius’ ’Vtip’ , or ) ’Vtip’ , ’sigma’ ’CWs’ , ’ratio’ thrust ), _ ’CWs’ or ’Vtip’ ), , ’ratio’ ’span’ or DESIGN ’scale’ ’CWs’ ), ( or ’DL’ ’WL’ parameters ’ratio’ GW or _ = input two of SET ’DL’ , or disk loading or ’aspect’ rotor perf _ _ or wing loading to calculate radius from radius of another rotor to calculate tail rotor radius from parametric equation, SET ’radius’ ’radius’ SIZE : ’area’ to calculate span from rotor radius to calculate span from rotor radius, fuselage width, and clearance (tiltrotor) to calculate span from wing panel widths to calculate span from span of another wing to calculate span from rotor hub position (tiltrotor) , aspect ratio for drive state #1 ref _ perf=’rotor’ ’chord’ _ ’ratio+XX+XX’ ’scale+XX+XX’ ’ratio+XX’ ’radius+XX’ ’width+XX’ ’hub+XX’ ’panel+XX’ Vtip that have zero thrust margin are not used to size jet group = input three of ( = = input three of ( = = input two of ( = = = = = = power from maximum of power required for all designated conditions and missions SIZE : use to force parameter iteration = thrust required not used to size jet group = power required not used to size charge group : = thrust from maximum of thrust required for all designated conditions and missions flight conditions and missions (max GW, max effort, or trim) , rotor parameters: required for each rotor rotor rotor rotor rotor wing wing wing wing wing wing : input DGW restricts , wing parameters: for each wing; input two quantities, other two derived _ _ _ _ _ _ _ _ _ _ using main rotor radius and disk loading : : fixed weight empty obtained by adjusting contingency weight param’ ’jet’ ’none’ ’charge’ ’none’ jet charge _ rotor wing _ _ SET except if SET tip speed is SET SET thrust from designated sizing conditions and missions ( SET SET SET SET SET SET DGW WE _ _ _ _ SIZE SIZE ’SIZE SET rotor parameters: input three or two quantities, others derived rotor parameters for an antitorque or aux thrust rotor: SET FIX FIX Structure: SizeParam cap − cap fuel − =off) W fuel * ) ) W Plimit + _ xmsn xmsn _ _ ) ) DGW _ fFuelSDGW wmto + sdgw _ _ ) DESIGN DESIGN Wfuel xx – _ DGW FltState%SET D _ W DESIGN DESIGN (energy) = fPlimit Wfuel – G cap ) D _ W , W G xmsn = Efuel _ W * G SD W , M T O W D G * DESIGN W W * W fWMTO D * ) or calculate (from * + (weight) or *((maximum mission fuel)+(reserve fuel)) W xx _ * ) cap cap fWMTO _ _ + eng fSDGW es fWMTO + fSDGW P _ Plimit + + *(maximum mission fuel), (maximum mission fuel)+(reserve fuel)) dWMTO fSDGW fFuel = fFuelSDGW eng Wfuel + , limit cap + ) N ) _ _ ( , rotor limits only use power required (or input) dWMTO dSDGW av req M T O rs = cap ∑ = dSDGW _ P P _ dWMTO = W fFuel ds = fWMTO cap dSDGW fSDGW _ SD , eng eng _ , = SD M T O limit W N N _ dFuel ( ( W W SD M T O = = max( Efuel ∑ ∑ fPlimit W W ) ) or dSDGW dWMTO cap cap _ _ prim prim cap EngineGroup%SET Ω Ω _ Rotor%SET / / Efuel Efuel input ref ref or or Wfuel ds (Ω (Ω = calculate from mission fuel used and mission battery discharge power _ cap cap ds ds _ _ _ _ = based on WMTO; = based on SDGW; = function of mission fuel used = based on DGW; = based on DGW; = based on fuel state; = based on maximum fuel; Plimit = input input parameters: , maximum takeoff weight: input parameters: , drive system torque limit: input (use = calculate from mission fuel used , structural design gross weight: Wfuel Wfuel = input = input = = from takeoff power, = from engine power required at transmission sizing conditions and missions ( ds = calculated from maximum gross weight at SDGW sizing conditions ( = calculated from maximum gross weight at WMTO sizing conditions ( = from engine power available at transmission sizing conditions and missions ( fPlimit fPlimit , fuel tank sizing: usable fuel capacity _ ’input’ ’miss’ ’miss+power’ ’f(miss)’ tank SDGW WMTO limit ’input’ ’f(DGW)’ ’f(WMTO)’ ’maxfuel’ ’perf’ Aircraft ’input’ ’f(DGW)’ ’f(SDGW)’ ’maxfuel’ ’perf’ Aircraft ’input’ ’ratio’ ’Pav’ ’Preq’ for transmission sizing conditions and mission segments ( _ _ _ _ SET SET SET SET engine shaft limit also uses rotor shaft limit also uses convergence may be improved if do not apply drive system limits to power available ( Structure: SizeParam perf=1 _ WMTO=’input’ _ DGW=1 SET _ , Cases%TASK , FIX ) , span panel=’width+taper’,’span+taper’ _ _ not ) mission=1 ) ) , _ SDGW=’input’ ) _ perf Preq perf , , panel , charge=’none’ SET radius , _ , _ ) Wing%SET wing=’area+span’ Pav ) , _ , hub not , SIZE , maxfuel none maxfuel , ) , Cases%TASK , SET fmiss , ratio , , , ) , scale width ) ds=’input’ , or none , , _ ) , param none input , fSDGW size=0 , area _ fWMTO , _ ratio _ jet=’none’ limit none , , radius _ _ , param , misspower rotor not limit , , , DL , _ wing=’WL+panel’ thrust , SET , fDGW _ SIZE , ratio fDGW , , WL param chrg miss , , SET _ , , , SET engine radius _ radius jet input _ ACTION=’nosize’ _ _ span area _ Cases%TASK input _ _ input _ perf jet charge _ _ _ _ rotor rotor _ _ wing wing perf=’none’ rotor=’radius+Vtip+sigma’ tank=’input’ _ tank _ _ _ _ _ WMTO SIZE SIZE SIZE SDGW _ SET _ (1 fixed, 0 not) SIZE SET SET SIZE SET SET SET (1 fixed, 0 not) /σ SET SET (1 fixed, 0 not) W tip turn off sizing: fix aircraft: use with wing panels: C V σ input required to transmit sized rotorcraft to another job (through aircraft description file) or to following case: performance ( performance ( performance ( rotor sized ( rotor radius ( rotor rotor rotor wing area ( wing span ( wing chord (1 fixed, 0 not) wing aspect ratio (1 fixed, 0 not) fuel tank ( SDGW ( WMTO ( drive system torque limit ( conditions and missions for size engine or rotor conditions and missions for size jet group conditions and missions for size charge group design conditions and missions for DGW design conditions and missions for transmission Specification Number of conditions and missions int int int int int int int int int int int int int int int int int int int int int radius(nrotormax) area(nwingmax) span(nwingmax) ds(npropmax) _ CWs(nrotormax) Vtip(nrotormax) sigma(nrotormax) _ _ _ GW xmsn(npropmax) chord(nwingmax) AR(nwingmax) _ _ _ _ _ _ _ engine(npropmax) jet(njetmax) charge(nchrgmax) perf(npropmax) jet(njetmax) charge(nchrgmax) rotor(nrotormax) _ _ _ rotor wing wing tank(ntankmax) SDGW WMTO limit _ _ _ _ _ _ _ _ _ _ _ rotor rotor rotor wing wing _ _ _ _ _ Structure: SizeParam iSIZE iSIZE iSIZE iSIZE iSET FIX FIX FIX iSET iSET FIX FIX iSET iSET iSET iSET nSIZE nSIZE nSIZE nDESIGN nDESIGN 0 0 ) ) nfltmax nmissmax (max all sizing conditions and missions) (max all sizing conditions and missions) (max all sizing conditions and missions) (max all sizing conditions and missions) avP G avEG avCG avJG /P /P /P /T Plimit Wfuelcap WE reqP G reqEG reqCG reqJG P P T P design conditions for SDGW design conditions for WMTO design missions for fuel tank design conditions and missions for rotor thrust kind iteration, performance (0 none, 1 size engine or radius or jet group or charge group) kind iteration, parameters (0 none, 1 calculate parameters) converged (0 not) number of iterations, performance loop number of iterations, parameter loop total number of iterations error ratio, engine error ratio, jet error ratio, charge error ratio, rotor error ratio, DGW error ratio, error ratio, structural design gross weight error ratio, maximum takeoff weight error ratio, error ratio, thrust error ratio, ratio ratio ratio ratio number of conditions for output number of missions for output number of conditions (maximum number of missions (maximum Size aircraft Sizing Flight Conditions Design Missions + + + + int int int int int int int int int int real real real real real real real real real real real real real real real int int int int sdgw wmto tank thrust out out _ _ _ _ size param _ _ _ _ size param total _ _ _ engine(nengmax) jet(njetmax) charge(nchrgmax) rotor(nrotormax) DGW xmsn(npropmax) sdgw wmto tank thrust(nrotormax) WE iter iter _ _ _ _ _ _ _ _ _ _ _ _ _ Structure: SizeParam nDESIGN nDESIGN nDESIGN nDESIGN kind kind issizeconv count count count error error error error error error error error error error error Pratio(npropmax) PratioEG(nengmax) Jratio(njetmax) Cratio(nchrgmax) nFltCond nMission nFltCond nMission namelist variables are arrays namelist SizeMission FltState and SizeCondition variables) in MissSeg FltState variables) in , and FltState MissSeg , and FltCond MissParam input one condition ( input one mission ( all mission segments are defined in this namelist, so each variable gets one more dimension, first array index is always segment number Structure: SizeParam Default Mission Analysis Parameters Missions Description Type OffParam Mission Chapter 9 Structure: OffDesign Variable OffParam Mission(nmissmax) Default namelist variables are arrays OffMission FltState and variables) in MissSeg Mission=1 _ FltState ) , and Cases%TASK MissSeg nmissmax , MissParam mission analysis input required if input one mission ( all mission segments are defined in this namelist, so each variable gets one more dimension, first array index is always segment number title notes number of missions for output number of missions (maximum Mission Analysis Analyze mission Missions Description + + + + + Type c*100 c*1000 int int out _ Chapter 10 Structure: OffParam Variable title notes nMission nMission Default conditions conditions Flight Performance Analysis Parameters Performance Flight Conditions Description Type PerfParam FltCond FltState Chapter 11 Structure: Performance Variable PerfParam FltCond(nfltmax) FltState(nfltmax) Default namelist Perf=1 _ PerfCondition Cases%TASK variables) in ) FltState nfltmax and FltCond flight performance analysis input required if input one condition ( title notes number of conditions for output (including sweeps) total number of sweep conditions number of conditions (maximum Flight Performance Analysis Analyze performance Performance Flight Conditions Description + + + + + Type c*100 c*1000 int int int out _ total _ Chapter 12 Structure: PerfParam Variable title notes nFltCond nsweep nFltCond 1 1 0 0 2 0. 0.

’std’ 1000.

20000.

Default engine=1 _ Map _ Cases%TASK ) alt _ Units ( h first set second set minimum maximum increment baseline engine map only available for RPTEM model and reciprocating engine model (performance only) engine map input required if only performance parameters or only model parameters used title notes engine group independent variables (0 none, 1 altitude, 2 temperature, 3 flight speed, 4 engine speed, 5 power) output format (1 single header, 2 header for inner variable) atmosphere specification altitude Map of Engine Performance Identification Kind (1 performance, 2 model) Performance Description + + + + + + + + + + + + + + + + + Type c*100 c*1000 int int int int int c*12 real real real real header min max inc base _ _ _ _ _ map _ var(5) var2(5) atmos _ _ _ Chapter 13 Structure: MapEngine Variable title notes kEngineGroup KIND SET SET WRITE SET altitude altitude altitude altitude 2 2 0 0 0. 0. 0. 0. 5. .1 1. .1 1.

10. 50. 90.

100. 200. 110. 100.

) ) xxx _ xxx _ temp , temp , xxx _ xxx _ altitude altitude ) xxx _ ) temp altitude _ Units ( T Δ is temperature or temperature increment xxx for other options (polar, tropical, and hot days) _ ) vel temp _ atmos _ Units (1 rpm, 2 percent) (TAS, or temperature increment = standard day at specified altitude, plus temperature increment (use N = standard day at specified altitude, and specified temperature (use τ , atmosphere specification: V = standard day at specified altitude (use FltState%SET atmos determines whether ’std’ ’temp’ ’dtemp’ see _ minimum maximum increment baseline minimum maximum increment baseline minimum maximum increment baseline minimum maximum increment baseline independent variables: 1 to 5 variables, last is innermost loop; outer loop is always rating quantities not identified as independent variables fixed at baseline values SET temperature flight speed engine speed power required (1 power, 2 fraction of power available (0. to 1.+) IR suppressor system state (0 off, hot exhaust; 1 on, suppressed exhaust) installation losses (0 for none) + + + + + + + + + + + + + + + + + + + + + + real real real real real real real real int real real real real int real real real real int int min max inc base _ _ _ _ IRS _ min max inc base loss min max inc base _ _ _ _ min max inc base _ rpm power _ _ _ _ _ _ _ _ _ _ Structure: MapEngine temp temp temp temp Vkts Vkts Vkts Vkts SET Nturbine Nturbine Nturbine Nturbine SET power power power power STATE KIND 1 1 0. 0. 1. .8 5. .1 2. .1 50. 1.1 .02 90.

400. 110.

) model _ V and fN ) model _ ) V and model _ fP V ) and model _ theta theta C ) and ) vel P/P _ V xxx _ (percent) Units spec atmos T /T _ (TAS, N/N V SET vs power fraction and airspeed (use vs temperature ratio (use vs airspeed (use performance: fuel flow, mass flow, net jet thrust, optimum turbine speed turbine speed: power ratio vs turbine speed and airspeed (use power available: specific power, mass flow, power, fuel flow number (maximum 10) values minimum maximum increment number (maximum 10) values minimum maximum increment minimum maximum increment minimum maximum increment RPTEM model flight speeds temperature ratio engine speed, fraction static MCP power, engine model atmosphere ( number of independent variable sets Model Specification + + + + + + + + + + + + + + + + + + + + + int real real real real int real real real real real real real real real real int int int model var _ model(10) min max inc atmos _ _ _ _ _ _ model min max inc min max inc _ model(10) min max inc _ _ _ _ _ _ _ _ _ _ Structure: MapEngine nV V V V V ntheta theta theta theta theta fN fN fN fP fP fP kEngineModel iSET nSET 1 1 1 1 1 1 1 0 Default ’retract’ aero=1 _ Map _ lift(1) _ var Cases%TASK AEROn//load (each aux tank size) or auxtank N aero//load load=DRAG, SIDE, LIFT, ROLL, PITCH, YAW _ FILE one file for each filename= function of single independent variable = fuselage and landing gear tails wings rotors engines and fuel tank airframe aerodynamics map input required if multi-dimensional: generate 6 files of three-dimensional tables one-dimensional: generate 1 file of all six loads title notes kind (1 one-dimensional, 2 multi-dimensional) aerodynamic loads (0 for components off) aircraft control state landing gear state number of auxiliary fuel tanks Map of Airframe Aerodynamics Tables Operating Condition Description + + + + + + + + + + + + + + + Type c*100 c*1000 int int int int int int int c*12 int control LG _ _ table _ fuselage tail wing rotor engine _ _ _ _ _ Chapter 14 Structure: MapAero Variable title notes KIND SET SET SET SET SET STATE STATE Nauxtank(nauxtankmax,ntankmax) 1 1 0 0. 0. 0. 0. 2. 5. 2. 2.

40. 45. 90. 20. 20. 60. 60.

180.

) ret , ext ) c + AC T c (keyword = = c ’retract’ , ; 1 for AC T c = c LG=’extend’ _ STATE α β low range increment (deg) high range increment (deg) low range value (deg) maximum value (deg) low range increment (deg) high range increment (deg) low range value (deg) maximum value (deg) low range increment (deg) high range increment (deg) low range value (deg) maximum value (deg) landing gear state: wing extension kit on aircraft (0 none, 1 present) angle of attack and sideslip angle representation (1 conventional, 2 reversed) use component control (0 for aircraft controls tilt angle of attack sideslip angle lift drag side force pitch moment roll moment yaw moment angle of attack and sideslip variation control variation third independent variable Independent variables Variable range + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + int int int real real real real c*16 c*16 c*16 c*16 c*16 c*16 real real real real real real real real real real real real control _ lowinc highinc low max lowinc highinc low max _ _ _ _ _ _ _ _ alpha lowinc highinc low max _ extkit comp _ _ _ _ _ _ lift(3) drag(3) side(3) pitch(3) roll(3) yaw(3) _ _ _ _ _ _ Structure: MapAero SET KIND SET control(ncontmax) tilt alpha beta var var var var var var angle angle angle angle control control control control gamma gamma gamma gamma +max ) and -max to control(ktilt) highinc replace ; by tilt +low to blank for 2D table , or control number) -low ) beta for , =one-dimensional load(3) ) _ ) ) retract alpha var , table lowinc control(ncontrol) , _ _ none naeromax extend _ KIND naeromax _ naeromax IDENT , LG naeromax VAR _ _ fixed if not independent variable ( variation: by ’beta’ used for , tilt / AERO STATE blank for 1D table, gamma lift(1) , _ ’alpha’ controls / var load(2) identify independent variables _ control beta / , only values: var load _ var alpha assume control system defined so aircraft controls connected to flaperon, elevator, aileron, rudder angle maximum total values = landing gear state ( number of independent variables variables (drag, side, lift, roll, pitch, yaw) number of angles (maximum angle values number of controls (maximum control values number of gamma (maximum gamma values Operating Condition Independent variables ( Tables int int int int real int real int real LG _ Structure: MapAero iSTATE nvar(6) ivar(3,6) nang ang(naeromax) ncnt cnt(naeromax) ngam gam(naeromax) 0 1 0 0 1 0 0. 0. 1. 0. 1. 0. 1. 0. 0. 0. 0. 0.

’pay’ ’DGW’ Default ) ) pay ) pay _ ) _ ) pay _ pay _ Units Units ( pass ) ( source N _ Units ( pay Units load pay _ ( − W G KIND ext ammo W Units ( W W weapons W cargo W number of passengers cargo external load ammunition weapons input gross weight gross weight increment gross weight factor power increment, each propulsion group power factor, each propulsion group thrust increment, each jet group thrust factor, each jet group kind (1 size mission, 2 size condition, 3 off design mission, 4 performance condition) mission or condition number segment number segment (1 start, 2 midpoint) input payload weight title label gross weight altitude (0 input, 1 from source for gross weight and altitude useful load Sizing or Performance Flight Condition Specification Description + + + + + + + + + + + + + + + + + + + + + + + + + Type c*100 c*8 c*12 real real real real real real real int int int int int c*12 real int real real real real source cargo extload ammo weapons _ _ _ _ _ GW alt UL _ _ _ source _ Chapter 15 Structure: FltCond Variable title label SET GW dGW fGW dPav(npropmax) fPav(npropmax) dTav(njetmax) fTav(njetmax) SET KIND kSource kSegment seg SET Wpay Npass Wpay Wpay Wpay Wpay 1 1 0 0 0 1 1 1 0 1 1 1 1 1 1 1 1 0. 1. -1 0. 0. 0.

-th size) m for ) m wmto=0 _ categories) DESIGN (each aux tank size) , , 2 only increase, 0 no change) seat − nWoful crew seat auxtank pass sdgw=0 − _ δN N δN Nauxtank crew δN DESIGN , GW=0 _ DESIGN tank size changed (–1 first, –2 first size already used, number tanks added or dropped number of crew increment crew seat increment passenger seat increment folding kit wing extension kit wing kit on aircraft other kit on aircraft (set fuel weight or energy increment fuel capacity factor auxiliary fuel tanks (1 adjust number of auxiliary fuel tanks crew weight increment other fixed useful load increment ( equipment weight increment kits on aircraft (0 none, 1 present) fixed gross weight conditions not used to determine DGW, SDGW, WMTO condition not used to size engine or rotor if power margin fixed (max GW, max effort, or trim) fuel tank system fixed useful load label is short description for output sizing flight condition: use all parameters except sweep design condition for power (1 to use for engine sizing) design condition for jet thrust (1 to use for jet group sizing) design condition for charge power (1 to use for charge group sizing) design condition for DGW (1 to use for DGW calculation) design condition for transmission (1 to use for transmission sizing) design condition for SDGW (1 to use for SDGW calculation) design condition for WMTO (1 to use for WMTO calculation) design condition for antitorque or aux thrust (1 to use for rotor sizing) + + + + + + + + + + + + + + + + + + + + + + + + + + + real real int int int int real int real real int int int int int int int int int int int int int int engine jet charge GW xmsn sdgw wmto thrust seat _ _ _ _ _ _ _ _ seat _ _ auxtank(ntankmax) foldkit extkit(nwingmax) wingkit(nwingmax) otherkit _ _ _ _ _ Structure: FltCond dFuel(ntankmax) fFuel(ntankmax) SET mauxtank(ntankmax) dNauxtank(ntankmax) Nauxtank(nauxtankmax,ntankmax) dWcrew dNcrew dWoful(10) dWequip dNcrew dNpass SET SET SET SET DESIGN DESIGN DESIGN DESIGN DESIGN DESIGN DESIGN DESIGN dPav + dTav ) av P cap + − av fPav source aux T _ = W ∗ fTav req KIND = P req T Nauxtank ∑ ) + cap ignored) − fuel UL _ ) or calculated (may depend on DGW) , W ); payload weight fallout SET (may depend on DGW) dGW cap dGW DGW + M T O _ + − , over all propulsion groups SD , over all jet groups dGW W fuel FIX + W xx source _ M T O W dGW ) = 0 SD Nauxtank ∗ W ) = 0 + W , * * W D ); fuel weight fallout * same (actual ; input ( W fFuel reqP G * , over all propulsion groups, engine groups, and rotors fFuel D fGW reqJG fGW + DESIGN P , UL fGW T _ W Wpay − = maximum GW for most restrictive of power, torque, and thrust margins − fGW ) ) = 0 ) dFuel d dFuel SET d : req + G ) + P W − GW = min( = maximum GW for power required equal specified power and transmission torque equal limit ’maxPQJ’ avP G avJG , , assume limit fuel f P f T = maximum GW for power required equal specified power: P values determine which input parameters used W = function WMTO: = function SDGW: = function of source: = input payload and fuel weights; gross weight fallout = input payload and fuel weights; gross weight fallout = function DGW: UL = maximum takeoff gross weight ’maxQP’ ’maxQJ’ = structural design gross weight _ , min(( min( most restrictive of power and torque margins min(( = gross weight from specified mission segment or flight condition ( ’max’ , = maximum GW for transmission torque equal limit: zero torque margin = design gross weight = input (use = maximum GW for jet thrust required equal specified thrust: , , source for gross weight or altitude: source must be solved before this condition = input fuel weight ( = input payload weight ( SET , , set gross weight , set useful load: with fixed useful load adjustments in fallout weight GW=’pay+fuel’ _ source ’DGW’ ’SDGW’ ’WMTO’ ’f(DGW)’ ’f(SDGW)’ ’f(WMTO)’ ’input’ ’source’ ’f(source)’ ’maxP’ ’maxQ’ ’maxPQ’ ’maxJ’ ’maxPJ’ ’pay+fuel’ ’pay’ ’fuel’ ’pay+fuel’ _ GW GW UL condition not used to size transmission if zero torque margin (max GW, max effort, or trim) _ _ _ SET performance flight condition: not use SET SET SET if KIND calculation order: size missions, size conditions, off design missions, performance conditions input fuel weight: Structure: FltCond 0 0 1 ignored inc _ sweep dWequip , ) saved (last value executed) first ); sign of _ of other quantities not used dWcrew , UL=’pay’ first inc _ _ _ sweep > 0 fFueln SET , sweep - sweep nsweepmax last auxtank _ dFueln _ , quantities SET sweep Wpay , sweep _ = no details ) first fTavn ) _ require non-zero weight per seat , = no details nquant seat sweep _ = parameter name dTavn used for fallout fuel weight ( , to Wpayload Wcrew _ nsweepmax _ qsweepmax last sweep dNpass _ fPavn SET _ , SET auxtank _ and sweep quant dPavn for first fuel tank system with SET used if seat , used if _ fGW (first value executed) should be condition that will converge of first quantity determines number of values, Wpay , dNcrew dWcrew last inc Nauxtank : control/pitch/roll values of trim iteration initialized from previous condition of sweep _ _ dGW , sweep adjust otherwise number of auxiliary fuel tanks fixed at input value maximum total number of values for all conditions is sweep analyzed using single data structure, only solution for sweep sign of parameter step determined by sign of ( sweep GW _ first parameter value last parameter value parameter increment parameter values auxiliary fuel tanks: payload: only crew: only equipment: Parameter sweep: only for performance flight conditions, not sizing flight conditions Single sweep, simultaneously varying Sweeps executed from INIT Available parameters: sweep (0 for none, 1 from list, 2 from range) initialize trim (0 for not) number of swept quantities (1 to list, number of values (maximum quantity (parameter name) range list Parameter sweep + + + + + + + + + + + + int int int int c*12 real real real real sweep _ first(qsweepmax) last(qsweepmax) inc(qsweepmax) sweep(qsweepmax) _ _ _ sweep _ sweep _ _ Structure: FltCond SET INIT nquant nsweep quant sweep sweep sweep sweep(nsweepmax,qsweepmax) ) pullup _ maxPQJ nz , , pullup maxQJ _ , rate fCharge , , maxPJ , turn _ ), fuel tank system, control number, or trim quantity; targetn fThrust maxJ _ , bank , wmto=0 , fTav _ , trim , turn _ fPower linear dTav maxPQ , _ nz , , DESIGN nz , HAGL dWLcg , , , Npecn turn maxQ _ , , linear _ rate dBLcg sdgw=0 ), jet group ( for all charge groups , , for all jet groups _ ny Vtipn viscosity for all engine groups, all propulsion groups maxP , , , _ roll fPav , tilt , , dSLcg GW linear _ , ) _ fCharge fThrust DESIGN fPower csound pitch dPav , nx , , pay , pedal , SET _ , side payfuel , , GW=0 linear density Nspec _ _ DoQV , ) , lngcyc , fuel az , , climb , ) xxx Vtip , _ pay temp _ pay , _ latcyc DESIGN GW linear , ROC _ _ UL , _ fDoQ ay coll , dtemp , , SET , , value is factor on input SET , value is factor on input Performance Mach pay , value is factor on input _ , linear , 2 _ Vkts ax altitude controln DoQ Size fPower fThrust fCharge number = propulsion group ( n 1 if absent for for for gross weight ( max gross weight (0 no iteration; useful load ( power margin as quantity (3 max GW, 2 max effort, 1 trim); not used to size engine or rotor torque margin as quantity (3 max GW, 2 max effort, 1 trim); not used to size transmission jet thrust margin as quantity (3 max GW, 2 max effort, 1 trim); not used to size jet group charger power margin as quantity (1 trim); not used to size charge group battery power margin as quantity (2 max effort, 1 trim); not used to size fuel tank fixed gross weight; quantity number quantity column label parameter values fraction of rated engine power available fraction of rated jet thrust available fraction of rated charger power available parent (1 FltCond performance output column (first for sweep) Specification Parameter sweep int int int int int int int int int int int int int c*8 real real real real original(nchrgmax) original(njetmax) sweep(qsweepmax) original(nengmax) _ _ _ _ GW GW maxGW UL sweep _ out _ _ _ _ _ Structure: FltCond parent kFltCond kcol iSET iSET iSET iSETPmargin(npropmax) iSETQmargin(npropmax) iSETJmargin(njetmax) iSETCmargin(nchrgmax) iSETBmargin(ntankmax) isFIX kquant label vsweep(nsweepmax,qsweepmax) fPower fThrust fCharge Default mission segments flight conditions Mission Profile Parameters Mission Segments Description Type MissParam MissSeg FltState Chapter 16 Structure: Mission Variable MissParam MissSeg(nsegmax) FltState(nsegmax) 0 0 1 1 1 0. 0. 1. 0. 0. 0. 0. 0. 0. 1. -1 ’delta’ Default ’pay+miss’ ’pay+miss’ -th size) m for m ) (each aux tank size) pay , 2 only increase, 3 increase at start and drop, 0 no change) _ ) auxtank Units ( pay ) N _ ) Nauxtank pay pay _ pay _ Units W ( pass ) G N Units W ( pay Units load _ ( − ext ammo Units ( W W weapons W cargo W tank size changed (–1 first, –2 first size already used, number tanks added or dropped number of passengers cargo external load ammunition weapons payload changes mission fuel weight (0 calculated, 1 fixed) fuel weight or energy increment fuel capacity factor auxiliary fuel tanks (1 adjust number of auxiliary fuel tanks folding kit on aircraft (0 none, 1 present) input gross weight gross weight increment gross weight factor input takeoff payload weight fuel tank systems fixed useful load title label mission takeoff gross weight useful load Mission Profile Specification Description + + + + + + + + + + + + + + + + + + + + + + + + + + Type c*100 c*8 c*16 real real real c*16 real int real real real real c*16 int real real int int int int int cargo extload ammo weapons _ _ _ _ GW UL pay auxtank(ntankmax) foldkit missfuel(ntankmax) _ _ _ _ _ _ Chapter 17 Structure: MissParam Variable title label SET GW dGW fGW SET Wpay Npass Wpay Wpay Wpay Wpay SET FIX dFuel(ntankmax) fFuel(ntankmax) SET mauxtank(ntankmax) dNauxtank(ntankmax) Nauxtank(nauxtankmax,ntankmax) SET 1 1 1 1 1 1 1 1 0.

30. 10. 10.

100.

2000.

) GW=0 _ DESIGN ) or calculated (may depend on DGW) (may depend on DGW) dGW DGW M T O _ + SD dGW W FIX + W : MissSeg%SizeZZZ=0 M T O G dGW SD + W W * W D * ; input ( W * D fGW xx fGW _ W fGW ) ) ) ) dist alt _ DESIGN _ GW time res _ f Units Units Units values determine which input parameters used = function WMTO: = function SDGW: = function DGW: UL = maximum takeoff gross weight = structural design gross weight _ = design gross weight = input (use SET , , set mission takeoff gross weight ’DGW’ ’SDGW’ ’WMTO’ ’f(DGW)’ ’f(SDGW)’ ’f(WMTO)’ ’input’ GW GW fixed gross weight missions not used to determine DGW (set mission segment not used to size engine or rotor if power margin fixed (max GW, max effort, or trim) mission segment not used to size transmission if zero torque margin (max GW, max effort, or trim) mission segment not used for sizing if set _ _ fuel reserve fraction distance increment ( time increment ( altitude increment ( takeoff velocity increment takeoff height increment label is short description for output sizing mission: use all parameters off design mission: not use SET SET fuel reserve (1 fraction mission fuel, 2 fraction fuel capacity, 3 only mission segments) split segments design mission for power (1 to use for engine sizing) design mission for jet thrust (1 to use for jet group sizing) design mission for charge power (1 to use for charge group sizing) design mission for DGW (1 to use for DGW calculation) design mission for transmission (1 to use for transmission sizing) design mission for fuel tank (1 to use for fuel tank capacity) design mission for antitorque or aux thrust (1 to use for rotor sizing) + + + + + + + + + + + + + + + int real real real real real real int int int int int int int engine jet charge GW xmsn tank thrust _ _ _ _ _ _ _ inc inc reserve inc _ inc _ _ _ _ inc _ Structure: MissParam SET fReserve dist time alt VTO hTO DESIGN DESIGN DESIGN DESIGN DESIGN DESIGN DESIGN ) must xWpay and dPav + pay _ dTav av P + SET av fPav T = fTav ) req ) = P ) req ignored) T UL _ xWpay(seg1) – SET xWpay(seg1) MissSeg%xWpay / xWPay xWPay , with more than one fuel tank system = no details same (actual ); initial payload weight fallout ) (so payload is fallout), payload (from UL , over all propulsion groups _ , over all jet groups Wpay SET ’miss’ ’pay+miss’ ) = 0 Nauxtank Wpayload ) = 0 _ , or or , minimum gross weight of designated segments , minimum gross weight of designated segments , minimum gross weight of designated segments , minimum gross weight of designated segments ); fuel weight fallout (not use SET reqP G , over all propulsion groups, engine groups, and rotors fFuel reqJG P , T , assume Wpay − = maximum GW for most restrictive of power, torque, and thrust margins − MaxGW MaxGW MaxGW MaxGW ) ) = 0 ) UL=’fuel’ xWpay d dFuel used if UL=’miss’ d _ _ req + + P SET SET ’pay+miss’ − xWpay = maximum GW for power required equal specified power and transmission torque equal limit ’maxPQJ’ avP G or avJG , and limit and f P f T = maximum GW for power required equal specified power: P = input payload, fuel weight from mission; gross weight fallout = input payload, fuel weight from mission; gross weight fallout = input payload and fuel weights; gross weight fallout = input payload and fuel weights; gross weight fallout ’maxQP’ ’maxQJ’ , Wpay at mission segment min(( at mission segment min( at mission segment most restrictive of power and torque margins at mission segment min(( ’max’ , = maximum GW for transmission torque equal limit: zero torque margin = maximum GW for jet thrust required equal specified thrust: = value; payload = , only used for = increment; payload = (initial payload weight)+( = no changes = factor; payload = (initial payload weight)*( = fuel weight from mission; initial payload weight fallout = input fuel weight ( GW=’max’ = input payload weight ( _ , set payload changes: mission segment payload (use of , set useful load: GW=’pay+fuel’ _ SET ’maxP’ ’maxQ’ ’maxPQ’ ’maxJ’ ’maxPJ’ ’pay+fuel’ ’pay+miss’ ’pay’ ’fuel’ ’miss’ ’pay+fuel’ ’pay+miss’ ’none’ ’input’ ’delta’ ’scale’ UL pay missfuel _ _ _ SET SET if FIX SET when not be zero at the maximum GW segments payload: only Structure: MissParam 1 0 1 0. 0. 0. 0.

) ) cap − fuel fuel W E + res O f W or cap = − cap − aux pay W W fuel ∗ − W G res f W , then drop Nauxtank at start (fixed ∑ fuel E ) + or Nauxtank to exceed tank capacity cap − fuel Nauxtank fuel W fuel E or , W cap for each segment; or fuel ), fixed − W ) or fraction of fuel capacity ( fuel ), start with input ) Nauxtank with change in W burn ∗ Nauxtank , allow E fuel ’pay+miss’ W res fFuel nsegmax f or ’pay+fuel’ + or or dFuel ’miss’ with change in ), adjust burn = Nauxtank=0 ’fuel’ O W = UL W ’pay’ res _ f = min( = UL _ SET UL fuel _ ) , adjust SET W options: fixed; or adjust UL SET _ : maximum of fuel for designated reserve mission segments SET OffDesign auxtank _ increase at mission start, then constant; or increase at start, then drop , 2 reserve for fallout fuel weight, this is the initial value for the mission iteration SET for input fuel ( for mission fuel ( for fallout ( for all fuel tank design mission: and fraction of fuel ( _ Size number input fuel weight: auxiliary fuel tanks: SET input all mission segments as arrays in single mission namelist method (0 segment start, 1 segment midpoint, 2 trapezoidal) relaxation factor (mission fuel) relaxation factor (range credit) relaxation factor (max takeoff GW) tolerance (fraction reference) trace iteration (0 for none) number of mission segments (maximum Segment integration Mission iteration (supersede Solution input if nonzero) Mission Segments parent (1 Mission performance output column + + + + + + + + + + int real real real real int int int int int SegInt miss _ miss range gw miss _ out _ _ _ _ _ Structure: MissParam KIND relax relax relax toler trace nSeg parent kMission kcol ) maxPQJ , maxQJ , maxPJ , ) maxJ ) , scale paymiss , ) , maxPQ , to _ delta miss , , maxQ Wfuel , input , payfuel and , maxP GW=0 _ to none _ ) _ _ fuel , GW xxx _ pay GW _ _ burn pay _ DESIGN GW W SET _ SET UL , range credit, takeoff GW) _ SET SET Wfuel gross weight ( max gross weight ( number max gross weight segments useful load ( payload changes ( power margin as quantity (all mission segments); not used to size engine or rotor torque margin as quantity (all mission segments); not used to size transmission jet thrust margin as quantity (all mission segments); not used to size jet group charger power margin as quantity (all mission segments); not used to size charge group battery power margin as quantity (all mission segments); not used to size fuel tank fixed gross weight; number reserve segments number adjustable segments kind adjustable (0 none, 1 distance, 2 time) kind range credit (0 none, 1 all forward, 2 all backward, 3 both) number takeoff segments kind iteration (0 none, 1 calculate mission fuel, 2 adjust mission, 3 only range credit or integration) converged (0 not) number of iterations error ratio ( first solution (initialize takeoff gross weight (start of mission) gross weight (end of mission, excluding reserve segments; last non-reserve segment) gross weight (end of mission; last segment) takeoff fuel weight (start of mission) added fuel weight (fill/add/drop during mission) fuel weight (end of mission, excluding reserve segments; last non-reserve segment) fuel weight (end of mission; last segment) weight fuel burned Specification Segments Iteration Mission quantities int int int int int int int int int int int int int int int int int int int real int real real real real real real real real to(ntankmax) add(ntankmax) endmiss(ntankmax) end(ntankmax) maxGW miss GW GW maxGW UL pay _ _ _ _ _ _ miss(3) _ adjust range iter _ _ _ _ to endmiss end _ _ _ _ _ _ _ Structure: MissParam iSET iSET nSET iSET iSET iSETPmargin(npropmax) iSETQmargin(npropmax) iSETJmargin(njetmax) iSETCmargin(nchrgmax) iSETBmargin(ntankmax) isFIX nreserve nadjust kind kind ntakeoff kind ismissconv count error isFirstSol GW GW GW Wfuel Wfuel Wfuel Wfuel Wburn(ntankmax) cap − fuel E ) (1 + > fuel avCG E P ) or limit avJG T cap DS avP G ) (1 + − P P ) ) max > fuel P (min) ) ) W (1 + ) (1 + (1 + > − reqCG > 2 > (1 + P / (1 + > )) | reqJG > (nm) T reqP G reqP G batt )) burn ) P P ˙ fuel ) E | W res W res burn − W E W to + + − W ( (ton-nm/lb or ton-nm/kg) to / burn burn (ton-kt/lb or ton-kt/kg) W to W E ( / W (ton-kt/lb or ton-kt/kg) burn ( to ( O burn W (maximum of fraction or reserve segments) R/W (maximum of fraction or reserve segments) E/ ln( V /W V /W res res pay = burn R/ W E pay pay E W W W = EF = e = = , block time (min) p p RF E (nm) R weight reserve fuel calculated mission fuel weight ( takeoff fuel energy (start of mission) added fuel energy (fill/add/drop during mission) fuel energy (end of mission, excluding reserve segments; last non-reserve segment) fuel energy (end of mission; last segment) energy fuel burned energy reserve fuel calculated mission fuel energy ( exceed power available: any mission segment exceed torque available: any mission segment exceed jet thrust available: any mission segment exceed charger power available: any mission segment exceed fuel capacity: any mission segment exceed battery power: any mission segment endurance range air distance (nm) block speed (kts; range/endurance) range factor efficiency factor fuel efficiency productivity productivity available seat miles direct operating cost total fuel burned energy used carbon dioxide Total mission, excluding reserve segments Cost Emissions Trading Scheme (kg CO2, per mission) Weight of emissions (kg, per mission) real real real real real real real real real int int int int int int real real real real real real real real real real real real real real real o f _ _ miss(ntankmax) factor to(ntankmax) add(ntankmax) endmiss(ntankmax) end(ntankmax) miss(ntankmax) _ _ fuel energy _ _ _ _ _ factor eff _ _ _ _ CO2 _ Structure: MissParam Wres(ntankmax) Wfuel Efuel Efuel Efuel Efuel Eburn(ntankmax) Eres(ntankmax) Efuel exceedP exceedQ exceedJ exceedC exceedWf exceedB endurance range airdist blockspeed range end fuel productivity productivity ASM DOC ETS ETS ETS W - methane - ozone (long life) - ozone (short life) x x x x NO water vapor soot sulphates total total without AIC carbon dioxide NO NO NO water vapor soot sulphates aviation induced cloudiness Average Temperature Response (deg C) real real real real real real real real real real real real real real noAIC CO2 CH4 O3L O3S H2O soot SO4 AIC _ _ _ _ _ _ _ _ _ NOx H2O soot SO4 _ _ _ _ Structure: MissParam W W W W ATR ATR ATR ATR ATR ATR ATR ATR ATR ATR 0 0 0 0 0 0 0 0 0 0 0. 0. 0. 1. 0. 0. 0. 0. 1. 0. 1. 0.

’dist’ Default ) mWfuel , 6-7 fill/add below rWfuel pass δN ) pay _ Units ) dist ) _ time _ Units ( D Units number of passengers increment ( T reserve (0 for not) adjustable for flexible mission (0 for not) segment number for range credit (0 for no reassignment) ignore segment (0 for not) copy segment (source segment number) split segment (number segments; –1 calculated; 0 for not split) fuel increment fuel factor fuel weight or energy change threshold fraction threshold weight or energy maximize gross weight (0 not) power increment, each propulsion group power factor, each propulsion group thrust increment, each jet group thrust factor, each jet group payload weight change ( kind distance time segment segment fuel use or replace refuel (0 not, 1 fill all tanks, 2 add fuel, 3 drop fuel, 4-5 fill/add below gross weight useful load Segment definition Description + + + + + + + + + + + + + + + + + + + + + + + + + + + Type c*12 real real int int int int int int int real real int real real real int real real real real real int credit _ tank(ntankmax) refuel(ntankmax) _ _ Chapter 18 Structure: MissSeg Variable kind dist time reserve adjust range ignore copy split SET dTank(ntankmax) fTank(ntankmax) SET xWfuel(ntankmax) rWfuel(ntankmax) mWfuel(ntankmax) MaxGW dPav(npropmax) fPav(npropmax) dTav(njetmax) fTav(njetmax) xWpay xNpass 0 0 0 1 1 0 0 0 0 1 1 1 1 1 1 0. 0. 0. 0. 0.

) alt _ kSeg ), fuel burned but no distance added to range ) time time ) ) dist time categories) seat − nWoful crew seat *altitude) pass − *altitude) δN δN crew fWind + fWind δN + dWind : use or replace specified fuel amount, calculate time but no distance added to range dWind : takeoff distance calculation : fly segment for specified time (use : taxi/warm-up mission segment (use ’TO’ , ’charge’ ’loiter’ : climb/descend from present altitude to next segment altitude : climb/descend from present altitude to next segment altitude, fuel burned but no dist added to range ’idle’ , : fly segment for specified time (use , , : fly segment for specified distance (use : use or replace specified fuel amount, calculate time and distance number of crew increment crew seat increment passenger seat increment wing extension kit wing kit other kit kind crew weight increment other fixed useful load increment ( equipment weight increment kits on aircraft (0 none, 1 present) kind=’taxi’ kind=’dist’ kind=’time’ kind=’hold’ kind=’climb’ kind=’spiral’ kind=’fuel’ kind=’burn’ kind=’takeoff’ fixed useful load source of altitude wind increment, knots ( wind gradient, knots ( power jet thrust charger power DGW transmission antitorque or aux thrust segment altitude at start of segment (0 input, 1 from previous segment, 2 from wind specification (0 none, 1 headwind, 2 tailwind) design mission (0 to not use segment for sizing) + + + + + + + + + + + + + + + + + + + + + + + real int real real int int int int int int int int real real int int int int int int seat seat _ _ alt extkit(nwingmax) wingkit(nwingmax) otherkit alt wind _ _ _ _ _ _ Structure: MissSeg dWcrew dNcrew dWoful(10) dWequip dNcrew dNpass SET SET SET SET kSeg SET dWind fWind SizeEngine SizeJet SizeCharge SizeGW SizeXmsn SizeThrust can not be ignored xWfuel source _ inc _ alt , cap inc _ − FltCond%KIND cap , fuel time − E , * fuel less distances from other segments fuel inc credit kind=’time’/’hold’ E _ E _ * * credit dist fTank fuel _ + E or all range * fTank , fTank range + + ) dTank fTank ’burn’ or + dTank dTank MaxGW or refuel _ cap MissParam%dest kind=’dist’ or or − dTank SET cap fuel fuel or − of fuel capacity (including auxiliary tanks), fill all tanks of fuel capacity (including auxiliary tanks), add segment is specified , specified distance is for group of segments W W * * fuel fuel xWfuel kind=’fuel’ W W * * designations for each segment credit rWfuel rWfuel _ fTank fTank , fill all tanks , add + + kind=’dist’ fTank fTank credit , can adjust one or more segments range + + _ dTank dTank xWfuel mWfuel mWfuel xWfuel ’miss’ range dTank dTank , not (not based on keyword, increment always positive) UL adjust : fill all tanks (including any auxiliary tanks installed) : add fuel : drop fuel : if below fraction : if below fraction : if below : if below _ 0 : no requirement : target : target : increment : increment , segment must be SET ˙ E < reserve positive (add or drop determined by credit , refuel: change at start of segment; weight or energy _ tank = 0 tank = 1 tank = 2 tank = 3 tank = 4 refuel = 1 refuel = 2 refuel = 3 refuel = 4 refuel = 5 refuel = 6 refuel = 7 : segment fuel use or replace for _ _ _ _ _ _ _ _ _ _ _ _ actual distance flown in tank refuel range calculated for this segment credited to segment = range if credit to earlier segment, iteration required if more than one segment adjusted, must be all adjust time or distance based on fuel burn (proportional to initial values) SET SET SET SET SET charge if target limited by capacity, if target already achieved then no requirement increment limited by current fuel (use) or capacity minus current fuel (replace) SET SET SET SET SET SET SET added fuel limited by capacity; not used for first segment xWfuel _ _ only one of reserve: time and distance not included in block time and range range credit: to facilitate specification of range adjustable: for split segment: number specified, or calculated from ignore segment: removed from input; segments using SET SET Structure: MissSeg 0. 0. 0. 0.

1.5 2.0 1.2 0.04 ’none’ ), alt=1 ) _ fail SET or eng can not be climb or spiral ’maxPQ’ or kind (keyword = alt pullup=0 ’maxQ’ _ _ or SET , turn=SET takeoff=’start’ ’engine fail’ _ _ ), MaxGW segments , GW=’maxP’ SET SET run _ ), or altitude of previous segment ( ) or refuel ’brake’ SET = no details = 0 _ , alt=0 , or is _ V ( SET ground , SET ’climb’ superseded; kind=’takeoff’ , Wpayload credit _ _ require non-zero weight per seat HAGL = no details kind=’takeoff’ SET (keyword = range seat takeoff=’start’ ’transition’ , _ (sec) _ ) and , designate segments if t vel Wcrew used if _ _ SET adjust , dNpass (deg) MissParam%Npass SET SET γ ’rotation’ ’ground run’ , MaxGW xWpay and , reserve , T R and (ft or m) seat used if n _ h not allowed in maximum effort (+ for uphill; deg) μ time G , (sec) γ Wpay R dist is change from dNcrew t dWcrew takeoff=’start’ takeoff=’liftoff’ var=’alt’ _ _ _ begin altitude is that input for this segment ( xNpass first segment identified by last segment if next segment is not SET SET not use max velocity specification ( splitting liftoff or engine failure segment produces additional ground run segments maximize gross weight: climb/descend or spiral segment: end altitude is that of next segment; last segment payload: only crew: only equipment: takeoff distance calculation: set of consecutive takeoff segment kind each segment requires appropriate configuration, trim option, max effort specification can split segment (except start, rotation, transition): split height for climb, velocity for others takeoff segment kind ground speed or climb speed (knots, CAS) climb angle relative ground height during climb slope of ground friction coefficient decision delay after engine failure rotation time transition load factor Takeoff distance calculation + + + + + + + + + + c*12 real real real real real real real real takeoff takeoff _ ground takeoff takeoff _ _ _ _ transition _ decision rotation _ _ Structure: MissSeg SET Vkts climb height slope friction t t nz , fraction of engine failure segment segment fPower ) segment burn takeoff=’liftoff’ , _ for each segment fuel , SET specification must be consistent spiral , nEngInop takeoff=’liftoff’ _ nEngInop climb , SET hold , time , dist , ) after engine failure segment, use engine rating, taxi _ kind _ decision _ OffDesign t , 2 integrate acceleration vs velocity to obtain time and distance segments correspond to ends of integration intervals analysis checks for consistency of input velocity and calculated acceleration analysis checks for consistency of input height and input/calculated climb angle order: start, ground run, engine failure, ground run, liftoff, rotation, transition, climb only one liftoff; only one engine failure, rotation, transition (or none) engine failure before liftoff; all ground run before liftoff, all climb after liftoff order: start, ground run, engine failure, brake only one engine failure (or none) until so engine failure segment corresponds to conditions before failure if engine failure segment present, MissSeg Size number number separate definition of multiple ground run, climb, brake segments allows configuration variations define takeoff profile in terms of velocities takeoff distance definition: includes accelerate-stop distance definition: does not have engine failure segment (if present) identifies point for decision delay number of inoperative engines specified by kind ( folding kit on aircraft (0 none, 1 present) parent (1 Mission MissSeg performance output column Specification int int int int int int foldkit out _ _ Structure: MissSeg parent kMission kMissSeg kcol ikind SET ) maxGW + maxPQJ , maxQJ , maxPJ ) , maxJ , xWpay(seg1) / maxPQ , xWpay maxQ , maxP _ ) or factor ( GW ) _ ratio) xxx _ SET dist or xWpay(seg1) takeoff – _ time SET xWpay (ground axes) (ground axes) zG L D μF − − T W first segment after split last segment after split this segment receives range credit (0 not, 1 source forward, 2 source backward, 3 both) adjustment ratio (initial split segment (number segments; 0 for not split) payload increment ( max gross weight (0 no iteration; power margin as quantity (3 max GW, 2 max effort, 1 trim) torque margin as quantity (3 max GW, 2 max effort, 1 trim) jet thrust margin as quantity (3 max GW, 2 max effort, 1 trim) charger power margin as quantity (1 trim) battery power margin as quantity (2 max effort, 1 trim) converged (0 not) number of iterations error ratio gross weight increment takeoff segment kind ( ground speed or climb speed (CAS) ground speed (ft/sec or m/sec) angle relative ground (deg) consistent acc and V change, climb and h change net force net force friction drag acceration (ground axes) height (ft or m) time (sec) distance (ft or m) cumulative time (sec) cumulative distance (ft or m) Segments Maximum gross weight Takeoff distance calculation int real int int int real int int int int int int int int real real int real real real int real real real real real real real real real TO _ TO _ TO maxgw first last TO maxGW takeoff _ _ maxgw _ _ _ TO range TO TO _ inc _ _ TO TO _ _ _ _ _ _ _ TO TO TO TO _ _ _ _ Structure: MissSeg kind fadjust wassplit ksplit ksplit dWpay iSET iSETPmargin(npropmax) iSETQmargin(npropmax) iSETJmargin(njetmax) iSETCmargin(nchrgmax) iSETBmargin(ntankmax) ismaxgwconv count error GW iSET VCAS V climb isConsistent FxG FzG FzGmu acc h t s time dist ) ’spiral’ or kind=’climb’ ) FltAircraft ; at start or midpoint) h h (knots) V V c c (knots) ) V V h w (ft/sec or m/sec) ˙ ˙ V V E E c FltState (MJ/hr) V − ˙ ˙ w w ˙ h E (lb/hr or kg/hr) V ˙ (knots) ( at start of segment (ft or m) at end of segment (from start of next segment, only used for w w h h V engine rating engine rating fraction of rated engine power available jet rating jet rating fraction of rated jet thrust available charger rating charger rating fraction of rated charger power available friction coefficient horizontal speed climb velocity fuel flow energy flow horizontal speed climb velocity fuel flow energy flow original value for engine failure decision (from engine failure segment (0 for none) horizontal speed climb velocity fuel flow energy flow trapezoidal integration altitude altitude Performance (from Headwind Ground speed c*12 int real c*12 int real c*12 int real real int real real real real real real real real real real real real real real real real original(nchrgmax) original(nchrgmax) _ _ start(ntankmax) end(ntankmax) original(njetmax) _ _ original(njetmax) _ _ original(nchrgmax) start(ntankmax) end(ntankmax) original(njetmax) TO original start end original(nengmax) _ _ _ _ _ original(nengmax) jet charge _ _ _ _ start end original(nengmax) jet charge _ _ _ _ _ _ _ _ start end _ _ Structure: MissSeg rating krate fPower rating krate fThrust rating krate fCharge friction kSegEF speed Vclimb fuelflow(ntankmax) energyflow(ntankmax) speed Vclimb fuelflow energyflow speed Vclimb fuelflow energyflow alt alt Wind groundspeed (nm) (nm) (lb or kg) dR (MJ) (segment start) (segment start) (segment start) D G burn fuel burn fuel W W E W E (minutes) - methane - ozone (long life) - ozone (short life) T x x x x time ground distance distance from past range credit (nm) distance from future range credit (nm) range contribution air distance (nm) fuel burned fuel added at start of segment fuel weight fuel burned fuel added at start of segment fuel energy gross weight total fuel burned energy used carbon dioxide NO water vapor soot sulphates total total without AIC carbon dioxide NO NO NO water vapor soot sulphates Mission segment quantities Emissions Trading Scheme (kg CO2, per mission) Weight of emissions (kg, per mission) Average Temperature Response (deg C) real real real real real real real real real real real real real real real real real real real real real real real real real real real real real real add(ntankmax) start(ntankmax) add(ntankmax) start(ntankmax) _ _ noAIC CO2 CH4 O3L O3S H2O soot SO4 fuel energy _ _ start _ _ _ _ _ _ _ _ _ _ _ CO2 NOx H2O soot SO4 _ _ _ _ _ Structure: MissSeg T D otherDpast otherDfuture dR airdist Wburn(ntankmax) Wfuel Wfuel Eburn(ntankmax) Efuel Efuel GW ETS ETS ETS W W W W W ATR ATR ATR ATR ATR ATR ATR ATR ATR , input or turboshaft calculated, weighted for engine group ˙ w ∑ ˙ ˙ w w/ for EI to ∑ /P = q P x = NO P aviation induced cloudiness EI f real real real AIC _ NOx(ntankmax) _ Structure: MissSeg ATR EI fPto(nengmax) Default fuselage landing gear rotors wings tails fuel tank systems propulsion groups engine groups jet groups charge groups Flight State Aircraft Components Description Type FltAircraft FltFuse FltGear FltRotor FltWing FltTail FltTank FltProp FltEngn FltJet FltChrg Chapter 19 Structure: FltState Variable FltAircraft FltFuse FltGear FltRotor(nrotormax) FltWing(nwingmax) FltTail(ntailmax) FltTank(ntankmax) FltProp(npropmax) FltEngn(nengmax) FltJet(njetmax) FltChrg(nchrgmax) 0 0 1 1 0 0 0 1. 0. 0. 0. 0. 0. 0. 0. 0. 1. 0. 0. 1. 0. 0.

77 ’ ’ ’ ’ Default ’general’ input) input) ) FltState FltState vel ) ) _ 2 2 Units value, 1 ) value, 1 or m/sec or m/sec ROC 2 2 _ Aircraft Aircraft Units (ft/sec (ft/sec ( c (TAS or CAS, (Mach number) V h ACx ACy (deg) V M a a (deg) (g) (deg/sec) (g) (deg/sec) F (deg) V φ n n F F ˙ ψ ˙ θ V ψ θ F θ F φ pitch roll turn rate load factor bank angle pitch rate load factor x-acceleration y-acceleration quantity variable switch quantity if exceed limit (0 not, 1 power margin, 2 torque margin, 3 both) flight speed factor horizontal velocity horizontal velocity vertical rate of climb climb angle sideslip angle pitch motion specification (0 roll motion specification (0 turn specification (0 zero, 1 turn rate, 2 load factor, 3 bank angle) pullup specification (0 zero, 1 pitch rate, 2 load factor) linear acceleration specification (0 zero, 1 acceleration, 2 load factor) maximum effort performance (maximum 2, 0 to analyze specified condition) flight speed aircraft motion Flight State Specification Description + + + + + + + + + + + + + + + + + + + + + + + + + + + + Type int c*12 c*12 int real c*12 real real real real real int int real real int real real real int real real int real real turn max vel pitch roll turn pullup acc quant(2) var(2) limit(2) turn pullup _ _ _ _ _ _ _ _ _ _ _ _ _ turn pullup linear linear _ _ _ _ Chapter 20 Structure: FltAircraft Variable SET max max max fVel(2) SET Vkts Mach ROC climb side SET SET pitch roll SET rate nz bank SET rate nz SET ax ay 0 1 1 1 1 1 1 1 1 0 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0.

’std’ 999.

’default’ ) c + conversion schedule) AC input) T c = LG Aircraft c ) h FltState ) ; 1 for input, 2 temp _ or m/sec (g) (g) AC (g) value, 1 ACs T c Lx Ly Lz c Units n n n ( = FltState ) c T (ft/sec ACc Aircraft 0 Δ c temp AC _ ACz c a value, 1 s ) Units c ( alt τ _ ACp Aircraft c μ tilt ρ α Units ( z-acceleration x-load factor increment y-load factor increment z-load factor increment collective stick lateral cyclic stick longitudinal cyclic stick pedal tilt (0 collective stick lateral cyclic stick longitudinal cyclic stick pedal tilt h temperature temperature increment density speed of sound viscosity height of landing gear above ground level control specification (0 aircraft controls stationline buttline waterline altitude atmosphere specification ground effect (0 OGE, 1 IGE) landing gear state aircraft control state use component control (0 for center of gravity specification (0 baseline plus increment, 1 input) + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + real real real real real c*12 real real real real real int real c*12 int int int int int int int real real real real real real int int real real real control _ LG control _ _ atmos GE control(ncontmax) coll latcyc lngcyc pedal tilt comp cg _ _ _ _ _ _ _ _ _ _ linear linear linear linear _ _ _ _ Structure: FltAircraft az nx ny nz altitude SET temp dtemp density csound viscosity SET HAGL STATE STATE SET SET SET SET SET SET control(ncontmax) coll latcyc lngcyc pedal tilt SET SET dSLcg dBLcg dWLcg 1 1 1 0 1 0 0 2 0 0 2 0 0. 1. 1. 1. 0. 1. 1.

79 ’ ’ 20.

’MCP’ ’MCT’ ’MCP’ ’hover’ ) jet ) AT eng AP , 4 fixed av , 4 fixed AT av AP , 3 fixed (0. to 1.+) ) (0. to 1.+) A C ) P f max f , 3 fixed (0. to 1.+) chrg A CC T x f AP inop or inop N N acc , each fuel tank mbd inop eq dP x N , 4 fixed dP A at μ M tip speed tip speed from tip speed from engine speed (rpm) rotor tip speed specification drive system state drive system rating drive system limit (0 not applied to power available) rotor shaft limit (0 not used for torque margin) accessory power increment engine rating fraction of rated engine power available number of inoperative engines power required (1 distributed, 2 fixed IR suppressor system state (0 off, hot exhaust; 1 on, suppressed exhaust) jet rating fraction of rated jet thrust available number of inoperative jets thrust required (1 from component, 2 fixed IR suppressor system state (0 off, hot exhaust; 1 on, suppressed exhaust) charger rating fraction of rated charger power available number of inoperative chargers power required (2 fixed battery capacity fade factor cell temperature (deg C) maximum current (fraction Specification, each propulsion group Specification, each engine group Specification, each jet group Specification, each charge group Equipment power increment Specification, each fuel tank (battery) Deice system state (0 off) + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + c*12 real real real real int c*12 int int real c*12 real int int int c*12 real int int int c*12 real int int real real real real int jet(njetmax) rs(npropmax) _ _ gear(npropmax) IRS(nengmax) IRS deice _ _ _ _ ds(npropmax) jet(njetmax) charge(nchrgmax) _ _ _ Vtip(npropmax) Plimit(npropmax) Qlimit Preq(nengmax) Jreq(njetmax) Creq(nchrgmax) Vtip(npropmax) _ _ _ _ _ _ Vtip(npropmax) _ _ Structure: FltAircraft SET Vtip(npropmax) mu Mat Nspec(npropmax) STATE rating SET SET dPacc(npropmax) rating(nengmax) fPower(nengmax) nEngInop(nengmax) SET STATE rating fThrust(njetmax) nJetInop(njetmax) SET STATE rating fCharge(nchrgmax) nChrgInop(nchrgmax) SET dPeq(ntankmax) ffade(ntankmax) Tcell(ntankmax) fcurrent(ntankmax) STATE 0 0 0 0 0 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0.

’none’ ) pay _ Units , drag _ Units ( pay ) /W ) drag _ ) D/q for no trim) Units Δ( drag ( _ D/q ’none’ Units ( , trim D/q _ input if nonzero) IDENT κ d c Solution all rotors trim maximum effort maximum gross weight all rotors trim maximum effort maximum gross weight trim maximum effort trim maximum effort maximum gross weight maximum effort maximum gross weight payload forward flight drag increment payload drag increment scaling with weight payload vertical drag increment induced power factor profile power mean inplane forces, tip-path plane axes (1 neglect, 2 blade-element theory) inplane forces, profile (1 simplified, 2 blade element theory) control mode (1 thrust and TPP, 2 thrust and NFP, 3 pitch and TPP, 4 pitch and NFP) aircraft trim state (match trim quantity targets relaxation factor tolerance (fraction reference) reinitialize aircraft controls (0 no, 1 force retrim) variable perturbation amplitude (fraction reference, 0. for no limit) maximum derivative amplitude (0. for no limit) Performance Rotor (nonzero to supersede rotor model) Trim solution Iterations (supersede + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + real real real real real int int int c*12 real real real real real real real real real int int real real real real real fly(2) maxgw Ftpp(nrotormax) Fpro(nrotormax) _ _ _ _ trim trim fly(2) maxgw pay _ _ _ _ control(nrotormax) pay _ pay _ rotor trim fly(2) maxgw rotor trim fly(2) maxgw _ target(mtrimmax) _ _ _ _ _ _ _ _ _ trim fly _ _ _ Structure: FltAircraft DoQ fDoQ DoQV Ki(nrotormax) cdo(nrotormax) MODEL MODEL KIND STATE trim relax relax relax relax toler toler toler toler init init perturb perturb perturb maxderiv maxderiv 0 0 0 0 0 0. 0.

tran _ ROC=0 Rotor%CTs as outer loop uses ’speed’ / ’power’ maximum (1/fuelflow) 0.99 maximum (V/fuelflow) maximum (V/fuelflow) 0.99 maximum (V/fuelflow), low side maximum (ROC) quant=’rotor(t) n’ Vtip=’input’ for convergence _ _ ) max , n=1 as inner loop, fly=0.1 quant _ _ steady ’alt’ ’ROC’ / _ max ’end’ ’range’ ’range(100)’ ’range(low)’ ’climb’ or maxinc FltAircraft%SET ’Pmargin’ Rotor%CTs var=’speed’ : requires _ uses max ’Nspec’ or description endurance range (high side) range range (low side) climb or descent rate ” is absent, use first component ( n : switch quantity to power and/or torque margin if margin negative; useful for best range quant=’rotor(s) n’ var=’Vtip’ limit is only used for _ _ _ two variables must be unique two variables can be identified for same maximized quantity (endurance, range, climb) ROC or altitude can be outer loop quantity only if it is also inner loop variable fVel ceiling calculation should use best range calculation often requires ROC for zero power margin initialized based on level flight power margin if input max if energy burned (not weight) or multiple fuels, use equivalent fuel flow obtained from weighted energy flow max if trailing “ max maximum effort maximum gross weight maximum effort all rotors trim (2 for component controls) maximum effort maximum gross weight maximum effort performance: one or two quantity/variable identified; first is inner loop maximum increment fraction (0. for no limit) solution method trace iteration (0 for none) + + + + + + + + + + real real int int int int int flymax(2) fly(2) maxgw _ _ _ rotor trim fly(2) maxgw _ _ _ _ Structure: FltAircraft maxinc maxinc method trace trace trace trace (all fuel tanks) (transient) (sustained) (all limits) (all propulsion groups) (all jet groups) = 0 = 0 ) = 0 | ) = 0 fVel fVel /σ /σ ) = 0 T T req ) = 0 batt ˙ C C Q E = 0 req req times times sideslip angle Euler angle rates linear, airframe axes linear, inertial axes linear, ground axes body axes relative inertial axes − − P − T L − | C − − max max − limit ) ) av av max P Q T P /σ /σ max T T L C C ’zaccG’ maximum (1/Power) maximum (ROC/V) maximum (V/Power) maximum (altitude) min( min( min( most restrictive most restrictive most restrictive most restrictive min( ( ( C ’zaccI’ ’, control ’zacc’ _ ’, ’, , , ’turn’ , ’yaccG , ’yaccI ’roll’ , , , ’yacc IDENT , , , , var _ quant _ max ’speed’ ’ROC’ ’side’ ’alt’ ’pullup’ ’xacc’ ’xaccI’ ’xaccG’ match ’pitch’ ’Vtip n’ ’Nspec n’ max ’power’ ’angle’ ’power/V’ ’alt’ ’P margin’ ’Q margin’ ’J margin’ ’PQ margin’ ’PJ margin’ ’QJ margin’ ’PQJ margin’ ’B margin’ ’rotor(t) n’ ’rotor(s) n’ ’stall n’ description horizontal velocity vertical rate of climb aircraft velocity altitude aircraft angular rate aircraft acceleration aircraft acceleration aircraft acceleration aircraft control aircraft orientation propulsion group tip speed propulsion group engine speed description climb rate (power) climb or descent angle climb angle (power) ceiling power margin torque margin jet thrust margin power and torque margin power and thrust margin torque and thrust margin power, torque, thrust margin battery power margin rotor thrust margin rotor thrust margin wing lift margin Structure: FltAircraft ) F φ ) , V F ) ψ θ , , ) V F θ ψ climb dtemp tan( , h V = for flight speed altitude c V ) Vkts ) ) , climb ) ) side side for flight speed , , ) ) side side ) ) , Vkts ROC climb sin( ) ) side h ) , ROC altitude V altitude altitude climb=0/+90/–90 ) , = ROC s altitude side=90 V ; ; calculated using input side=–90 ; Vkts=0 ROC ; side=180 V /g , side , ) , ) ROC , ROC turn side Vkts =horizontal, ROC =velocity magnitude, , =velocity magnitude, =forward velocity, = ( Vkts , pitch and roll motion specification: ; calculated using input Vkts cos( Vkts Vkts h roll Vkts Vkts Vkts _ − not V V /g ) = n SET f √ Mach is CAS not TAS V pullup ( = specified altitude, plus temperature increment (use and values (perhaps function speed) or flight state input ) = , load factor in pullup: use pullup rate or load factor Vkts , bank angle and load factor in turn: use turn rate, load factor, or bank angle = use = general (use roll , linear acceleration: use acceleration or load factor = = general (use = hover (zero velocity) = general (use = general (use = polar day at specified altitude (use = right sideward (use = 1 + , atmosphere specification: = tropical day at specified altitude (use = hover or VROC (use = rearward (use PITCH turn pullup acc = left sideward (use = hot day at specified altitude (use = standard day at specified altitude (use _ _ _ _ sideslip positive aircraft moving to right, climb positive aircraft moving up specify horizontal velocity, vertical rate of climb, and sideslip angle yaw positive to right, pitch positive nose up, roll positive to right Aircraft initial values specified if motion is trim variable; otherwise fixed for flight state tan( n , velocity specification: vel atmos ’general’ ’hover’ ’vert’ ’right’ ’left’ ’rear’ ’Vfwd’ ’Vmag’ ’climb’ ’+Mach’ ’+CAS’ orientation velocity relative inertial axes defined by climb and sideslip angles ( orientation body relative inertial axes defined by Euler angles, yaw/pitch/roll ( SET SET SET SET ’std’ ’polar’ ’trop’ ’hot’ ’xxx+dtemp’ _ _ SET velocities: forward aircraft motion: SET Structure: FltAircraft ) gear _ control d _ + gear _ IDENT WL ) to use state#n – ’ref’ ) hub _ ret ) , control ) or , based on _ WL xmsn ext _ , ) + ( Nspec def nstate control viscosity Vtip , temp , , HAGL and cruise oei (1 to _ _ csound FltState%control , (use input Vtip altitude control , _ ) ’ (keyword = ) control=n ) man density _ _ SET ) ; 1 to use FltState%tilt Aircraft%Vconv ’retract Vtip cruise viscosity altitude , _ , ) Vtip=’input’ temp with radius ratio , _ STATE , and tip Vtip Vtip cruise _ V SET ) _ or csound ’extend’ ; 1 to use density hover , Mat _ gear ) ( Vtip _ Aircraft%Vcont ref(1) at hover _ values (perhaps function speed) or flight state input _ Vtip ) M density _ and STATE Vtip Vtip Mat _ ( Aircraft Aircraft%Vtilt at mu ( = use default Aircraft%Vconv M specification and values put in and μ hover = Vtip _ for this flight state tilt to use conversion schedule, / Aircraft%cont Vtip(speed) ’xmsn’ Vtip , (for drive state Vtip pedal / ref to use _ ; out-of-ground-effect (OGE) if rotor more than 1.5Diameter above ground ’ (based on retraction speed), ’OEI’ Aircraft%tilt , control=0 = specified altitude, and specified temperature (use = hot day table at specified altitude (use = for variable diameter rotor, scale _ = for tip speed limited by Vtip lngcyc , aircraft control state: identifies control matrix / HAGL ’man’ = input, not air on earth (use ’default , control specification: = use default = use default = input density, speed of sound, and viscosity (use = use input , = input density and temperature (use = use conversion schedule ( : = use tip speed from control=0 = use tip speed from , primary rotor tip speed: for primary rotor of propulsion group = use _ STATE latcyc : use LG control / : 0 to use , center of gravity position: input for this flight state; or _ _ GE control tilt cg Vtip ’xxx+temp’ ’hot+table’ ’dens’ ’input’ ’notair’ height rotor = landing gear above ground + hub above landing gear = coll initial values specified if control is trim variable; otherwise fixed for flight state SET 2 to use conversion speeds baseline cg position plus shift due to nacelle tilt, plus input cg increment ’input’ ’ref’ ’speed’ ’conv’ ’hover’ ’cruise’ ’mu’ ’Mat’ ’xxx+Mat’ ’xxx+diam’ without rotors, specify engine group speed by _ _ _ _ _ SET STATE STATE SET SET SET tip speed, engine, transmission: for each propulsion group SET Structure: FltAircraft charge=’takeoff’ _ nCharge nEngine rating or ’MCP’ , of rotor jet=’takeoff’ nJet _ ’IRP’ , react F rating eng = charge=’idle’ ’MRP’ or , P _ , for each operable jet A , for each operable engine A reqJG ’ERP’ rating =off), first engine shaft limit and then drive system limit T ) to use gear ratio #n jet=’idle’ fPower , for each operable charger _ = 0 to 1; > 1 is an acceptable input) A Plimit nGear _ , proportional = 0 to 1; > 1 is an acceptable input) rating SET = 0 to 1; > 1 is an acceptable input) (1 to fCharge reqP G fPower P equals rotor power required fThrust state=n from (flight condition or mission), the gross weight is determined defines power distribution for sizing d reqEG ): only for reaction drive, + P reqEG P fPsize , drive system rating not used can be used to reduce the available power , 1 ): ’maxPQ’ avP G Jreq=1 ≤ _ or f P ds fPav Power _ _ rating=’takeoff’ = SET or Preq=1 _ or SET reqP G fPower P GW=’maxP’ to use conversion schedule, next torque limits are applied (unless with identical results, unless the engine group is operating at a torque limit _ , drive system state: identifies gear ratio set for multiple speed transmissions , number inoperative chargers: 1 for one charger inoperative (OEI), maximum : distribution of propulsion group power required among engine groups : use charge group amplitude control variable : fixed options use jet group amplitude control variable , number inoperative engines: 1 for one engine inoperative (OEI), maximum then the power available is reduced by the factor either except for reaction drive, SET , number inoperative jets: 1 for one jet inoperative (OEI), maximum gear reduces charger group power available ( reduces jet group thrust available ( rating=’idle’ reduces engine group power available ( _ Propulsion%nrate Preq Jreq Creq state=0 if or the engine model gives the power available, accounting for installation losses and mechanical limits for such that distributed ( fixed options use engine group amplitude control variable engine group that consumes shaft power (generator or compressor) only uses fixed option engine group that produces no shaft power (converted to turbo jet or reaction drive) only uses fixed option EngineGroup%SET from component ( _ _ _ STATE drive system rating: match rating designation in propulsion group; blank for same as rating of first engine group engine rating: match rating designation in engine model; e.g. fPower nEngInop SET jet rating: match rating designation in jet model; or fThrust nJetInop SET charger rating: match rating designation in charger model; or fCharge nChrgInop SET Structure: FltAircraft ’comp’ , ) ’ground’ Mat , , mu , quant ’power’ _ , xmsn , OEI , man ’windtunnel’ Aircraft%trim , , for no trim with radius ratio) ) cruise ’rotor’ tip , ’none’ , V ) , Mach _ hover trim ) , _ retract ’thrust’ vel2 , , _ conv , SET IDENT PerfCond , extend ’hover’ , , speed , 4 CAS , _ depends on designation of ref , defines trim states with following identification: default ’symm’ vel2 _ _ , OffMiss ) at ) target LG _ input _ M , 3 _ SET xxx xxx , or control number) _ _ , trim ) xxx Vtip _ _ TAS STATE _ trim=’free’ xxx atmos _ SizeMiss _ _ QUANT SET VAR _ vel vel2 _ , 2 , aircraft trim state: match _ _ SET IDENT trim MAX MAX SET SET _ identifies trim variables and quantities ACTION=’configuration’ requirement for SizeCond number number number STATE quantity ( quantity structure number quantity is slope (maximize) variable ( variable structure number velocity ( velocity ( sideward flight (1 for sideward flight) atmosphere ( landing gear state ( aircraft trim state (number, 0 for no trim) rotor tip speed ( rotor tip speed limited by rotor tip speed for variable diameter rotor (1 to scale power margin as quantity (2 maximum effort, 1 trim) torque margin as quantity (2 maximum effort, 1 trim) drive system rating number parent (1 Mission MissSeg FltState performance output column Maximum effort Specification Specification, each propulsion group int int int int int int int int int int int int int int int int int int int int int int Mat(npropmax) VarDiam(npropmax) _ _ LG trim _ _ ds(npropmax) vel vel2 atmos Vtip(npropmax) Vtip Vtip quant(2) quantn(2) isslope(2) var(2) varn(2) _ out _ _ _ _ _ _ _ _ _ _ _ _ Structure: FltAircraft parent kMission kMissSeg kFltState kcol imax imax imax imax imax iSET iSET isSideward iSET iSTATE iSTATE iSET iSET iSET iSETPmargin(npropmax) iSETQmargin(npropmax) krate ) ) ) ) other ) _ equip _ auxtank extkit _ _ fixUL ) _ fixUL ) _ ) fixUL W _ fixUL W _ fixUL W _ foldkit W _ ) W wingkit otherkit _ _ crew fixUL _ _ fixUL fixUL W _ _ W W fixUL _ W load − fuel ext O F U L ammo W pay pass W W W W W weapons G W W W cargo other categories W W crew (replace weight statement auxiliary fuel tanks (replace weight statement other fixed useful load (replace weight statement equipment increment (replace weight statement folding kit (replace weight statement wing extension kit (replace weight statement wing kit (replace weight statement other kit (replace weight statement usable fuel weight standard tanks auxiliary tanks passengers cargo external load ammunition weapons other fixed useful load increment (relative weight statement engine rating number jet rating number charger rating number jet thrust margin as quantity (2 maximum effort, 1 trim) charger power margin as quantity (1 trim) battery power margin as quantity (2 maximum effort, 1 trim) gross weight usable fuel weight payload weight fixed useful load operating weight number of crew number of passengers Specification Weight int int int int int int real real real real real real real real real real real real real real real real real real real real real real real real int int other _ total std(ntankmax) aux(ntankmax) pass cargo extload ammo weapons other jet(njetmax) charge(nchrgmax) _ _ _ _ _ _ _ _ _ _ _ fixUL _ fixUL _ Structure: FltAircraft krate(nengmax) krate krate iSETJmargin(njetmax) iSETCmargin(nchrgmax) iSETBmargin(ntankmax) GW Wfuel Wfuel(ntankmax) Wfuel Wfuel Wpayload Wpay Wpay Wpay Wpay Wpay Wpay WFixUL dW Wcrew Wauxtank W Woful(10) Wequip Wfoldkit Wextkit Wwingkit Wotherkit WO Ncrew Npass fuel fuel fuel E W E G W so for flight state, additional fixed useful load = auxiliary fuel tank and kits and increments usable fuel energy standard tanks auxiliary tanks standard tanks auxiliary tanks standard tanks auxiliary tanks weight statement defines fixed useful load and operating weight for design configuration gross weight = weight empty + useful load = operating weight + payload + usable fuel useful load = fixed useful load + payload + usable fuel operating weight = weight empty + fixed useful load xx yy zz xy yz xz number of crew seats number of passenger seats usable fuel energy gross weight usable fuel weight usable fuel energy stationline buttline waterline I I I I I I Weight at mission segment start Center of gravity position Moments of inertia int int real real real real real real real real real real real real real real real real real real real real real start(ntankmax) start(ntankmax) start(ntankmax) _ start(ntankmax) _ _ _ seat seat start(ntankmax) std aux _ total std(ntankmax) aux(ntankmax) start(ntankmax) std aux _ _ _ _ _ _ _ start _ _ _ _ Structure: FltAircraft Ncrew Npass Efuel Efuel(ntankmax) Efuel Efuel GW Wfuel Wfuel Wfuel Efuel Efuel Efuel zcg(3) SLcg BLcg WLcg Ixx Iyy Izz Ixy Iyz Ixz ) ) δ, M ( σf √ V s /c h in F axes V | (knots) ( (ft/sec or m/sec) | AC (ft/sec or m/sec) h cal v V (ft/sec or m/sec) | | s V V AC f / V q V T μ/ρ (knots) (deg) AC Δ h v (deg) 0 = (ft/sec or m/sec) V in F axes V R c θ − V ν Ω V p ψ AC T /T 0 d h 0 v spec N h Ω ref p/p ρ/ρ τ Ω R / N/N Ω horizontal velocity horizontal Mach number climb angle sideward velocity velocity magnitude forward velocity calibrated airspeed velocity drag vector, dynamic pressure h rotor tip speed rotor rpm rotor engine rpm engine propulsion group reference speed ratio horizontal speed climb velocity sideslip angle derived altitude temperature temperature increment density ratio temperature ratio pressure ratio kinematic viscosity density altitude pressure altitude rotor radius rotational speeds flight speed Atmosphere Flight condition real real real real real real real real real real real real real real real real real real real real real real real real real real real real real rotor(nrotormax) eng(nengmax) ref(npropmax) _ _ _ trim trim(nrotormax) _ trim(nrotormax) trim _ _ _ trim trim trim trim(nengmax) _ _ _ _ Structure: FltAircraft alt tmp dtmp sigma theta delta kinvis altdens altpress radius(nrotormax) Vtip rpm rN N rN rN speed Vclimb side Vhoriz Mhoriz climb Vside Vmag Vfwd VCAS VAC(3) ed(3) qAC ) F φ , F θ , F ψ ) V ψ , V θ MissSeg or isSideward FltCond and Aircraft )*time turn (each aux tank size), from MissSeg auxtank N , yaw angle = ( , including auxiliary tanks , including auxiliary tanks yaw yaw Z cap Z cap − − FltCond side (linear acc and angular rate) fuel pitch Z fuel (+uphill; deg), from Y (deg) W E (deg) G T AC F γ roll θ F climb R n (pitch rate) φ X Y (yaw rate) F = ˙ = θ F ˙ in F axes in F axes (linear) , body axes relative velocity axes ψ , velocity axes relative inertial axes , body axes relative inertial axes F I V I yaw positive to right, pitch positive nose up, roll positive to right C sideslip positive aircraft moving to right, climb positive aircraft moving up C F I V I F V AC AC turn pullup turn radius ω a load factor pitch angle roll angle C C C angle of attack and sideslip angle representation: from orientation body relative inertial axes defined by Euler angles, with yaw/pitch/roll sequence ( orientation velocity relative inertial axes defined by climb and sideslip angles ( angular velocity acceleration angle of attack and sideslip angle representation (1 conventional, 2 reversed) orientation of body axes relative inertial axes rotation matrices aircraft controls number of auxiliary fuel tanks wing extension kit on aircraft (0 none, 1 present) wing kit on aircraft (0 none, 1 present) total fuel capacity total fuel capacity slope of ground parameter sweep, from real real real real real real int real real real real real real int int int real real real int trim(ncontmax) trim _ cap(ntankmax) alpha cap(ntankmax) ground _ trim _ _ extkit(nwingmax) wingkit(nwingmax) sweep trim _ _ _ _ _ _ _ trim _ Structure: FltAircraft turn pullup turnRadius wAC(3) aAC(3) nAC(3) KIND pitch roll CFI(3,3) CVI(3,3) CFV(3,3) control Nauxtank(nauxtankmax,ntankmax) SET SET Wfuel Efuel slope SET climb − Y side − Z pitch Y roll X = F V C (jet thrust, momentum drag) (fuselage, rotor, wing, tail, tank, engine, jet, charger) F F eng charge (turn) F aero F F F grav F F F jet F F inertial F F rotor F aerodynamic rotor engine groups jet groups charge groups gravitational inertial orientation body relative velocity axes: converged (0 not) number of iterations error ratio gain matrix converged (0 not) number of iterations error ratio quantity switched (1 P margin, 2 Q margin, 3 both) converged (0 not) number of iterations error ratio converged (0 not, –1 no iteration) count of solution (0 at start, get aircraft controls) trim derivative matrix exist (0 for not) forces (F axes, about cg) Trim (last) Maximum effort (principal iteration, 99% range iteration; inner, outer loops) Maximum gross weight (flight condition or mission takeoff) Rotor flap equation (all converged or any not converged) Solution state Loads int int real real int int real int int int real int int int real real real real real real real exist _ trim fly(2,2) maxgw control _ trim(mtrimmax) _ fly(2,2) _ maxgw _ trim(mtrimmax,mtrimmax) deriv _ _ _ _ _ Structure: FltAircraft istrimconv count error gain isflyconv(2,2) count error isSwitched(2) ismaxgwconv count error isrotorconv count trim Faero(3) Frotor(3) Fengine(3) Fjet(3) Fchrg(3) Fgrav(3) Finertia(3) avCG limit P ) DS avP G P P ) ) (1 + avJG T > (1 + ) (1 + > > reqCG (1 + P > reqP G reqP G P P jet V T reqJG T inertia + ) F req − inertia ) req P P M ) ) (propulsion groups and converted engine groups) = − grav − Fvert ) req req F P av T P M req + P P − − F − av (jet thrust, momentum drag) min( T , from energy flow limit av (charge groups) (engine groups) P (fuselage, rotor, wing, tail, tank, engine, jet, charger) P (jet groups) (I axes; sum equiv min( F F z req chrg engine charge (turn) ˙ jet F w aero F P P W (I axes); set to 0 if V>10 knots min( T min( M M − DL/W M z F jet DL/T F ˙ climb M F w inertial V ˙ F E rotor M ˙ M w aerodynamic rotor engine groups jet groups charge groups inertial power required power margin, torque margin, exceed power available: any propulsion group exceed torque available: any propulsion group thrust required jet thrust margin, exceed jet thrust available: any jet group power required charger power margin, exceed charger power available: any charge group moments (F axes, about cg) total force (F axes, about cg); total moment (F axes, about cg); download, aero rotor thrust, rotor download/thrust download/weight power thrust charging equivalent aircraft power required climb power, fuel flow total fuel flow equivalent fuel flow energy flow Aircraft performance real real real real real real real real real real real real real real real int int real real int real real int real real real real real real total equiv _ _ Structure: FltAircraft Maero(3) Mrotor(3) Mengine(3) Mjet(3) Mchrg(3) Minertia(3) Ftotal(3) Mtotal(3) Download Thrust DLoT DLoW Preq Pmargin Qmargin exceedP exceedQ Tjet Jmargin exceedJ Pchrg Cmargin exceedC Pequiv Pclimb fuelflow(ntankmax) fuelflow fuelflow energyflow(ntankmax) cap − fuel E ) (1 + max P > ) fuel E (1 + (MJ/hr) or /P > ) ) | | cap ref − batt batt W V /P ˙ ˙ ρA E E | = fuel (2 e W − | ) ; set to 0 if V<10 knots W/ (ton-nm/lb or ton-nm/kg) L/D √ max AC P (1 + W /q (ton-kt/lb or ton-kt/kg) equiv P/V ˙ > O = w AC min( = D V / fuel e F M = V /W D W pay equiv pay W ˙ ˙ w D/q E W = V / AC W/P e = equiv D p ˙ w / equiv battery power margin, exceed battery power: any fuel tank P total energy flow exceed fuel capacity: battery sfc, specific range, aircraft figure of merit aircraft effective lift-to-drag ratio aircraft effective drag aircraft drag aircraft drag area power loading range for fuel=1%GW (nm) fuel efficiency productivity length width area Performance indices Operating size real int real int real real real real real real real real real real real real real real total _ op op onepcW _ range _ _ op eff _ _ _ Structure: FltAircraft energyflow exceedWf Bmargin exceedB sfc spec FM LoDe Drage DragAC DoQAC WoP range fuel productivity length width area Default (F axes, about cg) (F axes) (B axes) M F B N C (F axes, about cg) F , from rotors (F axes) v v C aero Y F int M C v F aero F (I axes) D (deg) z (deg) L C F α C β D/q F aero D/q F T d e BA interference velocity total velocity relative air total velocity relative air angle of attack sideslip angle C velocity magnitude dynamic pressure payload contingency lift coefficient pitch moment coefficient drag coefficient side force coefficient yaw moment coefficient lift pitch moment drag side force yaw moment aerodynamic force aerodynamic moment drag download, aero Flight State - Fuselage aerodynamics loads Description Type real real real real real real real real real real real real real real real real real real real real real real real real pay cont _ _ Chapter 21 Structure: FltFuse Variable VintR(3,nrotormax) Vaero(3) VB(3) alpha beta CBA(3,3) Vmag q DoQ DoQ CL CM CD CY CN L M D Y N Faero(3) Maero(3) Drag Download Default ) retracted , extended _ (F axes, about cg) (F axes) LG F _ (F axes, about cg) F v aero M F aero in F axes F (I axes) STATE | v z | F v/ − F aero F T d e landing gear state ( total velocity relative air velocity magnitude dynamic pressure drag vector, aerodynamic force aerodynamic moment drag download, aero Flight State - Landing Gear aerodynamics Description Type int real real real real real real real real LG _ Chapter 22 Structure: FltGear Variable iSTATE Vaero(3) Vmag q ed(3) Faero(3) Maero(3) Drag Download Default ) a/ blade A tip ρV , a/ (1, S (–1 TPP command, 1 NFP command) (advancing tip or helical) S at BF M SF C C nac tip z R M pylon Ω cont z hub = Ω C z matrix scale factor tip T (thrust weighted) V matrix scale factor σ R T collective control mode (1 thrust command, 2 pitch command) cyclic control mode (1 TPP command, 2 NFP command) collective cyclic collective longitudinal cyclic lateral cyclic incidence cant diameter gear ratio factor shaft control, shaft relative airframe, hub position, pylon position, nacelle cg position, pylon relative airframe, radius tip speed rotational speed tip Mach number maximum Mach number solidity Flight State - Rotor control mode controls geometry condition Description Type int int real real real real real real real real real real real real real real real real real real real real real coll cyc _ _ control control _ _ Chapter 23 Structure: FltRotor Variable KIND KIND Scoll Scyc coll lngcyc latcyc incid cant diam fgear Ccont(3,3) CSF(3,3) zhub(3) zpylon(3) znac(3) CBF(3,3) radius Vtip Omega Mtip Mat sigma (F axes) (S axes) M y P S F S s M x C v v θ C C H Y c C blade C β (0.75R) I c Q T θ C .

C 0 s θ β i ν κλ hub ν = γ K β λ x y tip μ μ thrust drag force side force roll moment pitch moment torque thrust coefficient drag force coefficient side force coefficient roll moment coefficient pitch moment coefficient torque coefficient collective pitch longitudinal cyclic pitch lateral cyclic pitch coning longitudinal flapping lateral flapping inflow V /V total velocity relative air total velocity relative air Lock number blade moment of inertia flap frequency coning frequency hub stiffness shaft axis loads control and motion tip-path plane relative shaft, velocity and inflow performance real real real real real real real real real real real real real real real real real real real real real real real real real real real real real real Structure: FltRotor gamma Iblade flapfreq conefreq Khub T H Y Mx My Q CT CH CY CMx CMy CQ theta75 thetas thetac beta0 betac betas lambda0 CPS(3,3) VoVtip VF(3) VS(3) mux muy i λ /D T g /U C z ) m f = | 8) g W cos f /σ f T D i ideal T ( σa/ f λ P C A / | / from wings, normal (F axes) from wings, inplane (F axes) f g C (S axes) T = ( C S C F int F int ) g P ) x y m s v v ω z F √ H κ κ κ E 2 y /μ F , tpp , profile , , tpp , profile z μ = c H H μ Y Y h ( E + C C C C , 1 λ t 2 x − E μ z x y ˙ ˙ √ μ α α for inflow solution = = tan T angular velocity μ α ducted fan area ratio ducted fan thrust ratio ducted fan far wake ratio height rotor hub above ground, effective height, ground effect inflow ratio C reference ideal induced velocity ideal induced power inflow gradient inflow gradient inflow gradient drag force side force drag force side force thrust coefficient/solidity, profile power factor profile drag factor interference velocity interference velocity inplane forces converged (0 not, –1 no iteration) iteration count error ratio ( rotor derivative matrix exist (0 for not) rotor flap equations real real real real real real real real real real real real real real real real real real real real real real real real real real real real int int real int exist _ ideal _ rotor deriv _ rotor(3) _ _ Structure: FltRotor muz omegaS(3) dax day mu alphas fDuctA fDuctT fDuctW zg zge fg CTe lambdah lambda CPideal kappax kappay kappam CTs FPpro FHpro VintWn(3,nwingmax) VintWp(3,nwingmax) CHtpp CYtpp CHo CYo isrotorconv count error rotor ) ds _ rating limit and RS P ) ) (sustained) (transient) trim /P tip _ ) F v /σ /σ P t D T T rpm /C f V /V (1 + C ( C C P t / P i P p + (at ) − − P P o C v/P C C V P > P C D P o − ( mean = C d C limit T max max t c (F axes, about cg) κ T f T V /P ) ) λ t + RS limit P /σ /σ P (F axes, about cg) X P i T T λ F RS i p rotor L C C C C o P P P ( C ( ( M P F rotor P (sustained) (transient) F /σ /σ T T C C rotor force rotor moment lift (wind axis) drag (wind axis) lift coefficient drag coefficient vertical force (inertia axes) max max thrust margin, thrust margin, drive system limit torque margin, exceed torque available: rotor power induced power interference power profile power parasite power rotor power coefficient induced power coefficient interference power coefficient profile power coefficient parasite power coefficient induced velocity interference velocity induced power factor mean drag coefficient hover figure of merit, propulsive efficiency, momentum efficiency, effective drag, loads power real real real real real real real real real real real real real int real real real real real real real real real real real real real real real real real real steady tran rs rs _ _ _ _ rs _ steady tran _ _ Structure: FltRotor Frotor(3) Mrotor(3) L X CL CX Fvert CTs CTs Tmargin Tmargin Plimit Qmargin exceedQ P Pind Pint Ppro Ppar CP CPind CPint CPpro CPpar lambda lambdat Ki cdmean FM etaprop etamom CDe react r Ω / react P (F axes) (B axes) (F axes) (B axes) = F B F B v v v v (F axes, about cg) De pylon react (F axes, about cg) duct F D /C hub aero in F axes in F axes F D L | | C D react M v (deg) v (deg) C C P | | C F aero α α F (I axes) v/ v/ react z − − r F spin Ω pylon shaft D duct P hub D D D F aero F T d e total velocity relative air velocity magnitude dynamic pressure drag vector, total velocity relative air angle of attack total velocity relative air velocity magnitude dynamic pressure drag vector, total velocity relative air angle of attack effective lift-to-drag, shaft power reaction drive power reaction drive net force blade velocity hub pylon drag coefficient, hub drag coefficient, pylon drag coefficient, duct drag, hub drag, pylon drag, duct drag, spinner aerodynamic force aerodynamic moment drag download, aero shaft power and reaction drive aerodynamics loads real real real real real real real real real real real real real real real real real real real real real real real real real real real real hub pylon hub(3) pylon(3) hub pylon _ _ _ _ _ _ hub(3) pylon(3) hub(3) pylon(3) _ _ hub _ pylon _ _ _ Structure: FltRotor LoDe Pshaft Preact Freact rOmegareact Vaero Vmag q ed VB alpha Vaero Vmag q ed VB alpha CDhub CDpylon CDduct Dhub Dpylon Dduct Dspin Faero(3) Maero(3) Drag Download int α ) T C at fuselage at wing, left panel at wing, right panel at tail t t t t f f f f r r r r (from f f f f i z z z z λ f f f f (F axes) W W W W f f f f F ind t v f χ ideal induced velocity induced velocity wake angle interference factor interference factor interference factor interference factor fuselage inside wake wing inside wake, left panel wing inside wake, right panel tail inside wake twin rotor factor induced power interference at wing interference real real real real real real real int int int int real real fus wingLp(nwingmax,npanelmax) wingRp(nwingmax,npanelmax) tail(ntailmax) _ _ _ _ int _ wing(nwingmax) fus wingLp(nwingmax,npanelmax) wingRp(nwingmax,npanelmax) tail(ntailmax) _ _ _ _ _ wake _ Structure: FltRotor lambda vind(3) chi Fint Fint Fint Fint isInWake isInWake isInWake isInWake ftwin Aint Default (F axes) (B axes) , from rotors F B int v v α DV at left wing panel, from rotors (F axes) at right wing panel, from rotors (F axes) (panel area weighted), from rotors (F axes) , from other wings (F axes) C , from other wings F int F int F int F int (deg) v v v v (deg) int α α β , with interference (deg) α i f δ a δ BA F δ total velocity relative air total velocity relative air angle of attack sideslip angle C velocity magnitude dynamic pressure flap flaperon aileron incidence interference velocity interference velocity interference velocity interference velocity interference angle induced power interference with mean interference angle of attack vertical drag coefficient Flight State - Wing controls aerodynamics Description Type real real real real real real real real real real real real real real real real real real real Lp(3,nrotormax,npanelmax) Rp(3,nrotormax,npanelmax) int _ _ _ Chapter 24 Structure: FltWing Variable flap(npanelmax) flaperon(npanelmax) aileron(npanelmax) incid(npanelmax) VintR VintR VintR(3,nrotormax) VintW(3,nwingmax) AintW(nwingmax) AintR(nrotormax) Vaero(3) VB(3) alpha beta CBA(3,3) Vmag q alpha CDV (F axes) (B axes) (F axes) (B axes) M p M p Dpp Dpp F B F B p p C C C C v v v v C C (deg) (deg) (deg) Lp (deg) Lp α α C C β β (weighted by panel area) BA BA /S total velocity relative air total velocity relative air angle of attack sideslip angle C velocity magnitude dynamic pressure lift coefficient drag coefficient, parasite pitch moment coefficient roll moment coefficient lift drag, parasite pitch moment roll moment total velocity relative air total velocity relative air angle of attack sideslip angle C velocity magnitude dynamic pressure lift coefficient drag coefficient, parasite pitch moment coefficient roll moment coefficient lift drag, parasite pitch moment roll moment ) (sum over panels) qS left panel right panel qS ( real real real real real real real real real real real real real real real real real real real real real real real real real real real real real real real real Lp(npanelmax) Rp(npanelmax) Lp(3,npanelmax) Rp(3,npanelmax) Lp(npanelmax) Rp(npanelmax) _ _ _ Lp(3,3,npanelmax) _ Rp(3,3,npanelmax) _ Lp(npanelmax) Lp(npanelmax) _ Rp(npanelmax) Rp(npanelmax) _ Lp(npanelmax) _ Rp(npanelmax) _ _ _ _ Lp(3,npanelmax) Lp(npanelmax) Lp(npanelmax) Rp(3,npanelmax) Rp(npanelmax) Rp(npanelmax) Lp(npanelmax) Rp(npanelmax) _ _ _ _ Lp(npanelmax) _ _ Rp(npanelmax) _ _ _ _ Lp(npanelmax) Rp(npanelmax) Lp(npanelmax) Lp(npanelmax) Rp(npanelmax) Rp(npanelmax) _ _ _ _ _ _ _ _ Structure: FltWing Vaero VB alpha beta CBA Vmag q CL CDp CM CR L Dp M R Vaero VB alpha beta CBA Vmag q CL CDp CM CR L Dp M R qS qeff ) B /v ind v int K = Lα B C /v int (F axes, about cg) max v M Di L Dp (F axes) C C C (F axes, about cg) = C F L aero C F int C M v int (F axes) F aero − α , normal (F axes) , inplane (F axes) F (I axes) (F axes) F ind F int F int z v v v L max F int F L v C C F aero F T d e max compressible 3D lift curve slope α lift coefficient drag coefficient, parasite drag coefficient, induced pitch moment coefficient roll moment coefficient maximum lift coefficient lift drag, parasite drag, induced drag pitch moment roll moment stall margin, aerodynamic force aerodynamic moment drag download, aero velocity at tail induced velocity velocity at other wing angle at other wing ( velocity at rotor velocity at rotor loads interference real real real real real real real real real real real real real real real real real real real real real real real real real rotor(3,nrotormax) rotor(3,nrotormax) max tail(3,ntailmax) wing(3,nwingmax) wing(nwingmax) _ _ _ _ _ _ Structure: FltWing dCLda3D AoA CL CDp CDi CM CR CLmax L Dp Di D M R Lmargin Faero(3) Maero(3) Drag Download Vint vind(3) Vint Aint Vintn Vintp Default Lα C (F axes, about cg) (F axes) (B axes) max Di L Dp F B C C (F axes, about cg) C F , from rotors (F axes) , from wings (W axes) v v aero F int F int M v v F aero F (I axes) (deg) z (deg) L F α C β i F aero F δ T d e BA max control incidence interference velocity interference velocity total velocity relative air total velocity relative air angle of attack sideslip angle C velocity magnitude dynamic pressure compressible 3D lift curve slope α lift coefficient drag coefficient, parasite drag coefficient, induced maximum lift coefficient lift drag aerodynamic force aerodynamic moment drag download, aero Flight State - Tail controls aerodynamics loads Description Type real real real real real real real real real real real real real real real real real real real real real real real max _ Chapter 25 Structure: FltTail Variable cont incid VintR(3,nrotormax) VintW(3,nwingmax) Vaero(3) VB(3) alpha beta CBA(3,3) Vmag q dCLda3D AoA CL CDp CDi CLmax L D Faero(3) Maero(3) Drag Download Default max P (MJ/hr) ) (MJ/hr) | cap − (1 + batt ˙ (MJ/hr) E fuel > charge; power and current positive) E | − | ) (MJ/hr) (MJ/hr) (MJ/hr) comp crit batt mbd ˙ d (1/hr) ˙ E x max E eff max ˙ | ˙ E < P batt x E = ˙ E max batt x η cap (MJ/hr) mbd P loss eq x/x discharge, P P ˙ (1/hr) 0 = w V x ξ ˙ E > usable fuel weight usable fuel energy fuel weight capacity fuel energy capacity fraction weight capacity fraction energy capacity = state-of-charge = 1 - depth-of-discharge power capacity state (1 discharging, -1 CC charge, -2 CV charge) current current voltage component energy flow battery efficiency power loss battery energy flow effective capacity factor effective energy flow maximum current maximum power (for battery power margin exceed battery power: power loss fuel flow Flight State - Fuel Tank Systems all tanks (standard plus auxiliary) battery ( equipment power Description Type real real real real real real real int real real real real real real real real real real real real int real real cap cap cap _ _ _ Chapter 26 Structure: FltTank Variable Wfuel Efuel Wfuel Efuel rWfuel rEfuel Pfuel state x xi V Edotcomp etabatt Ploss Edotbatt dcrit Edoteff xmax Pmax Bmargin exceedB Peq fuelflow (F axes, about cg) (F axes) , from energy flow F (F axes, about cg) F v aero M equiv ˙ F aero w in F axes (I axes) F (I axes) | v z z | F F v/ ˙ E − F aero F T d D e energy flow equivalent fuel flow total velocity relative air velocity magnitude dynamic pressure drag vector, drag download, aero aerodynamic force aerodynamic moment drag download, aero auxiliary tanks aerodynamics loads real real real real real real real real real real real real equiv _ Structure: FltTank energyflow fuelflow Vaero(3,nauxtankmax) Vmag(nauxtankmax) q(nauxtankmax) ed(3,nauxtankmax) D(nauxtankmax) DL(nauxtankmax) Faero(3) Maero(3) Drag Download Default ) ds _ rating , propulsion group and ) limit acc P limit DS avP G P + P ) ) DS P (sum all engine groups producing shaft power) xmsn (sum all engine groups producing shaft power) P trim(primary) (1 + (1 + _ avEI + > > P avEG rpm P limited by reqP G comp (at P P , from energy flow reqP G reqP G reqP G P P P − = , propulsion group (sum all engine groups producing shaft power) avP G limit − , all components equiv P xmsn ˙ w DS P limit acc P avP G comp reqP G P DS avP G P P P , propulsion group P ˙ P w ˙ E avP G ˙ w /P rotor engine groups reqP G rotational speed increment, primary rotor or primary engine (rpm) power required transmission losses accessory power power required power available engine installed power available engine group power available P drive system limit at drive system limit ( torque margin, power margin, exceed power available: exceed torque available: torque margin, min(propulsion group, engine groups, rotors) exceed torque available: any propulsion group, engine groups, rotors fuel flow total fuel flow equivalent fuel flow energy flow Flight State - Propulsion Group drive system state control power propulsion group engines Description Type int real real real real real real real real real real real real int real real int int real int real real real real ds ds ds total equiv _ gear _ rotor eng _ _ _ _ _ _ ds _ trim _ Chapter 27 Structure: FltProp Variable STATE DN Pcomp Pcomp Pcomp Pxmsn Pacc PreqPG PavPG PavEIsum PavEGsum Pratio Plimit atPlimit Qmargin Pmargin exceedP exceedQ Qmargin exceedQ fuelflow(ntankmax) fuelflow fuelflow energyflow(ntankmax) (F axes, about cg) (F axes, about cg) F req prop M /P F prop F equiv w (F axes, about cg) sfc = ˙ (F axes, about cg) F aero M F aero F (I axes) ˙ z E F F aero F T d e total energy flow specific fuel consumption jet thrust and momentum drag force jet thrust and momentum drag moment aerodynamic force aerodynamic moment drag download, aero real real real real real real real real total _ Structure: FltProp energyflow sfc Fprop(3) Mprop(3) Faero(3) Maero(3) Drag Download Default aux D eng eng a − − q P P av req BF G P P C F N loss loss η F η mom f F gear loss loss e f P P ˙ E N ˙ A m ˙ i w B ψ amplitude mode incidence yaw gear ratio factor engine relative airframe, engine direction, uninstalled power required, installation loss installation efficiency installed power required, engine rpm mass flow fuel flow energy flow gross installed jet thrust momentum thrust net installed jet thrust momentum drag of auxiliary air flow uninstalled power available, installation loss installation efficiency installed power available, Flight State - Engine Group controls geometry engine Description Type real real real real real real real real real real real real real real real real real real real real real real real eng _ eng _ trim _ Chapter 28 Structure: FltEngn Variable amp mode incid yaw fgear CBF(3,3) ef(3) Pq Plossq etalossq Preq N mdot wdot Edot FG Fmom FN Daux Pa Plossa etalossa Pav f f d K q P ) eng − av NEngInOp – P ) eng − av Nengine P ) ) NEngInOp ds – ) _ limit ) ES limit P rating ) mech ES req NEngInOp ) Nengine P – P aux ; ( /P compreact or converted) and _ trim D _ (1 + N avEI equiv trim > _ P Nengine w (F axes, about cg) ( ) (at engine N _ limited by limited by reqEG F (F axes) motor (F axes, about cg) aux (at trim η P , from energy flow reqEG reqEG _ sfc = ˙ F mech eng M (generator or compressor); ( fPower P P N F v P − − aux ; N (F axes, about cg) avEG limit F − av equiv F P (at ˙ F P jet cell w comp ES in F axes limit (F axes, about cg) η | P avEG P M req v reqEG F P ES jet | avEG P Q P P (negative if generated) F v/ ˙ E − (negative if generated) avEG ˙ w /P reqEG engine mechanical limit at mechanical limit ( motor/generator efficiency fuel cell efficiency reaction drive mode (MODEL converted (KIND=RPTEM with mode=1; 0 shaft power, 1 reaction, 2 jet) shaft power (0 consumed (generator or compressor), 1 produced) component power power required engine installed power available power available, torque required P power margin, drive system limit at drive system limit ( torque margin, exceed torque available: fuel flow energy flow equivalent fuel flow specific fuel consumption net installed jet thrust momentum drag of auxiliary air flow jet thrust force jet thrust moment momentum drag force momentum drag moment total velocity relative air velocity magnitude dynamic pressure drag vector, engine group aerodynamics real int real real int int int real real real real real real real real int real int real real real real real real real real real real real real real real es es es equiv _ _ _ _ es _ Structure: FltEngn Pmech atPmech etamotor etacell ReactionMode Converted ProducePower Pcomp Preq PavEI Pav Qreq Pratio Pmargin Plimit atPlimit Qmargin exceedQ fuelflow energyflow fuelflow sfc FNEG DauxEG Fjet(3) Mjet(3) Faux(3) Maux(3) Vaero(3) Vmag q ed(3) (F axes, about cg) (B axes) B (F axes, about cg) F v aero M F aero F (I axes) D (deg) z C F α F aero F T d e total velocity relative air angle of attack drag coefficient drag aerodynamic force aerodynamic moment drag download, aero load real real real real real real real real Structure: FltEngn VB(3) alpha CD D Faero(3) Maero(3) Drag Download Default ) mech aux T D jet jet − a − q limited by T T av req G jet T T F − N loss loss BF η F η av mech C T T mom F loss loss T T f e ˙ E ˙ A m ˙ i w B ψ amplitude mode incidence yaw jet relative airframe, jet direction, uninstalled thrust required, installation loss installation efficiency installed thrust required, mass flow fuel flow energy flow gross installed jet thrust momentum thrust net installed jet thrust momentum drag of auxiliary air flow uninstalled thrust available, installation loss installation efficiency installed thrust available, jet mechanical limit at mechanical limit ( Flight State - Jet Group controls geometry jet Description Type real real real real real real real real real real real real real real real real real real real real real real int jet _ jet _ Chapter 29 Structure: FltJet Variable amp mode incid yaw CBF(3,3) ef(3) Tq Tlossq etalossq Treq mdot wdot Edot FG Fmom FN Daux Ta Tlossa etalossa Tav Tmech atTmech jet − av jet T − ) av T ) avJG T NJetInOp req ) – aux /T NJetInOp D – Njet react or converted) (1 + ; ( equiv _ Njet > w (F axes, about cg) ( jet _ avJI (F axes, about cg) F (F axes) (B axes) (F axes, about cg) aux T , from energy flow sfc = ˙ F B reqJG M reqJG (F axes, about cg) F F v v aero fThrust T aux T N (F axes, about cg) ; F M equiv F − ˙ F F aero jet w in F axes (F axes, about cg) F | M avJG D v (deg) F jet | reqJG T avJG C T α (negative if generated) F T (F axes) v/ ˙ E − (negative if generated) F jet ˙ (F axes) avJG w M F /T jet F reqJG reaction drive mode (MODEL converted (RPJEM with mode=1; 0 jet, 1 reaction) thrust required jet installed thrust available thrust available, T thrust margin, exceed thrust available: fuel flow energy flow equivalent fuel flow specific fuel consumption net installed jet thrust momentum drag of auxiliary air flow jet thrust force jet thrust moment momentum drag force momentum drag moment force moment total velocity relative air velocity magnitude dynamic pressure drag vector, total velocity relative air angle of attack drag coefficient drag aerodynamic force aerodynamic moment jet group loads aerodynamics load int int real real real real real int real real real real real real real real real real real real real real real real real real real real real real equiv _ Structure: FltJet ReactionMode Converted Treq TavJI Tav Jratio Jmargin exceedJ fuelflow energyflow fuelflow sfc FNJG DauxJG Fjet(3) Mjet(3) Faux(3) Maux(3) F(3) M(3) Vaero(3) Vmag q ed(3) VB(3) alpha CD D Faero(3) Maero(3) (I axes) z F F aero F T d e drag download, aero real real Structure: FltJet Drag Download Default ) cell q NChrgInOp ˙ E – ) Ncharge ( / NChrgInOp – f f d chrg K /η aux D req chrg Ncharge ˙ P reqCG E /η reqCG ˙ ; ( E ˙ = E = req = P ˙ = total cell E reqCG cell a BF ˙ q = ˙ E = req req ˙ E G C E ˙ P E F = N cell cell reqCG a F q ˙ P ˙ E E mom chrg f η F e reqCG P ˙ A m ˙ i w B ψ amplitude mode incidence yaw charger relative airframe, charger direction, cell power available cell energy flow available cell energy flow required cell power required charger efficiency installed power required mass flow fuel flow gross installed jet thrust momentum thrust net installed jet thrust momentum drag of auxiliary air flow power required energy flow required total cell power required Flight State - Charge Group controls geometry charger charger, fuel cell charge group Description Type real real real real real real real real real real real real real real real real real real real real real burn burn _ _ Chapter 30 Structure: FltChrg Variable amp mode incid yaw CBF(3,3) ef(3) Pacell Edotacell Edotqcell Pqcell etachrg Preq mdot wdot FG Fmom FN Daux Pchrg Edotchrg Preqtotal cell a ˙ E ) NChrgInOp – total req av P Ncharge ) /P ( aux equiv D (1 + w fCharge > ; (F axes, about cg) , from energy flow total sfc = ˙ total total (F axes, about cg) F (F axes) (B axes) (F axes, about cg) aux req av , from energy flow req equiv F B P M P ˙ (F axes, about cg) F P F v v w aero aux N − (F axes, about cg) F M equiv F ˙ F F aero jet w in F axes (F axes, about cg) F total | M D v (deg) ˙ F jet | av E C (F axes) α P (negative if generated) F ˙ v/ total w ˙ E − (F axes) av F chrg /P M F chrg F total fuel flow energy flow equivalent fuel flow specific fuel consumption req total cell power available P power margin, exceed power available: energy flow equivalent fuel flow fuel burn net installed jet thrust momentum drag of auxiliary air flow jet thrust force jet thrust moment momentum drag force momentum drag moment force moment total velocity relative air velocity magnitude dynamic pressure drag vector, total velocity relative air angle of attack drag coefficient drag aerodynamic force aerodynamic moment charge group, fuel cell loads aerodynamics load real real real int real real real real real real real real real real real real real real real real real real real real real real real real burn _ burn _ equiv burn equiv _ _ _ Structure: FltChrg Pavtotal Cratio Cmargin exceedC energyflow fuelflow fuelflow energyflow fuelflow sfc FNCG DauxCG Fjet(3) Mjet(3) Faux(3) Maux(3) F(3) M(3) Vaero(3) Vmag q ed(3) VB(3) alpha CD D Faero(3) Maero(3) (I axes) z F F aero F T d e drag download, aero real real Structure: FltChrg Drag Download 1 0 1 0 0 0 1 3 .5 4. .5 40 40 .01 .001 .002 Default maximum number of iterations tolerance (deg) relaxation factor maximum number of iterations tolerance (fraction reference) relaxation factor perturbation (1 first order, 2 second order) number of iterations between identification (0 for never recalculated) variable perturbation amplitude (fraction reference) title notes convergence control derivative (1 first order, 2 second order) maximum increment amplitude (0. for no limit) trace iteration (0 for none) convergence control perturbation identification of derivative matrix reinitialize aircraft controls (0 no, 1 force retrim) trace iteration (0 for none, 2 for component controls) method (1 secant, 2 false position) maximization method (1 secant, 2 false position, 3 golden section search, 4 curve fit) Solution Procedures Rotor Trim Maximum effort Description + + + + + + + + + + + + + + + + + + + + + + + + + Type c*100 c*1000 int real real int real int int real real int int real int int int int fly flymax trim rotor(nrotormax) _ _ _ _ rotor(nrotormax) rotor(nrotormax) trim trim trim rotor(nrotormax) rotor(nrotormax) rotor(nrotormax) trim trim trim _ _ _ _ _ _ _ _ _ _ _ trim _ Chapter 31 Structure: Solution Variable title notes niter toler relax deriv maxinc trace niter toler relax deriv mpid perturb init trace method method 3 0 0 1 0 0 .5 0. 0. .5 0. 0. 1. 1. 1.

80 40 40 40 40 .05 .98 .02 .01 .002 .002 maximum number of iterations tolerance (fraction reference) relaxation factor maximum number of iterations tolerance (fraction reference) relaxation factor maximum number of iterations tolerance (fraction reference) relaxation factor (mission fuel) relaxation factor (range credit) relaxation factor (max takeoff GW) maximum number of iterations (performance loop) maximum number of iterations (parameter loop) convergence control variable perturbation amplitude (fraction reference) maximum derivative amplitude (0. for no limit) maximum increment fraction (0. for no limit) extent of curve fit (fraction maximum) order of curve fit (2 quadradic, 3 cubic) reinitialize aircraft controls (0 no, 1 force retrim) trace iteration (0 for none) method (1 secant, 2 false position) convergence control variable perturbation amplitude (fraction reference) maximum derivative amplitude (0. for no limit) maximum increment fraction (0. for no limit) trace iteration (0 for none) convergence control trace iteration (0 for none) convergence control Maximum gross weight (flight condition or mission takeoff) Mission Size aircraft + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + int real real real real real real int int int int int real real real real real int int real real real real int int int fly maxgw _ _ maxgw fly maxgw fly maxgw _ _ _ _ _ fly maxgw miss fly fly fly maxgw maxgw maxgw miss miss miss range gw size param _ _ _ _ _ _ _ _ _ _ _ _ _ _ fly fly _ _ fly _ _ _ Structure: Solution niter toler relax perturb maxderiv maxinc rfit nfit init trace method niter toler relax perturb maxderiv maxinc trace niter toler relax relax relax trace niter niter 0 1 0 1. 1. 1. 1. 1. 1. 0. 0. 0. 0. 0. 0.

.01 .005 1) > param start _ _ niter trace > counter ; trace of iterations suppressed until or case=1 _ (fraction reference) trace , parameter iteration is part of performance loop (can be faster than Wfuel to identify point at which analysis diverges , param=1 Qlimit case=2 _ , _ power or radius gross weight drive system limit WMTO and SDGW fuel tank capacity rotor thrust power or radius gross weight drive system limit WMTO and SDGW fuel tank capacity rotor thrust niter counter written if Preq trace tolerance (fraction reference) relaxation factors maximum increment fraction (0. for no limit) with use then turn on trace selectively for mission/segment/condition trace iteration (0 for none, 2 for power) trace operation (0 for none, 1 trace, 2 for all iterations) counter at start trace of iterations counter check Case Flight condition and mission segment + + + + + + + + + + + + + + + + + + + + + real real real real real real real real real real real real real int int int int real size DGW xmsn wmto tank thrust _ _ _ _ _ _ size case start count size size DGW xmsn wmto tank thrust check _ _ _ _ _ _ _ _ _ _ _ _ Structure: Solution toler relax relax relax relax relax relax maxinc maxinc maxinc maxinc maxinc maxinc trace trace trace trace toler 1 1 1 1 1 ) /σ T C scale scale /σ T L weight scale (1 design gross weight, 2 nominal power scale (1 aircraft power, 2 derived from weight scale) length scale (1 rotor radius, 2 wing span, 3 fuselage length) rotor number wing number weight scale power scale length scale angle scale force scale moment scale horizontal velocity scale vertical velocity scale angular velocity scale C C altitude scale acceleration scale range scale Tolerance and perturbation scales Derived tolerance and perturbation scales + + + + + + int int int int int real real real real real real real real real real real real real real Wscale Pscale Lscale _ _ _ Structure: Solution KIND KIND KIND scaleRotor scaleWing Wscale Pscale Lscale Ascale Fscale Mscale Vscale Rscale Oscale Tscale Cscale Hscale Gscale Xscale 3 1 1 0.

5.0 123 100 0.04 100.00 Default inf _ year ) (1994) table /F ) inf _ ) ($/gal)/126.2 (based on 42.8 MJ/kg and 6.5 lb/gal of JP-4/JP-8) year ∼ = inf ( _ table year F ( (relative 1994) inflation = i F inflation pass = N ($/MJ) always used, even with internal table i F if no credit for generated energy ($/gallon or $/liter) energy multiplies airframe purchase price and maintenance cost i G inflation fuel F G factor CPI or DoD table: input factor: inflation: cost factors and rates include technology and inflation, correspond to equivalent energy price for fuel burned: $/MJ EnergyCredit=0.

title notes model (1 only input factor; 2 CPI; 3 DoD) year for internal inflation factor inflation factor (per cent, relative 1994 or year beyond CPI/DoD table data (0 error, 1 extrapolate factor) model (1 CTM) CTM cost model fuel price energy price credit for generated energy ($/MJ) number of passengers Cost Inflation Cost Description + + + + + + + + + + + + + + Type c*100 c*1000 int int real int int CostCTM real real real int inf inf _ cost _ _ inf _ Chapter 32 Structure: Cost Variable title notes MODEL year inflation EXTRAP MODEL CostCTM FuelPrice(ntankmax) EnergyPrice(ntankmax) EnergyCredit(ntankmax) Npass 8.

12. 15. 15. 10. 25. 1.0 0.87 3751.

B (min) N F T (years) (% purchase price) D S (%) maint V (years) (%) χ i L AF χ available block hours per year non-flight time per trip depreciation period loan period interest rate residual value spares per aircraft airframe maintenance Direct Operating Cost Technology Factors + + + + + + + + + + + real real real real real real real real real af maint _ _ cost cost _ _ Structure: Cost BlockHours NonFlightTime DepPeriod LoanPeriod IntRate ResidValue Spares TECH TECH 1 1 2 0.

0.0 0.0 0.0 0.0 0.0 34. 18.

160. 1.45 10000. 10000. 0.0017 Default inf _ year major M for composite construction ($/lb or $/kg) engine comp r M , mission equipment package ($/lb or $/kg) , flight control electronics ($/lb or $/kg) labor M FCD MEP r r parts M negative for cost reduction additional cost rate composite weight in body (fraction body weight) composite weight in tail (fraction tail weight) composite weight in pylon (fraction pylon weight) composite weight in wing (fraction wing weight) cost factor cost factor cost factors and rates include technology and inflation, correspond to rComp aircraft (1 rotorcraft, 2 turboprop airliner) engine (1 turbine, 2 piston) airframe systems (fixed useful load) maintenance cost estimate (1 total only, 2 separate components) labor rate ($ per hour) maintenance man hours per flight hour MMH/FH factor parts factor engine overhaul factor major periodic maintenance factor CTM rotorcraft cost model Purchase Price Maintenance Description + + + + + + + + + + + + + + + + + + + + + Type int int real real real real real real real int real real real real real real body tail pylon wing aircraft maint _ _ _ _ _ _ engine _ Chapter 33 Structure: CostCTM Variable MODEL KIND rComp fWcomp fWcomp fWcomp fWcomp rFCE rMEP MODEL rLabor MMHperFH Mlabor Mparts Mengine Mmajor 1.0 1.0 ) and term scaling with weight MMHperFH Mmajor=28 , Mmajor=18 , Mengine=1.74 , Mengine=1.45 inf , _ year Mparts=56 , Mparts=34 cdi , K Mlabor=0.0027 Mlabor=0.0017 crew K ) Mlabor labor rate includes inflation, corresponds to current best practice: current average practice: maintenance man hours per flight hour calculated from sum of fixed term ( empty ( crew+depreciation+insurance estimate (1 total only, 2 separate components) crew+depreciation+insurance factor crew cost factor Direct Operating Cost + + + + int real real doc _ Structure: CostCTM MODEL Kcdi Kcrew 1 1 0. 0.

30. 1.0 3.75 0.14 350. 0.03 500. 3.16 1.26 0.01 0.14 0.004 0.0002 Default 0.00004 0.00748 0.0036739 , 1 DLR, 2 Swiss) x NO EI (kg/MJ) (kg/kg) energy AIC f K fuel K (yr) (yr) x H 2 (missions/yr) max NO t CO U O CO 0 1 2 O EI EI 4 K H EIs EI EI 4 H r K SO EI K K K SO EI K soot EI emissions from fuel used, emissions from energy used, carbon dioxide, water vapor, sulphates, soot, nitrogen oxides, model parameters (0 input, 1 low emissions, 2 high emissions) DLR model, DLR model, Swiss model, carbon dioxide, water vapor, sulphates, 2 2 title notes CO CO aircraft operating lifetime aircraft utilization rate ATR discount rate ATR integration period emission index (kg/kg) turboshaft engine NOx emission model (0 input aviation induced cloudiness factor, energy emission factor (kg/MJ) Emissions Emissions Trading Scheme (ETS) Average Temperature Response (ATR) Description + + + + + + + + + + + + + + + + + + + + + + + + + + + Type c*100 c*1000 real real real real real real real real real real real int int real real real real real real real NOx(ntankmax) _ NOx(ntankmax) _ CO2(ntankmax) H2O(ntankmax) SO4(ntankmax) soot(ntankmax) NOx(ntankmax) CO2(ntankmax) H2O(ntankmax) SO4(ntankmax) _ _ _ _ _ _ _ _ Chapter 34 Structure: Emissions Variable title notes Kfuel(ntankmax) Kenergy(ntankmax) H U r tmax EI EI EI EI EI MODEL KIND KEI0(ntankmax) KEI1(ntankmax) KEIs(ntankmax) fAIC K K K 0. 0.

) burn=2 _ ) SET burn=1 _ SET ∞ = r ) only used for tanks that store and use fuel as energy ( energy evaluated as K ) only used for tanks that store and use fuel as weight ( and fuel K x 100000 K to ≥ NO P and r P K ) f L EI = q P model soot and O ˙ x w K default values are for turboshaft engine (CH soot, nitrogen oxides, 2 x EI emission index ( energy emission factor ( ATR discount rate: CO NO short life power factor, fuel flow, Emissions credit for energy generated (0 for none) ATR factors turboshaft NO + + + real real int real real real real real credit _ soot(ntankmax) NOx(ntankmax) _ _ Structure: Emissions K K SET ZCO2 ZNOx Zs fPower(11,nengmax) wdot(11,nengmax) 4 1 0 0 0 Default ’helicopter’ ) airplane ’airplane’ , , ) multicopter , nvelmax ’multicopter’ , compound , ’compound’ , tiltrotor , ’tiltrotor’ , coaxial , 2 piecewise linear, maximum ) ≥ ’coaxial’ , ) tandem , ncontmax ’tandem’ , nstatemax helicopter , ’helicopter’ , rotorcraft _ RCconfig ’rotorcraft’ = collective stick lateral cyclic stick longitudinal stick : identifies rotorcraft configuration config labels of aircraft controls collective stick lateral cyclic stick longitudinal stick pedal tilt number of speeds (0 zero value; 1 constant; config title notes configuration ( number of main rotors number of aircraft controls (maximum number of control states (maximum pilot’s controls (control number) control values (function speed) Aircraft Configuration Aircraft Controls Description + + + + + + + + + + + + + Type c*100 c*1000 c*16 int int int c*16 int int int int int int int int int int control(ncontmax) main _ control _ _ Chapter 35 Structure: Aircraft Variable title notes config RCconfig nRotor ncontrol IDENT nstate kcoll klatcyc klngcyc kpedal ktilt nVcont(ncontmax) nVcoll nVlatcyc nVlngcyc 0 0 control _ 0 control IDENT c _ + c comp AC _ T c , based on SET = c Vcont - cont - control=’coll’,’latcyc’,’lngcyc’,’pedal’,’tilt’ _ nVcont IDENT (for each control state) and value T defined AC c control can be suppressed for flight state using _ input put in appropriate c ACs tilt c / IDENT ACc pedal c / AC c lngcyc / ACp c latcyc / tilt α flight state specifies control state, or that control state obtained from conversion schedule use of component control initial values specified if control is trim variable; otherwise fixed for flight state coll can be zero, constant, or function of flight speed (CAS or TAS, piecewise linear input) pedal tilt collective stick lateral cyclic stick longitudinal cyclic stick pedal tilt collective stick lateral cyclic stick longitudinal cyclic stick pedal tilt aircraft controls connected to individual controls of each component, for each component control, define matrix c typical aircraft controls are pilot’s controls; default available for trim (flight state specifies trim option) each aircraft control can be zero, constant, or function of flight speed (CAS or TAS, piecewise linear input) flight state input can override by connecting aircraft control to component control, flight state can specify component control value rotor controls are positive Fourier series, with azimuth measured in direction of rotation values speeds (CAS or TAS) control system: set of aircraft controls aircraft controls: identified by sign conventions for pilot’s controls: collective + up, lat cyclic + right, long cyclic + forward, pedal + nose right + + + + + + + + + + + + + + int int real real real real real real real real real real real real Structure: Aircraft nVpedal nVtilt cont(nvelmax,ncontmax) coll(nvelmax) latcyc(nvelmax) lngcyc(nvelmax) pedal(nvelmax) tilt(nvelmax) Vcont(nvelmax,ncontmax) Vcoll(nvelmax) Vlatcyc(nvelmax) Vlngcyc(nvelmax) Vpedal(nvelmax) Vtilt(nvelmax) 1 1 1 1 1 1 hover _ kcont , ) ) hover _ cruise _ nvelmax nvelmax kgear conv , _ FltState kcont , of hover kcont _ , cruise Vtip _ _ Vtip , conv _ SET kgear , , tilt kgear tilt=90 _ , cruise ) ) _ SET , hover ): use _ 2 piecewise linear, maximum 2 piecewise linear, maximum Vtip cruise cruise ) ) , − − ≥ ≥ Vtip hover control − hover hover conv conv _ tilt=0 − − , use conv V V conv conv V V STATE ): use ≤ < V < V ) < V < V ) ≤ ≤ V V V cruise − cruise cruise hover hover − − linear with − − conv V tilt conv conv conv conv V V V V ≥ F ≥ ≥ V θ F φ V V available for trim (depending on flight state) each motion can be zero, constant, or function of flight speed (CAS or TAS, piecewise linear input) flight state input can override; initial value if trim variable hover and helicopter mode ( cruise mode ( conversion mode: nacelle tilt angle: 0 for cruise, 90 deg for helicopter mode flight number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) hover and helicopter mode ( conversion mode ( cruise mode ( hover and helicopter mode ( conversion mode ( cruise mode ( aircraft motion conversion control: use depends on aircraft pitch angle aircraft roll angle maximum speed for hover and helicopter mode (CAS or TAS) minimum speed for cruise (CAS or TAS) control state drive system state (each propulsion group) Aircraft Motion Conversion + + + + + + + + + + + + + + + + + + + + int real real int real real real real int int int int int int hover cruise hover conv cruise hover(npropmax) conv(npropmax) cruise(npropmax) _ _ _ _ _ _ _ _ Structure: Aircraft nVpitch pitch(nvelmax) Vpitch(nvelmax) nVroll roll(nvelmax) Vroll(nvelmax) Vconv Vconv kcont kcont kcont kgear kgear kgear 1 1 0 1 ), trim _ IDENT match ) ) trim _ ) mtrimmax xxx , or control number) _ STATE ntrimstatemax xxx _ QUANT VAR _ _ , 2 component) TRIM TRIM FltState identifies trim state ( conversion, controls and motion, rotor tip speed, landing gear retraction, trim targets FltState velocity schedules: all described as function CAS or TAS trim state: one or more set of quantities and variables for trim iteration number of trim states (maximum label of trim state number of trim variables (maximum trim quantity name trim variable name target source (1 trim quantity name ( trim quantity structure number trim quantity kind (0 other, 1 rotor, 2 rotor lift, 3 rotor prop, 4 wing, 5 wing lift) trim variable name ( trim variable structure number Velocity schedules (1 CAS, 2 TAS) Trim states Derived trim states + + + + + + + + int int c*12 int c*16 c*16 int int int int int int trim(ntrimstatemax) trim _ _ quant(mtrimmax,ntrimstatemax) quantn(mtrimmax,ntrimstatemax) quantk(mtrimmax,ntrimstatemax) var(mtrimmax,ntrimstatemax) varn(mtrimmax,ntrimstatemax) Vschedule quant(mtrimmax,ntrimstatemax) var(mtrimmax,ntrimstatemax) target(mtrimmax,ntrimstatemax) _ _ _ _ _ _ _ _ _ Structure: Aircraft SET nstate IDENT mtrim(ntrimstatemax) trim trim trim itrim itrim itrim itrim itrim Rotor%Klift Rotor%Klift Rotor%Kprop Rotor%Kprop Rotor%Klift Wing%Klift Wing%Klift , , , , , , , target target target target target target target target target target target target target target target target target target target target target target target target target target target _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ target zero zero FltState%trim FltState%trim FltState%trim FltState%trim FltState%trim FltState%trim FltState%trim FltState%trim FltState%trim FltState%trim FltState%trim FltState%trim FltState%trim FltState%trim FltState%trim FltState%trim FltState%trim FltState%trim FltState%trim FltState%trim FltState%trim FltState%trim FltState%trim FltState%trim FltState%trim FltState%trim FltState%trim ’moment z’ , ’force z’ , ’fprop rotor n’ ’X/q rotor n’ ’vert rotor n’ ’T margin tran n’ , ’flift rotor n’ , ’flift wing n’ ’moment y’ , , ’roll n’ ’pitch n’ ’torque n’ , , , , , , ’lngflap n’ ’latflap n’ ’nz’ ’force y’ , , , , quant ’ny’ _ , trim ’force x’ ’moment x’ ’nx’ ’power n’ ’P margin n’ ’Q margin n’ ’power EG n’ ’PEG margin n’ ’jet n’ ’J margin n’ ’charge n’ ’C margin n’ ’tank n’ ’B margin n’ ’lift rotor n’ ’CLs rotor n’ ’prop rotor n’ ’CXs rotor n’ ’CTs rotor n’ ’T margin n’ ’betac n’ ’betas n’ ’hub Mx n’ ’hub My n’ ’hub Mz n’ ’lift wing n’ ’CL wing n’ ’L margin n’ ’lift tail n’ description aircraft total force aircraft total moment aircraft load factor propulsion group power power margin torque margin engine group power power margin jet group thrust jet thrust margin charge group power charge power margin fuel tank energy flow battery power margin rotor lift rotor lift rotor propulsive force rotor propulsive force rotor thrust rotor thrust margin rotor flapping rotor flapping rotor hub moment rotor hub moment rotor torque wing lift wing lift coefficient wing lift margin tail lift trim quantity: Structure: Aircraft 2 1 1 0 1 used ) Kprop tran _ or Klift GW*nAC(3) Rotor%CTs uses body axes relative inertial axes horizontal, vertical flight speed sideslip angle Euler angle rates ) ; otherwise component qAC*DoQ control _ target _ ’turn’ ’ROC’ ’roll’ quant=’T margin tran n’ , IDENT , , _ Vtip=’input’ var ) _ : target is fraction total aircraft lift ( _ trim ) ZoL , / trim match ’pitch’ ’speed’ ’side’ ’pullup’ ’Vtip n’ ’Nspec n’ FltState%trim n=1 YoL / steady _ XoL FltAircraft%SET quant=’flift wing n’ _ Rotor%CTs trim : target is fraction total aircraft drag ( : requires or uses ; 2 scaled, from WL / ’Nspec’ BL , trim quantity target value is or / SL ” is absent, use first component ( description aircraft control aircraft orientation aircraft velocity aircraft velocity aircraft angular rate propulsion group tip speed propulsion group engine speed n target=1 kind (1 rotor radius, 2 wing span, 3 fuselage length) identification (component number) _ quant=’flift rotor n’ quant=’fprop rotor n’ quant=’T margin n’ var=’Vtip’ _ _ _ _ reference length kind (0 input, 1 rotor, 2 wing, 3 fuselage, 4 center of gravity) identification (component number) stationline buttline waterline trim trim variable: if if trailing “ trim trim trim trim input (1 fixed, scaled geometry reference point Geometry + + + + + + + + + + + + int int int int int real real real geom _ scale Ref _ _ Ref Ref Ref _ _ _ Structure: Aircraft INPUT KIND kScale KIND kRef SL BL WL 0.

’std’ ) ) scale _ kScale , KIND scale ) _ geom KIND _ INPUT WL / , or position of identified component BL / Ref SL _ XX in Location override this global scale _ in Location override ): divided by reference length ( ): dimensional KIND geom _ FIX ; input dimensional geom = 2 + up; from reference point _ geom = 1 _ kRef ZoL , INPUT Ref T INPUT _ ) Δ + right, xxx _ KIND YoL s c atmos T /T 0 _ τ + aft, p/p μ ρ/ρ SET ρ stationline + aft, buttline + right, waterline + up; arbitary origin; units = ft or m XoL option to fix some geometry ( option to specify reference length ( component reference must be fixed ρ fixed geometry input ( scaled geometry input ( reference point: certain Locations can be calculated from other parameters (configuration specific) flight state has calculated or input actual cg location stationline buttline waterline temperature temperature increment density speed of sound viscosity Geometry: Location for each component center of gravity: baseline is for nacelle angle = 90 calculated reference point (input or component) baseline center of gravity location atmosphere specification altitude atmosphere ( density density ratio temperature ratio pressure ratio Takeoff flight condition Derived takeoff flight condition + + + + + + + + + real real real Location c*12 real real real real real real int real real real real to _ to to atmos to _ atmos _ _ _ _ cg _ Structure: Aircraft SLref BLref WLref loc SET temp dtemp density csound viscosity altitude iSET density sigma theta delta ) ) dtemp , temp ) , /S altitude D altitude /A viscosity W , D W W f W f csound , ) density ) ) viscosity temp (each engine group at takeoff rating) , (each jet group at takeoff rating) ) , jet eng T P altitude csound ; rotor disk loading = jet density , Systems eng A N ; individual wing loading = A N S f ∑ in rotor sizing ∑ ∑ density ∑ = = /σ = = av T T av C ref , ref P S , A av , , av /T ref for other options (polar, tropical, and hot days) ref /P D /S D /A W D D W atmos W W _ = standard day at specified altitude, plus temperature increment (use = input, not air on earth (use = standard day at specified altitude, and specified temperature (use = input density, speed of sound, and viscosity (use , atmosphere specification: = input density and temperature (use = standard day at specified altitude (use FltState%SET atmos ’std’ ’dtemp’ ’temp’ ’dens’ ’input’ ’notair’ see _ takeoff condition (density) used for SET aircraft disk loading = aircraft wing loading = aircraft power loading = aircraft thrust-to-weight = aircraft disk loading reference rotor area aircraft wing loading reference wing area total takeoff power available aircraft power loading total takeoff thrust available aircraft weight-to-thrust wing extensions (0 for none) wing extension kits (0 for none) wing kits (0 for none) other kit (0 for none) folding (0 none, 1 fold weights, 2 with kit) (from Size Configuration real real real real real real real real int int int int int fold _ Structure: Aircraft diskload Aref wingload Sref Pav powerload Tav thrustload nWingExt nWingExtKit nWingKit nWotherkit SET ) FCE W inflation + MEP W + AF W R = MEP K W MEP FCE major LG C C W K FCE comp C W EN comp relative 1994, including factor C K ($/lb or $/kg) inf _ ET engine ($/lb or $/kg) EK C K , folding kit not installed P year AF ( (2011 relative 1994, CPI) /W AF i i W F F maint maint AC na /W C C , folding kit not installed C labor parts SL AF C C AC AC C weight empty + airframe kits C C = EK stationline length width area first fuel tank that burns weight (0 none) reference specific energy (MJ/kg) aircraft aircraft maintenance maintenance inflation factor inflation factor composite cost increment mission equipment package cost flight control electronics cost composite weight increment mission equipment package weight flight control electronics weight configuration factor, airframe total aircraft airframe weight W rated takeoff power labor cost parts cost engine overhaul cost major periodic maintenance cost maintenance man hours per flight hour Neutral point Operating size (hover; controls = 0 except tilt = 90) Fuel tank system Cost real real real real int real real real real real real real real real real real real real real real real real real real real real real real real nokit op _ op inf inf2011 _ nokit _ _ _ op _ _ Structure: Aircraft SLna length width area burnweight eref CAC CAC Cmaint Cmaint factor factor Ccomp CMEP CFCE Wcomp WMEP WFCE Kconfig rAF rAC WAFcost WEKcost Pcost Clabor Cparts Cengine Cmajor MMHperFH 0. 1. 1. 0. 1.

6.0 cap − fuel ), maximum W * SDGW _ ) SET ( fFuelSDGW + sdgw _ SD W DGW _ DESIGN Wfuel – D W = G ) W , M T O WE G _ W * W D * FIX ( W * E ), structural design gross weight fSDGW W + fSDGW + DGW fSDGW _ + FIX ( dSDGW = dSDGW D = dSDGW W SD = SD at SDGW W ), weight empty W SD SD ult W z W ): adjust contingency weight to achieve n M T O usually calculated (identified as input so inherited by next case) WMTO W _ WE=1 _ SET DGW ( _ corresponding to DGW D FIX W fuel corresponding to DGW M T O Wpay = based on WMTO; W W = based on DGW; = based on fuel state; pay and , structural design gross weight: = input W = calculated from maximum gross weight at SDGW sizing conditions ( DGW _ , design gross weight: used for rotor disk loading and blade loading, wing loading, power loading, thrust loading SDGW if calculated, then input parameter is initial value to obtain aircraft moments of inertia from radii of gyration for tolerance and perturbation scales of the solution procedures optionally to define structural design gross weight and maximum takeoff weight optionally to specify the gross weight for missions and flight conditions ’input’ ’f(DGW)’ ’f(WMTO)’ ’maxfuel’ ’perf’ _ mission fuel payload structural design gross weight gross weight increment gross weight factor fraction main fuel tanks filled at SDGW maximum takeoff weight gross weight increment gross weight factor input or calculated: design gross weight takeoff weight fixed weight empty ( DGW Wfuel SET design gross weight structural design gross weight maximum takeoff weight design ultimate flight load factor Weight + + + + + + + + + + + + + + real real real real real real real real real real real DGW DGW _ _ ult _ Structure: Aircraft DGW Wfuel Wpay SDGW dSDGW fSDGW fFuelSDGW WMTO dWMTO fWMTO nz cap − fuel W + ) DGW _ wmto _ Wfuel – D W DESIGN = G W , G W * SD W D * W fWMTO * + fWMTO + fWMTO + dWMTO = ) fuel dWMTO M T O = W dWMTO W = − M T O W M T O scaled W W − /L /L y /L D z k x k W k ( / O D W E W W = based on SDGW; = based on DGW; = based on maximum fuel; , maximum takeoff weight: = input = calculated from maximum gross weight at WMTO sizing conditions ( , design ultimate flight load factor at SDGW: used for weights (fuselage, rotor, wing) WMTO ’input’ ’f(DGW)’ ’f(SDGW)’ ’maxfuel’ ’perf’ so for flight state, additional fixed useful load = auxiliary fuel tank and kits and increments flight state can also increment crew weight or equipment weight _ ult operating weight weight empty growth factor = roll radius of gyration pitch radius of gyration yaw radius of gyration _ SET SDGW used for weights (fuselage, rotor, wing) WMTO used for cost, drag (scaled aircraft and hub drag), and weights (system, fuselage, landing gear, engine group) nz weight empty = structure + propulsion + systems and equipment + vibration + contingency operating weight = weight empty + fixed useful load weight statement defines fixed useful load and operating weight for design configuration xx yy zz xy yz xz aircraft weight statement (operating weight, without payload and usable fuel) moments of inertia (based on design gross weight, scaled with reference length) I I I I I I Weight Derived moments of inertia (corresponding to aircraft weight statement) + + + + + + Weight real real real real real real real real real real real real factor _ Structure: Aircraft Weight WO WE growth kx ky kz Ixx Iyy Izz Ixy Iyz Ixz 0 0 2 0. 0.

2.5 0.05 0.008 ) k ) / V DL ) k /Mg ; 3 scaled, from D/q ( D C or m or / DL/T D/q /klb ∼ ; 2 scaled, from ) V D kDL C are ft D/q ; 2 scaled, input ref ) (others calculated) A kDrag D/q = Dscale kDrag _ , D/q (other calculated) CD Units , ( kDL / 2 , DoQ = reference rotor area, ref 1000) A DoQV / ref /A M T O (based on rotor area, V (0 calculated; 1 fixed, input W ) D = 9 for old helicopter, 2.5 for current low drag helicopters, ( = reference rotor area; units of / V C : minimum drag, excludes drag due to lift and angle of attack 1.6 for current tiltrotors, 1.4 for turboprop aircraft (English units) ) ) = 0.02 for old helicopter, 0.008 for current low drag helicopters , download: D/q D/q ref drag DL use only one of input A CD kDrag use only one of D/q D/q _ _ = ( D/q ( gross weight = weight empty + useful load = operating weight + payload + usable fuel useful load = fixed useful load + payload + usable fuel FIX FIX = ( DL area coefficient k area k flight state: gross weight, useful load (payload, usable fuel, fixed useful load), operating weight fixed drag or download: obtained by adjusting contingency angle of attack and sideslip angle: reversed definition best for sideward flight total aircraft total aircraft download (0 calculated; 1 fixed, input angle of attack and sideslip angle representation (1 conventional, 2 reversed for sideward flight) Drag Aerodynamics + + + + + + + + + + int real real real int real real int alpha _ drag DL _ _ Structure: Aircraft FIX DoQ CD kDrag FIX DoQV kDL KIND 2 0 1 1 1 1 0 0 1 0 0 0 0 0 0 0 0 (without contingency) ) (without contingency) (without contingency) comp ) comp ) nengmax comp / ) ) ) 2 ) D/q ) ( ) ) / D/q wet ) ( 2 ) ) ) D/q ) ) ( 1000) AC / ) nengmax D/q npropmax AC ) ) ntankmax ( ) 1000) nchrgmax AC ref nchrgmax / ) ) ) D/q M T O ntankmax nengmax nengmax ( nchrgmax nengmax /A nengmax D/q ) ref W ref wet ( ( D/q M T O njetmax ( AC / njetmax /A /A /S ) ) W V ( ) wet AC / wet nrotormax nwingmax ) ) S ) D/q D/q ntailmax ( ( D/q ( D/q D/q D/q ( ( ( sum component cruise drag, area sum component helicopter drag, area sum component vertical drag, area total cruise drag, area total helicopter drag, area total vertical drag, area total cruise total helicopter total cruise total helicopter total vertical total cruise wetted drag, area total wetted area total cruise rotors (maximum wings (maximum tails (maximum fuel tank systems (maximum propulsion groups (maximum engine groups (maximum jet groups (maximum charge groups (maximum engine models (maximum engine tables (maximum reciprocating engine models (maximum compressor models (maximum motor models (maximum jet models (maximum fuel cell models (maximum solar cell models (maximum battery models (maximum Derived aircraft drag Number of Components + + + + + + + + + + + + + + + + + + real real real real real real real real real real real real real real int int int int int int int int int int int int int int int int int AC AC AC _ _ comp AC _ comp comp AC AC _ _ _ _ _ _ AC AC AC AC _ _ _ AC _ _ Structure: Aircraft DoQC DoQH DoQV DoQC DoQH DoQV CDC CDH kDragC kDragH kDL DoQwet Swet CD nRotor nWing nTail nTank nPropulsion nEngineGroup nJetGroup nChargeGroup nEngineModel nEngineTable nRecipModel nCompressorModel nMotorModel nJetModel nFuelCellModel nSolarCellModel nBatteryModel used in one or more engine groups ParamN propulsion group is set of components and engine groups, connected by drive system engine model or engine table or reciprocating engine or motor model describes particular engine, jet model describes particular jet, used in one or more jet groups fuel cell model or solar cell model describes particular charger, used in one or more charge groups battery model describes particular batter, used in one or more fuel tanks Aircraft Systems Fuselage LandingGear Rotor Wing Tail FuelTank Propulsion EngineGroup JetGroup ChargeGroup EngineModel EngineTable RecipModel CompressorModel MotorModel JetModel FuelCellModel SolarCellModel BatteryModel Aircraft Input for case int int int int int int int int int int int int int int int int int int int int int int Structure: Aircraft inAircraft inSystems inFuselage inLandingGear inRotor(nrotormax) inWing(nwingmax) inTail(ntailmax) inFuelTank(ntankmax) inPropulsion(npropmax) inEngineGroup(nengmax) inJetGroup(njetmax) inChargeGroup(nchrgmax) inEngineModel(nengmax) inEngineParamN(nspeedmax,nengmax) inEngineTable(nengmax) inRecipModel(nengmax) inCompressorModel(nengmax) inMotorModel(nengmax) inJetModel(njetmax) inFuelCellModel(nchrgmax) inSolarCellModel(nchrgmax) inBatteryModel(ntankmax) ) span _ not ) ) ) , ) perf Preq perf , , panel , radius , _ ) Pav ) , hub not , , maxfuel none maxfuel ) , , fmiss , ratio , ) , scale width ) none , , ) , param none input , fSDGW , area _ fWMTO , ratio _ none , , radius , param , misspower rotor not limit , , , DL , _ thrust , , fDGW ratio fDGW , WL param chrg miss ) , , SET _ , , , engine radius _ radius jet Size input _ _ _ span area _ input _ _ input _ perf jet charge _ _ _ _ rotor rotor _ _ wing wing _ tank _ _ WMTO SIZE SIZE SIZE SDGW _ SET _ (1 fixed, 0 not) SIZE SET SET SET (1 fixed, 0 not) /σ SET SET (1 fixed, 0 not) W tip C V σ Cost Emissions performance ( performance ( performance ( rotor sized ( rotor radius ( rotor rotor rotor wing area ( wing span ( wing chord (1 fixed, 0 not) wing aspect ratio (1 fixed, 0 not) design gross weight (0 calculated, 1 fixed) weight empty (0 calculated, 1 fixed) fuel tank ( SDGW ( WMTO ( drive system torque limit ( kind iteration, performance (0 none, 1 size engine or radius or jet group or charge group) kind iteration, parameters (0 none, 1 calculate parameters) conditions and missions for size engine or rotor conditions and missions for size jet group conditions and missions for size charge group design conditions and missions for DGW design conditions and missions for transmission design conditions for WMTO design missions for fuel tank design conditions and missions for antitorque or aux thrust rotor design gross weight source (1 condition, 2 mission) Design specification (from Design data (from sizing) int int int int int int int int int int int int int int int int int int int int int int int int int int int int int int int radius(nrotormax) area(nwingmax) span(nwingmax) ds(npropmax) _ CWs(nrotormax) Vtip(nrotormax) sigma(nrotormax) _ _ _ GW xmsn(npropmax) wmto tank thrust chord(nwingmax) AR(nwingmax) _ _ _ size param _ _ _ _ _ _ _ _ _ engine(npropmax) jet(njetmax) charge(nchrgmax) source perf(npropmax) jet(njetmax) charge(nchrgmax) rotor(nrotormax) _ _ _ rotor wing wing tank(ntankmax) SDGW WMTO limit _ _ _ _ _ iter iter _ _ _ _ _ _ _ rotor rotor rotor wing wing DGW WE _ _ _ _ _ _ _ _ _ Structure: Aircraft inCost inEmissions iSIZE iSIZE iSIZE iSIZE iSET FIX FIX FIX iSET iSET FIX FIX FIX FIX iSET iSET iSET iSET kind kind nSIZE nSIZE nSIZE nDESIGN nDESIGN nDESIGN nDESIGN nDESIGN DGW ) ndesignmax design gross weight source number design gross weight segment number number design of conditions and missions (maximum design data int int int XAircraft kState kSeg _ _ Structure: Aircraft DGW DGW nDesignState XAircraft(ndesignmax) Default (each aux tank size) auxtank N load − ) ext ammo pay pass W ; at segment start W W weapons G W W W cargo other FltAircraft W W passengers cargo external load ammunition weapons other source (1 condition, 2 mission) source number segment number title kind (condition or mission) number (condition or mission/segment) label Set Gross Weight Set Useful Load design number of auxiliary fuel tanks number of crew number of passengers number of crew seats number of passenger seats kits gross weight payload weight Design Data Weights (from Description Type int int int c*100 c*12 c*12 c*12 c*12 c*12 c*12 int int int int int c*12 real real real real real real real real seat seat pass cargo extload ammo weapons other _ _ _ _ _ _ _ _ Chapter 36 Structure: XAircraft Variable source kState kSeg title kind number label setgw setul design Nauxtank(nauxtankmax,ntankmax) Ncrew Npass Ncrew Npass kits GW Wpayload Wpay Wpay Wpay Wpay Wpay Wpay ) ) Aircraft ; at segment start ; at segment start F U L (from fuel W E Aircraft%Weight fuel ) O W E W W U L catagories W FltAircraft crew fluids (from auxiliary fuel tanks other fixed useful load equipment increment folding kit wing extension kit wing kit other kit usable fuel weight standard tanks auxiliary tanks weight empty fixed useful load usable fuel energy standard tanks auxiliary tanks usable fuel weight operating weight useful load military load usable fuel energy Energy (from real real real real real real real real real real real real real real real real real real real real real real real fluid other _ _ total std(ntankmax) aux(ntankmax) total std(ntankmax) aux(ntankmax) _ _ _ _ _ _ fixUL fixUL _ _ Structure: XAircraft Wfuel Wfuel(ntankmax) Wfuel Wfuel WO WE WFixUL Wcrew W Wauxtank W Woful(10) Wequip Wfoldkit Wextkit Wwingkit Wotherkit WUL WML Efuel Efuel(ntankmax) Efuel Efuel 1 1 0 0. 0. 0.

’ ’ Default ) Ucrew*Ncrew+Wcrew ) or all terms ( Wcrew : payload specified by flight condition or mission : no details (only weight or adjustment weight per crew number of crew number of categories (0 for one value without name; maximum 10) category name baseline weight Wpayload Wcrew _ _ weight per passenger crew weight (1 no details; 2 all terms) trapped fluids and engine oil weight other fixed useful load other kit SET SET other fixed useful load: can include baggage, gun installations, weapons provisions, aircraft survivability equipment, survival kits, life rafts, oxygen title notes weight statement (systems) payload (1 no details; 2 all terms) fixed useful load Systems Weight Description + + + + + + + + + + + + + + + + + Type c*100 c*1000 Weight int real int real real int real int c*24 real real name(10) _ Wpayload Wcrew _ _ Chapter 37 Structure: Systems Variable title notes Weight SET Upass SET Wcrew Ucrew Ncrew Wtrap nWoful Woful Woful(10) Wotherkit 0 1 1 1 0. 0. 0. 0. 0. 0.

0.5 0.5 0.5 0.5 0.5 ) WE=1 _ FIX ( E cont W W + vib W for input cont W for no rotor/wing/tail/body fold weight , or adjust E W E cont W f fWxxfold=0 vib = f and = cont vib W (fraction wing/rotor/tail/body fold weight) vib cont W f f of fold weight in fixed useful load as kit, remainder kept in component weight foldkit input or dWxxfold=0 f input or cont vib W fWfoldkit vib cont W : : W W , folding: aircraft power supply Wvib Wcont fold = (structure + propulsion group + systems and equipment) + set component fraction kit weight removable, absent for specified flight conditions and missions _ _ _ wing rotor tail body (wing and rotor fold) body (tail fold) weight fraction weight empty weight fraction weight empty auxiliary power group (APU) instruments group pneumatic group environmental control group electrical group (1 no details; 2 all terms) E W SET SET SET folding (0 none, 1 fold weights, 2 with kit) folding weight in kit vibration treatment weight (1 fraction weight empty, 2 input) contingency weight (1 fraction weight empty, 2 input) systems and equipment + + + + + + + + + + + + + + + + + + + + + int real real real real real int real real int real real real real real real int real real supply _ fold Wvib Wcont Welectrical _ _ _ _ Structure: Systems SET fWfoldkitW(nwingmax) fWfoldkitR(nrotormax) fWfoldkitT(ntailmax) fWfoldkitFw fWfoldkitFt SET Wvib fWvib SET Wcont fWcont Wauxpower Winstrument Wpneumatic Wenviron SET Welectrical Welect 1 1 1 1 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0. 0.

armor cabin floor armor weight per area cabin wall armor weight per area armor weight per crew armament provisions gun provisions turret systems expendable weapons provisions each crew seat each passenger seat miscellaneous accommodation per crew seat miscellaneous accommodation per passenger seat oxygen system per crew seat oxygen system per passenger seat power conversion power distribution and controls lights and signal devices equipment supports avionics communications navigation identification control and display aircraft survivability mission system equipment armor (1 no details; 2 all terms) armament provisions (1 no details; 2 all terms) armament electronics (avionics group) furnishings and equipment accommodations for personnel miscellaneous equipment avionics group (1 no details; 2 all terms) armament group furnishings and equipment group (1 no details; 2 all terms) + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + real real real real int real real real real real real real int real real real real int real real real real real int real real real real real real real real gun turret expend com nav ident disp survive mission _ _ _ _ _ _ _ _ _ misc crew pass _ floor wall crew _ _ conv distrib lights support _ _ _ elect _ _ _ _ crew pass _ _ _ WMEQ Warmor Warmprov Wfurnish crew pass _ _ _ _ _ _ Structure: Systems Welect Welect Welect Welect SET WMEQ Wavionics Wavionics Wavionics Wavionics Wavionics Wavionics SET Warmor Uarmor Uarmor Uarmor SET Warmprov Warmprov Warmprov Warmprov Warm SET Wfurnish Useat Useat Uaccom Uaccom Uox Uox Wfurnish 1 0 0 0. 0. 0. 0. 0. 0. 0. 0. 0.

): armor=2 _ SET and furnish=2 _ SET crew _ crew+Uarmor _ pass _ crew+Uox pass+Uox _ _ ; equipment weights include installation crew+Uaccom pass+Uaccom _ _ Welectrical+WDIelect Warmprov Wfurnish Warmor WMEQ Wload : only : only : only inc=Useat inc=Useat : only : only _ _ : only seat seat trim acoustic and thermal insulation weight per cabin area fire detection and extinguishing other emergency equipment cargo handling weight per cabin floor area hoist external load provisions _ _ excluding vibration absorbers furnishings emergency equipment load and handling aircraft handling load handling Welectrical=1 WMEQ=1 Warmor=1 Warmprov=1 Wfurnish=1 Wload=1 miscellaneous accommodation includes galleys and toilets miscellaneous equipment includes cockpit displays trim includes floor covering, partitions, crash padding, acoustic and thermal insulation other emergency equipment includes first aid, survival kit, life raft Ucrew Upass _ _ _ _ _ _ load and handling group (1 no details; 2 all terms) number of crew seats number of passenger seats equipment weight increment per crew seat (0. for default) equipment weight increment per passenger seat (0. for default) SET SET SET SET SET SET equipment weight increment is for flight condition and mission; default (if systems and equipment + + + + + + + + + + + + + + + + + + real real real real int real real real real real int int real real inc inc aircraft _ cargo _ _ trim _ fire other _ _ _ hoist extprov seat seat seat seat _ _ _ _ _ _ Wload _ Structure: Systems Wfurnish Uinsulation Wemerg Wemerg SET Wload Whandling Uhandling Wload Wload Ncrew Npass Ucrew Upass 1 1 1 1 1 0. 0. 0. 0. 0.

reference rotor number for global rotary wing, boosted rotary wing, control boost mechanisms rotary wing, non-boosted fixed wing, control boost mechanisms fixed wing, non-boosted model (0 input, 1 NDARC, 2 custom) rotary wing flight controls (0 not present, 1 global, 2 for each rotor) fixed wing flight controls (0 for not present) conversion controls (0 for not present) flight control weight increment fuselage rotors wings tails cabin floor armor weight cabin wall armor weight crew armor weight seats miscellaneous accommodation oxygen system acoustic and thermal insulation weight flight control group and hydraulic group fixed useful load, fold kit armament group furnishings and equipment group cargo handling weight equipment weight increment per crew seat equipment weight increment per passenger seat weight on wing tip systems and equipment Derived weights Weight + + + + + + + + + + + + + + real real real real real real real real real real real real real real real int int int int int real real real real real fus rotor wing tail _ _ _ _ cargo _ b mb nb mb nb foldkit foldkit foldkit foldkit fc RWfc FWfc CVfc floor wall crew _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ fixUL fixUL fixUL fixUL _ _ _ _ Structure: Systems W W W W Warmor Warmor Warmor Wseat Waccom Wox Winsulation Whandling Ucrewseatinc Upassseatinc Wtip(nrotormax) MODEL MODEL refRotor MODEL MODEL dWRWfc dWRWfc dWRWfc dWFWfc dWFWfc 0. 0. 0. 0. 0. 0. 0. 0. 0. 0.

1.0 1.0 1.0 xx=0.

_ TECH or xx=0 _ MODEL ; for fixed (input) weight use dWxx + RW mb χ model _ Wxx * xx _ RW nb χ : global option is based on just main rotors RW b TECH χ conversion, boosted conversion, control boost mechanisms cockpit controls automatic flight control system rotary wing fixed wing conversion electrical system anti-ice system = RWfc _ fixed flight controls hydraulic weight increment equipment hydraulics NDARC model model (0 input, 1 NDARC, 2 custom) weight increment NDARC model Wxx “for each rotor” option sums separate contributions from all rotors sum rotary wing and conversion flight controls, hydraulic group, trapped fluids anti-icing group boosted control boost mechanisms non-boosted weight model result multiplied by technology factor and increment added: MODEL tiltrotor wing weight model requires weight on wing tip: distributed to designated rotor; rotary wing flight control weight Technology Factors + + + + + + + + + + + + + + + + + + + + real real real real real real real real WFltCont int real real WDeIce real real real b mb nb _ _ _ mb nb DI _ _ _ RWfc RWfc RWfc _ _ _ cc afcs _ _ Structure: Systems dWCVfc dWCVfc Wfc Wfc dWRWhyd dWFWhyd dWCVhyd WEQhyd WFltCont MODEL dWDIelect dWDIsys WDeIce TECH TECH TECH 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 F W mb CV mb χ χ elect sys DI hyd χ DI hyd hyd χ F W nb CV nb RW χ χ CV F W χ χ χ control boost mechanisms non-boosted control boost mechanisms non-boosted rotary wing fixed wing conversion electrical system anti-ice system fixed wing flight control weight conversion flight control weight flight control hydraulics anti-icing + + + + + + + + + + + + + real real real real real real real real real mb nb mb nb _ _ _ _ FWfc FWfc CVfc CVfc RWhyd FWhyd CVhyd DIelect DIsys _ _ _ _ _ _ _ _ _ Structure: Systems TECH TECH TECH TECH TECH TECH TECH TECH TECH 1 2 1 6 1 0.6 3.0 0.2 0.2 0.2 1.7 0.10 0.02 0.10 0.41 0.40 0.10 0.10 Default (fraction boosted + boost mech, or total) RW b (fraction total weight) f red RW f RW mb f (fraction boost mechanisms weight) RW nb f (fraction maximum takeoff weight) (fraction total fixed wing flight control weight) CV mb (fraction boost mechanisms weight) f , 2 scaled with DGW) F W nb CV nb (fraction rotary wing boost mechanisms + hydraulic weight) f f cc (fraction fixed wing boost mechanisms weight) (fraction conversion boost mechanisms weight) _ hyd hyd hyd Wfc RW CV F W f cc f f X cc K model (1 fraction, 2 parametric, 3 Boeing, 4 GARTEUR, 5 Tishchenko) AFDD: non-boosted control weight AFDD: hydraulic system redundancy/complexity factor AFDD: survivability (1 baseline, 2 UTTAS/AAH level of survivability) Boeing or GARTEUR or Tishchenko: boosted control weight GARTEUR or Tishchenko: boost mechanisms weight model (1 full controls, 2 only on horizontal tail, 3 GARTEUR, Raymer (4 transport, 5 general aviation)) non-boosted weight Raymer: number of control functions Raymer: number of mechanical functions (fraction total) boost mechanisms weight non-boosted weight model (1 fixed factor exponent rotary wing fixed wing conversion rotary wing flight controls fixed wing flight controls conversion controls cockpit controls flight control hydraulics Flight Control Group, NDARC Weight Model Hydraulic Group, NDARC Model Description + + + + + + + + + + + + + + + + + + + + + + + + + Type int real real int real real int real int real real real int real real real real real WRWfc WFWfc cc _ _ _ red nb b mb nb _ WRWfc mb nb _ _ _ _ _ _ _ Chapter 38 Structure: WFltCont Variable MODEL fRWfc xRWfc KIND fRWfc fRWfc MODEL fFWfc nfunction fmech fCVfc fCVfc MODEL Kcc Xcc fRWhyd fFWhyd fCVhyd 0.

) nb _ fRWfc = 0.4 fRWhyd = 0.6 (data range 0.3 to 1.8), = fraction: parametric except for non-boosted controls (from nb _ = 1.0 to 3.0 fRWfc WRWfc _ red _ flight controls = non-boosted (do not see aero surface or rotor loads) + boost mechanisms (actuators) + boosted MODEL typically xRWfc parameters Custom Weight Model + + real fc(8) _ Structure: WFltCont WtParam 0.

0.0 0.25 0.25 0.006 0.006 Default ) ) or kg/m or kg/m (lb/ft (lb/lb or kg/kg) elec air (lb/ft (lb/lb or kg/kg) K K jet rotor K (lb/ft or kg/m) K wing K weight factor for electrical system weight factor for main rotor weight factor for wing weight factor for engine air intake weight factor for jet air intake parameters Anti-Icing Group, NDARC Weight Model Custom Weight Model Description + + + + + + + + Type real real real real real real DI(8) _ elec(nrotormax) rotor(nrotormax) wing(nwingmax) air(nengmax) jet(njetmax) _ _ _ _ _ Chapter 39 Structure: WDeIce Variable kDeIce kDeIce kDeIce kDeIce kDeIce WtParam 1 1 1 2 1. 1.

Default (fraction fuselage length) boom ref f C boom proj w f wet f proj S wet S nose fus proj h aft S wet fus S w fus nose length nose length (fraction reference length) aft length aft length (fraction reference length) length nose length (distance forward of hub; 1 input, 2 calculated) aft length (distance aft of hub; 1 input, 2 calculated) fuselage SL location relative nose rotor-rotor longitudinal separation tail length (wing to horizontal tail) wetted area projected area factor for wetted area factor for projected area fuselage height tail boom effective circumference tail boom effective width input wetted area input projected area title notes fuselage location fuselage length (1 input, 2 calculated, 3 from rotor and tail only, 4 from rotor only) fuselage width fuselage wetted area (1 input, 2 input plus boom, 3 from nose length, 4 from fuselage length) Fuselage Geometry Description + + + + + + + + + + + + + + + + + + + + + + + Type c*100 c*1000 Location int real int real real int real real real real real real int real real real real real real real real real nose aft boom fus nose aft rotors tail fus _ _ fus boom _ _ _ _ _ _ _ _ _ in in length nose aft Swet fus _ _ _ _ _ _ _ fuselage _ Chapter 40 Structure: Fuselage Variable title notes loc SET Length SET Length fLength SET Length fLength fRef Length Length Width SET Swet Sproj fSwet fSproj Height Circum Width Swet Sproj 2 1 1 1 0.6 0.6 0.6 0.60 Swet and fus _ fus _ Length Length ), ); used for Sproj nose , ); used for _ aft , and fuselage or nose length) _ Swet fus _ fLength fLength not input (set to zero if no boom) Height , fus Swet _ _ ; used for operating length and sketch fus SET _ Width cabin , f floor wall − − ) or calculated (from nose and aft lengths) ) or calculated (from = input, otherwise calculated Length fus fSproj ) or calculated (from )/ _ cabin , cabin nose f f _ cabin aft _ length floor S wall nose _ _ − − fSwet Length SL Length SET – cabin Length cabin S S fuselage _ : cabin areas used for systems and equipment weights : input (use SL : both wetted area and projected area; input (use : input (use which can not then be scaled with fuselage length : input (use =( length nose aft Swet Scabin fus calculated uses rotor, tail, wing locations; or just rotor and tail, or just rotor input required if or calculated (from _ _ _ _ _ _ total cabin surface area cabin floor area cabin wall area factor for total cabin surface area factor for cabin floor area factor for cabin wall area rotor number (for rotor radius) wing number (for wing span) height of cargo ramp fraction of fuselage length used for cargo SET SET SET fRef SET boom circumference and width used if SET cabin area (1 input, 2 calculated) reference length (1 rotor radius, 2 wing span, 3 fuselage length) Geometry (for graphics) + + + + + + + + + + + + + int real real real real real real int int int real real cargo floor wall floor wall ramp _ _ _ scale _ _ _ _ Scabin _ Structure: Fuselage SET Scabin Scabin Scabin fScabin fScabin fScabin KIND refRotor refWing Height fLength 1 1 0. 0. 0. 0. 0. 0. 0. 0.

1.0 1.0 1.0 1.0 1.0 1.0 ); otherwise input DL _ ); otherwise input drag _ Aircraft FIX Aircraft FIX rb ) cont V ) fit ) fus D/q ( V ) D/q ( D/q ( cont ) D/q ( fus ) D/q cw ( χ D/q mar ( χ press χ calculated if total download fixed ( basic calculated if total drag fixed ( wfold χ χ tfold cont basic body body marinization pressurization body crashworthiness tail fold wing fold χ _ cont _ fuselage group model (0 input, 1 NDARC, 2 custom) weight increment AFFD model DoQ DoQV model (0 none, 1 standard) standard model contingency drag, area contingency vertical drag, area fuselage drag, area fuselage vertical drag, area fittings and fixtures drag, area rotor-body interference drag, area weight statement (component) fuselage group basic body body marinization pressurization body crashworthiness tail fold wing fold Aerodynamics Derived drag Weight Technology Factors + + + + + + + + + + + + + + + + + + + + + int AFuse real real real real real real Weight int real real real real real real WFuse real real real real real real aero weight _ _ cont fus body mar press crash ftfold fwfold _ _ _ _ _ _ _ _ cont fus fit rb _ _ _ _ Structure: Fuselage MODEL AFuse DoQ DoQV DoQ DoQV DoQ DoQ Weight MODEL dWbody dWmar dWpress dWcrash dWftfold dWfwfold WFuse TECH TECH TECH TECH TECH TECH xx=0.

_ TECH or xx=0 _ MODEL ; for fixed (input) weight use dWxx + model _ Wxx * xx _ TECH = Wxx weight model result multiplied by technology factor and increment added: Structure: Fuselage 2 2 2 2 0. 0. 0. 0. 0. 0.

10. 0.0 0.0 0.0 0.0 10. 0.0 0.0 0.0 0.0 Default ) ) M N SC SC = = M/q N/q ) ) M N ) SC Y SC = = SC ) = L M/q N/q SC Y /q = L/q ) ) (deg) ) ) M N L Y C C C C max (deg) β (per rad; based on wetted area, (per rad; based on wetted area and fuselage length, (deg) (per rad; based on wetted area and fuselage length, max zy /dβ /dα α β (per rad) /dβ (per rad) Y (per rad) (per rad; based on wetted area, ; 2 scaled, 0 M ; 2 scaled, 0 N ; 2 scaled, ) ; 2 scaled, (based on wetted area and fuselage length, ) (based on wetted area and fuselage length, dC 0 /dβ dC (deg) 0 /dα dC ) /dβ /dα ) L/q (per rad) M/q Y /q = N/q M N ) zl M/q = N/q L = ( C ( C α Y /q Y β ( dC /dα M/q N/q M α d C N β ( ( ) = d C d C L/q Lα ( d C specification (1 fixed, lift slope, lift slope, specification (1 fixed, moment at zero lift, moment at zero lift, moment slope, moment slope, specification (1 fixed, side force slope, side force slope, specification (1 fixed, moment at zero lift, moment at zero lift, moment slope, moment slope, zero lift angle of attack angle of attack for maximum lift lift pitch moment sideslip angle for zero side force sideslip angle for maximum side force side force yaw moment Aerodynamics, Standard Model Description + + + + + + + + + + + + + + + + + + + + + + + + + Type real real int real real int real real real real real real int real real int real real real real zl max lift moment side yaw _ _ _ _ _ _ zy max _ _ Chapter 41 Structure: AFuse Variable AoA AoA SET dLoQda dCLda SET MoQ0 CM0 dMoQda dCMda SS SS SET dYoQdb dCYdb SET NoQ0 CN0 dNoQdb dCNdb 2 2 2 2 2 2 0. 0. 0. 0. 2.

0.0 0.0 0.005 ) D C ) ) ) ) proj D D D (deg) D S SC SC = SC SC min = = D = = C d d D/q ); other parameter calculated ) ) ) ) ) X X | | D/q D/q D/q D D/q D D D D e e CX α α C C C C C | | Drb d d C K K D D C C ; 2 scaled, ; 2 scaled, ; 2 scaled, ; 2 scaled, ; 2 scaled, = = D D ) or scaled (use D/q D/q D/q D/q D/q C C Δ Δ , , XoQ d d (based on wetted area, (based on wetted area, (based on projected area, (based on wetted area, (based on wetted area, K X 0 fit wet D D Drb DV DS 0 fit rb V S C C C C C /S ) ) ) ) ) proj : fixed (use D/q D/q D/q D/q D/q S ( ( ( ( ( xxx _ DV specification (1 fixed, area coefficient specification (1 fixed, area coefficient specification (1 fixed, area coefficient total rotor-body interference drag, specification (1 fixed, area coefficient C specification (1 fixed, area coefficient model (0 none, 1 general, 2 quadratic) angle of attack for fuselage minimum drag drag increment drag increment SET forward flight drag fixtures and fittings rotor-body interference vertical drag sideward drag drag variation with angle of attack Drag, Standard Model + + + + + + + + + + + + + + + + + + + + + + + + + + int real real int real real int real real real int real real real int real real int real real real drag _ total _ fit rb(nrotormax) drag Dfit Drb Vdrag Sdrag Dmin _ _ _ _ fit _ rb(nrotormax) rb _ _ _ _ _ _ Structure: AFuse SET DoQ CD SET DoQ CD SET DoQ CD CD SET DoQV CDV CDVs SET DoQS CDS MODEL AoA Kdrag Xdrag 25.

(deg) t α (deg) (derived) t α (derived) d X model (1 input transition angle of attack, 2 calculate for quadratic) angle of attack for transition transition from forward flight drag to vertical drag angle of attack for transition exponent + + + int real real real trans _ tran _ Structure: AFuse MODEL AoA at Xd 1 0 0 8. 0.

0.6 0.0 0.0 0.0 0.0 0.0 200.

Default (knots) dive max V V .

= 1 1.0 for tandem, 0.3-0.6 for single main rotor (0.7-0.8 with ramp) ∼ = Vdive (fraction body weight) (fraction basic body weight) body (fraction basic body weight) cw / f ) press mar f f CG = 0.06 = 0.30 (AFDD84 or other) or 0.05 (AFDD82) for folding tail (fraction wing+tip (AFDD84 or other) or body+tailfold (AFDD82) weight) + Δ r (fraction tail (AFDD84 or other) or body (AFDD82) weight) = SLS max speed, crash tfold + wfold _ _ f c tfold f ( max V = fWbody fWbody crg _ AFDD84 (UNIV) is universal body weight model, for tiltrotor and tiltwing as well as for helicopters AFDD82 (HELO) is helicopter body weight model, should not be used for tiltrotor or tiltwing dive speed: fLength typically typically model (1 AFDD84, 2 AFDD82, 3 other) model (1 Boeing, GARTEUR (2 airplane, 3 helicopter), 4 Tishchenko, 5 Torenbeek, Raymer (6 transport, 7 gen av)) AFDD: rear cargo ramp (0 none) Boeing: cabin length + ramp length + cg range (fraction fuselage length) Boeing or Torenbeek or Raymer: design dive speed Raymer: number of cargo doors Raymer: cabin pressure differential (psi) body weight for marinization body weight for pressurization body weight for crashworthiness tail fold weight wing fold weight parameters Fuselage Group, NDARC Weight Model Custom Weight Model Description + + + + + + + + + + + + + + + Type int int int real real int real real real real real real real fuse(8) _ body other mar press crash tfold wfold crg _ _ _ _ _ _ _ _ ramp _ Chapter 42 Structure: WFuse Variable MODEL MODEL KIND fLength Vdive ndoor Pdelta fWbody fWbody fWbody fWbody fWbody WtParam 1 1 1 0.

Default ) gear _ d + gear _ WL – hub _ WL + ( HAGL LG d gear ) determines rotor height above ground level _ WL FltState ( (0 for retractable gear) D/q D/q HAGL : used for weight (fuselage and wing) height rotor = landing gear above ground + hub above landing gear = landing gear location: with place title notes landing gear location distance from bottom of landing gear to placement (1 located on body, 2 located on wing) retraction (0 fixed, 1 retracts) retraction speed (CAS or TAS, knots) model (0 none, 1 standard) standard model landing gear cruise drag, area landing gear helicopter drag, area Landing Gear Geometry Aerodynamics Derived drag Description + + + + + + + + + + + Type c*100 c*1000 Location real int int real int AGear real real aero _ LG LG LG _ _ _ gear _ gear _ Chapter 43 Structure: LandingGear Variable title notes loc d place KIND speed MODEL AGear DoQC DoQH 0. 0. 0.

1.0 1.0 1.0 xx=0.

_ TECH or xx=0 _ MODEL ; for fixed (input) weight use dWxx + model _ Wxx * cw LG xx _ χ LG χ ret TECH LG = χ basic landing gear retraction crashworthiness Wxx alighting gear group model (0 input, 1 NDARC, 2 custom) weight increment AFFD model weight model result multiplied by technology factor and increment added: weight statement (component) alighting gear group basic landing gear retraction crashworthiness Weight Technology Factors + + + + + + + + + + + Weight int real real real WGear real real real weight _ LG LGret LGcrash _ _ _ Structure: LandingGear Weight MODEL dWLG dWLGret dWLGcrash WGear TECH TECH TECH Default D/q drag area extended, Drag, Standard Model Description + + Type real Chapter 44 Structure: AGear Variable DoQ 2 3 0.

0.08 0.14 0.0325 Default (fraction basic+retraction weight) cw LG (fraction basic weight) = 0.14 f basic _ ret LG N LG crash _ f fWLG (fraction maximum takeoff weight) fWLG LG f = 0.0325 (fraction method) = 0.08, =fraction uses basic ret _ _ LG _ fWLG fWLG MODEL only typically typically model (1 fraction, 2 parametric rotary wing, 3 parametric fixed wing) number of landing gear assemblies basic landing gear weight landing gear weight for retraction landing gear weight for crashworthiness parameters Landing Gear Group, NDARC Weight Model Custom Weight Model Description + + + + + + + + Type int int real real real real gear(8) _ LG _ basic ret crash _ _ _ Chapter 45 Structure: WGear Variable MODEL nLG fWLG fWLG fWLG WtParam ’main’ Default ) ) multi isMainRotor , react , duct , var (keyword = twin _ ) int tilt _ , ) coax ’multirotor’ , twin , MODEL _ , tan not , twin _ Tdesign , tiltrotor , (keyword = ’reaction drive’ ) MODEL fThrust , , , ’prop’ prop , , coaxial geom rotor , _ _ tail ’tiltrotor’ ’tail’ , , , SET SET ’ducted fan’ , main tandem _ _ ’main’ ’tandem’ ’auxT’ , , ’antiQ’ ’coaxial’ ’variable diameter’ ROTORCONFIG ROTORCONFIG principal designation = antitorque = twin rotor = others = separately specify appropriate performance and weight models options defined by variables options defined by variables configuration designation: principal designation required, rest identify special characteristics principal designation determines where weight put in weight statement, and designates main rotors ( multiple rotor configurations have special options for geometry and performance antitorque or aux thrust rotor has special options for sizing reaction drive still requires propulsion group title notes configuration ( main rotor (0 not) antitorque rotor (0 not) auxiliary thrust rotor (0 not) variable diameter rotor (0 not) ducted fan (0 not) reaction drive (0 not) multiple rotors (0 not) configuration ( Rotor Configuration Description + + + + Type c*100 c*1000 c*32 int int int int int int int int int Chapter 46 Structure: Rotor Variable title notes config rotorconfig isMainRotor isAntiQRotor isAuxTRotor isVariableDiam isDuctedFan isReactionDrive isMultiRotor twinrotor 1 1 1 1 1 1.0 1.0 (control) included gear f ) 1 > n rotor _ for SET ) ( (depend on rotor radius) ref(1) ref(n) _ _ tip V gear , gear ratio factor Vtip Vtip nvelmax = var _ gear hover _ − tip and calculate ) V STATE ref used to change _ gear ref , 2 _ Vtip ) ref _ rVtip Vtip Radius 2 piecewise linear, maximum , then and rpm ratio) ≥ tip V and default tip speeds; (ratio rpm to rpm of primary rotor) ref(1) _ ref _ prim Vtip Ω / Vtip dep = Ω r can not specify gear ratio if sizing changes dependent rotor tip speed and gear ratio required for each drive system state primary: specify dependent: specify gear ratio, or specify if size task changes when evaluate rotational speed of dependent rotor ratio to state #1 reference rotational speed (state #1) gear ratio number of speeds (1 constant; speeds (CAS or TAS) cruise maneuvering flight drive system branch: only one primary rotor per propulsion group variable speed transmission: for drive system state reaction drive requires one and only one propulsion system (engine group or jet group) group number drive system branch (1 primary, 0 dependent) reference tip speed gear ratio input for dependent branch (1 effective radial station of force (fraction propulsion for reaction drive (group (1 engine, 2 jet), number, model) input form (1 tip speed, 2 hover function of flight speed tip speed rotor number Propulsion group Reaction drive Default rotor tip speeds (primary rotor) + + + + + + + + + + + + + + + + int int int real real real int real real int int int real real real gear Vtip ref _ _ _ xmsn ref(ngearmax) _ ref(ngearmax) react(3) cruise man _ _ _ _ _ react _ Structure: Rotor kRotor kPropulsion KIND Vtip rVtip Omega INPUT gear(ngearmax) r prop INPUT nVrpm Vrpm(nvelmax) Vtip Vtip 1. 1. 1. 1. 1. 1.

) xmsn _ ) ) rs _ DESIGN ’maxPQ’ fPlimit or GW=’maxQ’ FltState _ of SET ) ) or calculated (from Vtip _ rs _ SET for dependent branch Plimit quant=’Q margin’ _ gear(state) max = input (use ) ds _ req P ) ’input’ limit limit _ input ref = DS Ω rs ds (rotor P _ _ calculated from at × × tip rs rs limit rs V _ _ _ _ limit Plimit Size%SET limit : : from rotor power required at transmission sizing flight conditions ( ) RS SET RS P P Plimit fPlimit fPlimit : : : hover ’input’ − = tip ds=’input’ ds rs=0 rs=1 rs=2 _ _ _ _ _ /V tip limit limit limit limit limit V _ _ _ _ _ only for primary rotor; SET SET SET SET SET can be used for max effort in flight state ( can be used for max gross weight in flight condition or mission ( always check and print whether exceed torque limit OEI transmission sizing function of flight speed cruise maneuvering flight OEI transmission sizing function of flight speed default rotor tip speeds (including conversion): selectable by drive system torque limit: rotor shaft: options for rotor shaft power limit: corresponds to one rotor rpm ratio ( rotor shaft (0 input, 1 fraction power, 2 fraction drive system limit) rotor shaft power limit rotor shaft power limit factor rotor shaft torque limit ( Drive system torque limit + + + + + + + + + + + + + real real real real real real real real int real real real rs _ rs rs cruise man oei xmsn rs _ _ _ _ _ _ _ oei xmsn limit _ _ _ Structure: Rotor Vtip Vtip Vtip(nvelmax) fRPM fRPM fRPM fRPM fRPM(nvelmax) SET Plimit fPlimit Qlimit 0 0 1 1.0 1.0 ’std’ thrust _ SET Tdesign ) ) thrust _ , and thrust-weighted solidity ) to identify ref _ Tdesign * thrust _ Vtip DESIGN ) ) A fThrust A f DESIGN ( for dependent rotors ∑ diskload Tdesign / ref _ T = *DGW, 2 ref Vtip A , fDGW ref ), area = /A *DGW, takeoff condition, D (antitorque or aux thrust rotor); can specify using , and then W fDGW ref(1) A _ f Tdesign * from Vtip rotor=’DL+xx+xx’ /σ _ ref(1) W _ fThrust C SET Vtip , calculate , then , use and design gross weight obtained from thrust design conditions and missions ( ; aircraft disk loading = ’Vtip’ (for disk loading and blade loading) ’Vtip’ generally calculated (identified as input so inherited by next case) T /A (thrust-weighted) (thrust-weighted) Pdesign (main rotor) or fDGW W f Pdesign /σ D W W and specify not specify C Pdesign W N c/πR f = and = rotor rotor rotor=’CWs+xx+xx’ σ _ _ _ R W Tdesign otherwise use fraction rotor area for reference disk area fraction DGW thrust factor (antitorque or aux thrust rotor) tiltrotor (1 from clearance, 2 at wing tip, 3 at wing panel edge) ratio rotor radius to that of other rotor other rotor number wing number SET SET SET rotor disk loading = if rotor sized from disk loading ( if if if for antitorque or aux thrust rotor, need design conditions and missions ( Tdesign disk loading radius blade loading solidity thrust for antitorque or aux thrust rotor power for antitorque or aux thrust rotor rotor speed (rpm) at rotor thrust for disk loading and blade loading (0 default; 1 rotor thrust for disk loading and blade loading (1 from DGW, 2 from position (standard, tiltrotor, coaxial, tandem, tailrotor, multicopter) twin rotors Parameters Geometry + + + + + + + + + + + + + + + + + + + real real real real real real real real real real int int c*12 int real int int TRgeom thrust _ thrust geom _ _ _ Structure: Rotor diskload fArea fDGW fThrust Radius CWs sigma Tdesign Pdesign Ndesign SET iSET SET KIND fRadius otherRotor WingForRotor 1 1 0.

0.6 1.0 1.0 0.5 1.5 0.08 0.25 fus _ fclearance , ) fus , panel edge and wing span rotor BL _ _ loc rotor _ loc clearance multicopter , , fus PanelForRotor _ , tailrotor , Width , , wing span coaxial _ is relative calculated sep fus tandem WingForRotor d , tr d naccg ) is relative calculated _ ) is midpoint between hubs , first rotor is right loc coaxial fus WingForRotor ZoL ZoL , , , , WingForRotor d , first rotor is lower YoL YoL , pivot , _ tiltrotor XoL otherRotor loc XoL , , or otherRotor or for (fraction Diameter) s for (fraction Diameter) (fraction Diameter) : calculate lateral position (BL) pylon WL WL , _ : calculate position from , o m =clearance: from =wing tip: from =wing panel edge: from standard (fraction Diameter) _ BL loc BL , s , in SL coaxial SL ( ( _ geom _ (fraction Radius) TRgeom TRgeom TRgeom YoL _ _ _ WingForRotor sep rotor pylon (1 – separation/Diameter) or SET _ _ (clockwise from forward, deg) o : calculation override part of location input geom=’tiltrotor’ geom=’coaxial’ _ _ KIND KIND KIND same BL same loc loc ψ geom SET SET _ wing panel number tiltrotor clearance between rotor and fuselage tiltrotor clearance factor coaxial rotor separation tandem rotor overlap position ( clearance between rotor and fuselage horizontal separation vertical separation overlap overlap area fraction main rotor number radius scale factor clearance between tail rotor and main rotor angle arm length set diameter (1 conversion schedule, 2 function speed) ratio cruise radius to hover radius (variable diameter only) SET derived tail rotor multicopter variable diameter rotor + + + + + + + + + + + + + + + int real real real real int real real real real real int real real real real int real fus fus calc tr _ _ _ _ tr tandem twin _ _ _ geom twin twin VarDiam _ _ _ multicopter coaxial _ multicopter _ _ twin _ _ Structure: Rotor PanelForRotor clearance fclearance sep overlap iSET clearance Hsep Vsep overlap m mainRotor fRadius clearance ang len SET fRcruise 1 1 1 2 1. 1. 1.

-10.

multicopter _ len , ) ) fchord ) multicopter mainRotor rotor=’ratio’ _ hover , _ − in conversion mode tr ang _ V tandem SET conv _ DiskLoad V not rotor ≤ _ clearance overlap V to define control loc , 3 nonlinear from ) ); linear with )*function( g R otherRotor ; taper /σ twist t cruise σ Radius − rotor rotor _ _ conv V loc loc to calculate ≥ V ) ) is relative calculated , 2 linear from , first rotor is front *(main rotor Vdiam Radius(otherRotor) tr ZoL , , _ * , 2 nonlinear from nrmax Aircraft%config=’multicopter’ in hover and helicopter mode ( YoL , ) is center of rotors fdiam fTWsigma , twistL otherRotor fRadius fRadius YoL XoL , = = ref Radius for ) or only) is midpoint between hubs : calculate longitudinal and lateral position from in cruise mode ( /c = is relative calculated is relative calculated ) XoL nVdiam r /R R : calculate longitudinal position (SL) from is hover or reference radius; can be commanded by aircraft controls ( WL Radius : calculate longitudinal position (SL) from Radius or , also used for c XoL , XoL XoL , root tandem BL or BL r , (deg, root to tip) _ or or , fRcruise * L (deg) SL Radius (tip chord/root chord) SL SL identify as twin rotor ( θ ( ( identify as antitorque rotor ) t r N ( overlap Radius multicopter tw rotor pylon SL pylon SL rotor pylon config _ rotor=’ratio’ rotor=’scale’ θ _ _ _ _ _ config = _ _ geom=’tandem’ geom=’tailrotor’ geom=’multicopter’ _ _ _ same loc loc loc loc loc ang R SET SET ratio thrust-weighted solidity to geometric solidity taper ratio linear twist radial stations ( chord distribution twist SET SET SET if if conversion schedule: function of speed: use chord distribution (1 linear from twist distribution (1 linear from number of radial stations (maximum sizing: twin rotors: antitorque: variable diameter: direction of rotation (1 counter-clockwise, –1 clockwise) number of blades planform and twist Geometry + + + + + + + + + + + + + int int int real real int real int real real real chord twist _ _ Structure: Rotor rotate nBlade SET fTWsigma taper SET twistL nprop rprop(nrmax) fchord(nrmax) twist(nrmax) 0. 8. 0. 0.

5.7 0.1 0.0 1.04 0.97 ) taper (2 + 2 R / ) taper ) = (1 + 3 ) R .

( taper /c ) ) R f 2 (1 + 3 . / ( − c 1) (3 = − / f 1) − taper )*thrust f , thrust-weighted chord ; others calculated obtained from input .

75)4( .

to unit thrust-weighted chord R ) = 5 − fBlockage .

a (per-rad) r fchord(r) ; for linear taper (root) = (2 fchord α flapfreq c /c , or ; other calculated to tip) = geom _ = 1 + ( ) (per-rev at hover tip speed) a (tip) T taper c ν BR ref , B = f twist(r) sigma / (0. to use c/c or ν tw = _ : for SLS density, taper c (deg) f fTWsigma ) γ twistL δ sigma gamma /R = (lift zero from B determines whether Lock number input or calculated root determines how flap frequency and hub moment spring vary with rotor speed and r (deg) p hub β : use one of _ : use one of Iblade _ fTWsigma equivalent linear for linear taper chord twist for nonlinear distribution, scale input for nonlinear distribution, twist relative specified for hover radius and rotational speed KIND weight models can have separate blade and hub values for flap frequency SET _ _ hub type (1 articulated, 2 hingeless) first flapwise natural frequency coning natural frequency blade Lock number precone pitch-flap coupling blade section 2D lift-curve slope tip loss factor root cutout ( SET SET flap frequency and Lock number are used for flap dynamics and hub moments due to flap blade Lock number blockage factor: force acting on aircraft is (1– flap dynamics aerodynamics blockage factor (force increment + + + + + + + + + + + + int real real real real real real real real real hub _ Structure: Rotor KIND flapfreq conefreq gamma precone delta3 dclda tiploss xroot fBlockage 4 8 0.12 ’main’ (x)) ’prop’ (ry), ’tail’ ) (–z), ) ) mrmax mpsi mpsi ’main’ ) ; π/ π/ ’–z’ = 2 , = 2 ψ ψ mpsimax ’+z’ Δ Δ , ( ( mr / ) ’–y’ to 1; maximum , mpsi mpsi (scaled to unit thrust-weighted chord) to xroot to ’+y’ = hingeless) input) ref xroot = articulated) , − /c ) = 1 = 1 ) hub ’–x’ r R hub j blade _ j ( , _ I , , = (1 c ψ .

ψ hub conefreq geom 0 = ; mean geometric chord dr ’+x’ ( Δ _ Δ γ KIND K c ( KIND j j KE/P ( f e = cone = ) chord flap j 2 / j K K ψ ψ N c/πR (relative R , , πR ) ) ) = r j j ( ψ σ ψ ) tw 3 θ δ blade thickness-to-chord ratio number of radial stations ( number of azimuth angles (maximum radial increment cos( sin( direction of rotation (1 counter-clockwise, –1 clockwise) rotor area ( thrust-weighted chord solidity mean geometric chord aspect ratio, thrust-weighted blade area tan( chord distribution twist blade Lock number autorotation index blade moment of inertia flap stiffness flap hinge offset cone stiffness hub moment spring integration hub location pylon location pivot location nacelle cg location nominal orientation ( Geometry (for graphics) Geometry (derived) Blade element theory solution Geometry + + + + + + + + + + + + real real real real real real real real real real real real real real real real real real int int real real real Location Location Location Location c*16 calc _ geom geom _ _ rotor pylon pivot naccg calc _ _ _ _ _ Structure: Rotor thick frotate Arotor chord sigma chord AspectRatio Ablade KP fc(nrmax) tw(nrmax) gamma AI Iblade Kflap eflap Kcone Khub mr mpsi dr cspsi(mpsimax) snpsi(mpsimax) loc loc loc loc direction 0 1 0. 0. 0. 0. 1.

=fixed shaft tilt _ KIND pivot _ loc rotor relative w/N ref = 90 or _ naccg _ incid loc ) ES (drive system and engine system) total/nRotorOnWing _ ) W ES Wtip and W ( = ∑ gbrs + W Wmove , 3 , gbrs W gbrs (zero shaft control) W Wmove SF P F F P * C C C (deg) p (deg) φ (deg) p p ψ fWmove θ , orientation of pivot axes, and reference shaft control angles not used for (deg) (deg) (deg) (drive system) or 123) (deg) h ref pivot h ± ref for reference control) φ c _ θ i : cg shift calculated using incidence and cant rotation of gbrs 123) loc SF W , ± C ( = Wmove moving weight w _ naccg HF _ incidence cant angle pivot dihedral angle pivot pitch angle pivot sweep angle incidence cant angle weight (1 wing tip weight, 2 fraction moving weight loc for tiltrotor, locations and orientation specified in helicopter mode, so SET shaft control (0 fixed shaft, 1 incidence, 2 cant, 3 both controls) orientation of rotor shaft orientation of pivot axes reference shaft control moving weight for cg shift nominal orientation (1, –1, 2, –2, 3, –3, –3, r2, 1) axis incidence ( axis cant ( incidence (0 fixed, 1 controlled) cant angle (0 fixed, 1 controlled) pivot axes relative airframe, pivot axes relative airframe, W C rotor shaft relative airframe, Derived geometry + + + + + + + + + + + + + + int real real real real real real real int real int int int int int real real real real pivot _ pivot tilt incid cant pivot _ _ hub ref _ _ hub ref Wmove _ incid cant _ _ _ _ _ _ _ Structure: Rotor KIND incid cant dihedral pitch sweep incid cant SET fWmove iDirection axis axis KIND KIND CPF(3,3) CFP(3,3) WCHF(3,3) CSF(3,3) 1 1 2 1 1 0 1 0 1 0 1 0 ) ) ) ) nvelmax nvelmax nvelmax nvelmax matrix) matrix) matrix) matrix) T T T T 2 piecewise linear, maximum 2 piecewise linear, maximum 2 piecewise linear, maximum 2 piecewise linear, maximum ≥ ≥ ≥ ≥ ) /σ T C matrix (0 for none) T (nacelle tilt) i cyclic input (1 tip-path-plane tilt, 2 hub moment, 3 lift offset) collective input (1 thrust, 2 scale collective connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) rotor control mode (1 thrust and TPP, 2 thrust and NFP, 3 pitch and TPP, 4 pitch and NFP) collective (magnitude of thrust vector) longitudinal cyclic (tip-path plane tilt or no-feathering plane tilt) lateral cyclic (tip-path plane tilt or no-feathering plane tilt) incidence Controls + + + + + + + + + + + + + + + + + + + + + + + + + + + + + int int int int int real int real real int real int real real int real int real real int real int real real coll coll lngcyc latcyc incid _ _ _ _ _ control cyclic coll _ _ _ coll(ncontmax,nstatemax) lngcyc(ncontmax,nstatemax) latcyc(ncontmax,nstatemax) incid(ncontmax,nstatemax) _ _ _ _ Structure: Rotor KIND KIND KIND SCALE INPUT T nVcoll coll(nvelmax) Vcoll(nvelmax) INPUT T nVlngcyc lngcyc(nvelmax) Vlngcyc(nvelmax) INPUT T nVlatcyc latcyc(nvelmax) Vlatcyc(nvelmax) INPUT T nVincid incid(nvelmax) Vincid(nvelmax) 1 0 1 0 1 0 ) ) ) ) c nvelmax nvelmax nvelmax 0 nvelmax c + AC T c = c matrix) matrix) matrix) (for each control state) and value T T T 2 piecewise linear, maximum 2 piecewise linear, maximum 2 piecewise linear, maximum 2 piecewise linear, maximum T ≥ ≥ ≥ ≥ (variable speed transmission only) gear f (variable diameter only) diam flight state specifies control state, or that control state obtained from conversion schedule f can be zero, constant, or function of flight speed (CAS or TAS, piecewise linear input) for each component control, define matrix c by connecting aircraft control to component control, flight state can specify component control value initial values if control is connected to trim variable; otherwise fixed for flight state c connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) number of speeds (0 zero value; 1 constant; aircraft controls connected to individual controls of component, pylon moves with rotor; nontilting part is engine nacelle cant diameter gear ratio factor rotor lift Trim Targets + + + + + + + + + + + + + + + + + + + + + int real int real real int real int real real int real int real real int cant diam fgear _ _ _ cant(ncontmax,nstatemax) diam(ncontmax,nstatemax) fgear(ncontmax,nstatemax) _ _ _ Structure: Rotor INPUT T nVcant cant(nvelmax) Vcant(nvelmax) INPUT T nVdiam fdiam(nvelmax) Vdiam(nvelmax) INPUT T nVfgear fgear(nvelmax) Vfgear(nvelmax) nVlift 16 16 ) ) nvelmax X/q − , or /σ X C − , X − /σ T C 2 piecewise linear, maximum , or ≥ quant /σ _ L C Aircraft%trim ) μ vs /σ used to calculate rotor thrust margin, which available for max effort or trim T C = .170,.168,.161,.149,.131,.109,.084,.050,.049,.048,.047,.046,.045,.044,.043,.042 = 0.,.10,.20,.30,.40,.50,.60,.70,.71,.72,.73,.74,.75,.76,.77,.78 tran = .200,.197,.190,.177,.156,.135,.110,.080,.075,.070,.065,.060,.055,.050,.045,.040 _ = 0.,.10,.20,.30,.40,.50,.60,.70,.72,.74,.76,.78,.80,.82,.84,.86 CTs(1)=0.

steady tran CTS _ _ steady tran , _ _ CTs CTs mu mu can be fraction total aircraft drag, propulsive force /σ /σ steady can be fraction total aircraft lift, lift, T T _ target speeds (CAS or TAS) number of speeds (0 zero value; 1 constant; target speeds (CAS or TAS) number of points (maximum 20) advance ratio C number of points (maximum 20) advance ratio C target definition determined by Klift Kprop CTs defaults used if default default default default rotor propulsive force sustained transient Rotor Thrust Capability ( + + + + + + + + + + + + + + + real real int real real int real real int real real steady(20) tran(20) steady(20) _ tran(20) _ _ _ Structure: Rotor Klift(nvelmax) Vlift(nvelmax) nVprop Kprop(nvelmax) Vprop(nvelmax) nsteady mu CTs ntran mu CTs 1 2 2 1 1 2 2 0. 0.

0.0 0.0 1.0 0.0 1.2 high _ Vint ≥ V ) ES to 0 at W low _ and Vint gbrs ≤ W for no interference at all V at ) int=0 _ Kint duct ; 3 scaled, _ ) gbrs MODEL W fLength ref Dscale θ _ − Y Units = ( ; 2 scaled, / BS ) C Swet , ; 2 scaled, from ref rotor θ duct _ w/N S ( / pylon S = k pylon duct to suppress interference at component; S S standard model, induced power standard model, profile power table model low velocity (knots) high velocity (knots) pitch relative shaft axes area factor, area duct length (fraction rotor radius) if thrust and TPP command, and neglect inplane forces relative TPP, then pitch control angles not required Kint=0 with transition: interference factors linearly vary from power model (1 standard, 2 table model) inplane forces, tip-path plane axes (1 neglect, 2 blade-element theory) inplane forces, profile (1 simplified, 2 blade element theory) model (0 none, 1 standard, 2 with transition) transition standard model hub/pylon aerodynamic axes (0 input pitch, 1 helicopter, 2 propeller or tiltrotor) pylon wetted area (1 fixed, input duct area (1 fixed, input Performance Interference Geometry + + + + + + + + + + + + + + + + + + int PRotorInd PRotorPro PRotorTab int int int real real IRotor int real int real real int real real perf Ftpp Fpro int duct _ _ _ _ _ pylon aero pylon _ low high aeroaxes Spylon Sduct _ _ _ _ _ _ _ duct _ Structure: Rotor MODEL PRotorInd PRotorPro PRotorTab MODEL MODEL MODEL Vint Vint IRotor SET pitch SET Swet kSwet SET S fLength 2 1 0. 0.

0.0 1.0 ) ) and weight pylon _ duct S ) for propeller kSwet spin ) _ duct _ fSwet ); spinner radius used for drag and weight R * fLength ) spin for helicopter, _ fSwet ) or calculated (from (drive system and engine system) ; used for drag (wetted area fRadius ES pylon R aero=180 _ W * ) or calculated (from _ ∑ 3 spin Swet duct / _ pitch _ + ; ) or calculated (from ; 2 scaled, from /kg Swet gbrs aero duct _ _ W fLength (zero shaft control) S Swet = pylon or m ) pitch 3 hub BF / ) 2 pylon hub C duct ) ) D/q hub BS ( /lb , ) D/q C ( D/q spin D/q ( duct = spinner frontal area (from ( D/q ( /A are ft ) =input uses 2 spin (drive system) or spin πR spin S πR R kSwet gbrs , pylon wetted area: input (use = = (2 = , duct area: input (use , spinner wetted area: (use aeroaxes W k _ spin S = spin duct Spylon Sduct Sspin SET units of w pylon wetted area used for pylon drag rotor pylon must be consistent with engine group nacelle S A _ _ _ area factor, only SET SET SET spinner wetted area (1 fixed, input spinner radius (fraction rotor radius) pylon axes relative shaft, pylon axes relative airframe, spinner radius model (0 none, 1 standard) incidence angle for helicopter nominal drag (deg; 0 for not tilt) standard model hub cruise drag, area hub helicopter drag, area hub vertical drag, area pylon cruise drag, area pylon helicopter drag, area Derived geometry Drag Derived drag + + + + + + + int real real real real real real int real DRotor real real real real real drag spin _ spin hub pylon _ hub hub pylon spin _ _ _ _ _ _ spin Sspin _ _ _ Structure: Rotor SET Swet fSwet fRadius CBS(3,3) CBF(3,3) Radius MODEL Idrag DRotor DoQC DoQH DoQV DoQC DoQH 1 1 0. 0. 0. 0. 0. 0. 0. 0. 0. 0.

3.0 0.9 1.0 0.6 1.0 1.0 1.0 1.0 ) tip _ Wblade (sec) /P Ω blade I blade N 1 2 duct ) pylon = ) duct ) duct ) D/q D/q ( ( spin D/q spin ) hub ( D/q KE/P χ χ ( D/q wet ( shaft S χ blade χ blade hub and hinge inter-rotor shaft fairing/spinner blade fold tail rotor auxiliary thrust rotor support structure duct location tip weight (fraction blade radius) distributed weight for centrifugal force (fraction model (0 input, 1 NDARC, 2 custom) weight increment NDARC model autorotation index tip weight (per blade) radius of gyration for distributed mass (fraction blade radius) pylon vertical drag, area duct cruise drag, area duct helicopter drag, area duct vertical drag, area spinner drag, area total wetted area weight statement (component) rotor group (or empennage or propulsion group) blade moment of inertia (0 from Lock number, 1 from blade wt, 2 tip wt from Lock number, 3 tip wt from AI) blade weight (all blades; required for drive system weight) weight on wing tip (required for tiltrotor wing weight) blade weight hub and hinge weight inter-rotor shaft fairing/spinner weight Weight Technology Factors + + + + + + + + + + + + + + + + + + + + + + + + real real real real real real Weight int real real real real real real real real real WRotor int real real real real real real real real real real real weight tip tip tip _ _ _ duct _ pylon duct duct blade hub shaft spin _ _ _ _ _ _ _ _ rotor spin Iblade _ _ _ Structure: Rotor DoQV DoQC DoQH DoQV DoQ Swet Weight MODEL dWblade dWhub dWshaft dWspin dWrfold dWtr dWaux dWrsupt dWduct WRotor SET AI Wblade rWblade fWblade rblade Wblade Wtip TECH TECH TECH TECH 1.0 1.0 1.0 1.0 1.0 = 0 Iblade _ SET xx=0.

_ TECH or not used and tip xx=0 _ _ Iblade Iblade Wblade MODEL blade N tip = 2 or 3), W ) f calculated from calculated from config _ tip tip _ _ + (1 + Iblade MODEL Wblade Wblade blade ; for fixed (input) weight use dW + dWxx + blade , independent of blade weight , tip weight , tip weight w AI supt χ model _ blade gamma gamma χ at Wxx = χ * xx fold _ tr blade χ χ W duct TECH : calculate blade moment of inertia χ = Iblade Wxx 0 from Lock number 1 from blade weight 2 from Lock number 3 from autorotation index main rotor: rotor group tail rotor: empennage group (tail rotor) propeller: propulsion group (propeller/fan installation) _ weight model result multiplied by technology factor and increment added: blade weight: SET for tail rotor or auxiliary thrust weight model ( rotor weight = blade + hub + spinner + fold + shaft + support + duct rotor config determines where weight put in weight statement blade fold weight tail rotor weight auxiliary thrust weight rotor support structure weight duct weight + + + + + real real real real real rfold tr aux rsupt duct _ _ _ _ _ Structure: Rotor TECH TECH TECH TECH TECH 0. 0. 2. 0. 0. 2. 0. 2. 0. 0. 0.

2.0 2.0 1.0 4.5 1.12 1.08 0.08 0.08 0.65 1.176 Default edge κ prop _ (linear) (quadratic) prop Ki ) ) ) κ z z z ) for μ μ μ z p h h p p ( ( ( p h μ κ κ κ κ κ p κ κ κ variation κ ( κ variation κ κ h p prop for z for for for for for for for κ κ for for μ for for 1 2 1 2 1 2 3 climb h h p p hα in transition from hover to climb a a a in transition from hover to climb κ h hα a for for k k k k p k k k k X k X X ind ind hover axial axial ) ) κ X M /σ /σ T T C C hover axial climb axial cruise (propeller) edgewise flight (helicopter) ( coefficient coefficient exponent ( coefficient coefficient exponent coefficient exponent advance ratio coefficient coefficient coefficient exponent model (1 constant, 2 standard) induced velocity factors (ratio to momentum theory induced velocity) variation with thrust variation with shaft angle constant exponent variation with axial velocity Rotor Induced Power, Standard Energy Performance Method Description + + + + + + + + + + + + + + + + + + + + + + + + + + + Type int real real real real real real real real real real real real real real real real real real real real real ind _ Hind Pind _ _ prop hover climb prop edge _ _ _ _ _ Chapter 47 Structure: PRotorInd Variable MODEL Ki Ki Ki Ki CTs kh1 kh2 Xh2 CTs kp1 kp2 Xp2 kpa Xpa Maxial Xaxial mu ka1 ka2 ka3 Xa 3 1 0. 1. 0. 0. 8. 1. 1. 1. 1. 1.

0.8 4.5 10.

0.35 edge _ Ki , prop _ Ki , ; smaller effective height accounting for increased influence of ground compared hover .

m _ f Ki supersede calculated value x = 0 g f g edge C _ C (linear) (quadratic) Ki y ) ) ) eα f μ μ μ ) k FltState for ( ( ( off μ f κ κ κ ( in S κ Ki edge for off for for for μ f for 1 S 1 2 3 e e e o e k k k k for X =constant uses only max o min κ k ind κ _ : for tiltrotors, typically advance ratio coefficient coefficient coefficient exponent coefficient factor effective height correction longitudinal gradient factor lateral gradient factor MODEL nonzero values of Cge to isolated rotor variation with edgewise velocity variation with rotor drag variation with lift offset minimum maximum edgewise scale factor axial scale factor ground effect (0 none, 1 Cheeseman and Bennett, 2 BE Cheeseman and Bennett, 3 Law, 4 Hayden, 5 Zbrozek) inflow gradient in forward flight (0 none, 1 White and Blake, 2 Coleman and Feingold) hub moment inflow gradient factor Momentum theory + + + + + + + + + + + + + + + + + + + real real real real real real real real real real real real int real int real real real GE grad _ _ edge min max _ _ _ Structure: PRotorInd mu ke1 ke2 ke3 Xe kea ko1 ko2 Ki Ki fedge fprop MODEL Cge MODEL fGradx fGrady fGradm 1. 1. 1.

0.5 1.0 1.0 1.0 1.00 0.85 1.05 ’config’ ) D twin=0.90 _ Kh ’multirotor’ to 0.12 , or D ’tandem’ , ) not used) twin ’coaxial’ _ , multirotor Kh , ; ’tiltrotor’ tandem coaxial ¯ , α _ or A for hover for forward flight (use coaxial , with zero horizontal separation (typically = 0.88 to 0.81 for rotor separation 0.06 twin twin κ κ config=’duct’ ’side-by-side’ twin , _ sidebyside Kf for hover for forward flgiht , h f ’none’ twin=’coaxial’ x x , _ twin=’tandem’ none _ _ (fan edgewise velocity/free stream velocity) (fan axial velocity/free stream velocity) twin , forward flight, this rotor , forward flight, other rotor ’config’ _ , hover 1 2 V x V z MODEL (rotor thrust/total thrust) : h f f f f (fan area/far wake area) MODEL T x x x f in hover to forward flight transition A twin or f _ must identify rotor as twin or multiple rotors C MODEL coaxial: coaxial and tandem: ducted fan model used only if MODEL ’config’ model (1 specify area ratio, 2 specify thrust ratio) area ratio thrust ratio velocity ratio velocity ratio model (based on config, none, side-by-side, coaxial, tandem, multirotor) ideal induced velocity correction ideal induced velocity correction constant coaxial rotor nonuniform disk loading factor multirotor thrust factor multirotor thrust factor model ( thrust factor thrust factor thrust factor Ducted fan Twin rotors Derived twin rotors + + + + + + + + + + + + + + int real real real real c*12 real real real real real real int real real real twin duct twin _ _ _ twin _ twin twin multi(nrotormax) _ multi(nrotormax) coaxial _ _ _ _ Structure: PRotorInd MODEL fDuctA fDuctT fDuctVx fDuctVz MODEL Kh Kf Cind A xh xf iMODEL xh xf1 xf2 0. 2. 2.

1.0 0.0 0.2 0.0 0.5 0.0 0.5 4.0 3.0 0.0 0.0 0.07 0.07 0.009 0.009 Default (axial) sep ) d | ) (hover/edgewise) c sep min sep X D d + Δ ) ) c sep /σ Δ ) T + Δ /σ C 2 ( T 2prop Δ C d − ( 2hel /σ d | − Δ + T 1 2 C dp /σ | dp f f c f c Δ + T d d 1prop C X | d for for ( 1hel Δ = Re + d pα pα sep X d + d X = 0prop 0hel d d sep ) = d = c /σ dp (0. for no correction) Δ T c dh c C ref Re in drag, in drag (axial) in drag (axial) in drag, in drag (hover/edgewise) in drag (hover/edgewise) in drag increment for minimum profile drag ( 0hel 1hel 2hel 0prop 1prop 2prop (1 array, 2 equation) sep for separation ( in drag increment d d d d d d χ min X D sep sep basic ) ) d d vs thrust-weighted c /σ /σ d T T c C C number of points (maximum 24) blade loading drag coefficient ( coefficient coefficient coefficient coefficient coefficient coefficient variation with shaft angle, coefficient variation with shaft angle, exponent ( factor exponent variation with edgewise velocity, coefficient variation with edgewise velocity, coefficient variation with edgewise velocity, exponent profile power exponent for Reynolds number correction array ( equation Rotor Profile Power, Standard Energy Performance Method Technology factor Reference Reynolds number Basic model Description + + + + + + + + + + + + + + + + + + + + + + + + + + Type real real real int int real real real real real real real real real real real real real real real real real basic _ drag _ cd(24) Dmin sep _ _ _ ref hel hel hel prop prop prop _ Re _ _ _ _ _ _ _ Chapter 48 Structure: PRotorPro Variable TECH Re X MODEL ncd CTs cd(24) CTs d0 d1 d2 d0 d1 d2 dprop Xprop CTs dsep Xsep df1 df2 Xf 1 1 2. 0. 8. 0.

1.0 40. 2.0 3.0 1.0 0.08 ) s X s Δ s d + s X s Δ s d = d c Δ , s ) /σ T C )( off f τ α /f supersede calculated cdmean s = 0.0 to 0.23 (uniform increments) f ( /σ T | − C (0 none, 1 drag divergence, 2 similarity) FltState sα /σ d ): off in T f tip C | stall(1)=0. comp _ d cdo for off c = = 0.17,0.16,0.15,0.14,0.13,0.12,0.11,0.10,0.10,0.10 V /V f = 0.00,0.05,0.10,0.15,0.20,0.25,0.30,0.35,0.40,0.80 s o CTs cd(1)=0.

d Δ .01048,.01152,.01336,.01593,.01920,.02381,.03014,.04000,.08000,.16000,.32000,1.0000 (0 none) in stall drag increment in stall drag increment for stall _ stall 1 2 2 s _ in stall drag increment s s o ) = .01100,.01075,.01025,.01000,.01010,.01070,.01050,.00975,.00925,.00926,.00938,.00977, d f in stall drag increment in stall drag increment stall s X X d CTs mu f 1 2 /σ c s s cd at stall ( T d d C number of points (maximum 20) advance ratio ( coefficient factor factor blade tip thickness-to-chord ratio /σ T default array ( nonzero values of default used if default default C constant factor factor exponent exponent variation with lift offset variation with rotor drag similarity model Stall model Compressibility model + + + + + + + + + + + + + + + + + + int int real real real real real real real real real real int real real stall comp _ _ tip stall(20) _ stall(20) _ _ Structure: PRotorPro MODEL nstall mu CTs fstall dstall1 dstall2 Xstall1 Xstall2 do1 do2 dsa MODEL fSim thick 2.0 0.88 0.16 0.056 0.416 ) m X m Δ m d + ) c m Δ m d Mddcl − = d c Δ Mdd0 , dd = M dd /∂c − M dd at ∂M M = = κ m Δ in drag increment in drag increment 1 2 in drag increment m m m d d at zero lift X dd M derivative with lift coefficient coefficient exponent drag divergence Mach number ( drag divergence model ( + + + + + + + real real real real real Structure: PRotorPro dm1 dm2 Xm Mdd0 Mddcl 1 0 1 1 0 0 0 0 0 0 0 ’ ’ ’ ’ Default ) /σ P o C = 8 F ) mean d ) c ) ) ntablemax , 2 table ntablemax ) ntablemax ntablemax ) mean d c ntablemax ntablemax T P P at zT P P values (maximum α μ M values (maximum or values (maximum values (maximum or /T R T P P x HP μ zT P P κ α T P P zHP d T P P M μ c or μ μ α values (maximum or /σ or or values (maximum T HP /σ C μ at /T R T HP zHP HP x μ μ α C M M number independent variables (1 to 3) variables profile power model (0 standard, 1 table number independent variables (1 to 3) variables number of number of number of number of number of lift offset number of advance ratio axial advance ratio shaft angle-of-attack blade loading lift offset advancing tip Mach number induced power factor profile power mean induced power model (0 standard, 1 table, 2 table with equations) profile power model (0 standard, 1 table, 2 table with equations) independent variables table Performance, Table Method Description + + + + + + + + + + + + + + + + + + + + + + + + Type int int c*12 int int int c*12 int int int int int int real real real real real real real real indTab proTab _ _ proTab ind _ pro _ _ ind(3) pro(3) _ _ Chapter 49 Structure: PRotorTab Variable MODEL nVar var MODEL KIND nVar var nmu nmuz nalpha nCTs nMx nMat mu(ntablemax) muz(ntablemax) alpha(ntablemax) CTs(ntablemax) Mx(ntablemax) Mat(ntablemax) Ki(ntablemax,ntablemax,ntablemax) cdo(ntablemax,ntablemax,ntablemax) ) ) Mat , Mat , Mx , Mx , CTs , CTs , alphaTPP , alphaTPP , (hub plane) ) /μ z muzTPP (tip-path plane) μ , muzTPP ) ( , − /μ z μ ( muTPP , muTPP = tan − , (hub plane) α (hub plane) supersede table (or table with equations) values μ at alpha z , alpha (tip-path plane) = tan μ M , pro μ (tip-path plane) α _ muz z /σ , muz μ var T , FltState C mu in _ /T R mu and x _ M cdo ind _ tablevar var tablevar : shaft angle-of-attack and/or : axial velocity ratio Ki : edgewise advance ratio : blade loading : lift offset : shaft angle-of-attack : axial velocity ratio : edgewise advance ratio ’alphaHP’ ’muzHP’ , ’CT/s’ ’muHP’ ’offset’ , , : advancing tip Mach number , , ’mu’ ’muz’ ’alpha’ ’muTPP’ ’muzTPP’ ’alphaTPP’ ’CTs’ ’Mx’ ’Mat’ independent variables: arrays of independent variables used for both induced power and profile power nonzero values of induced power variables ( number of variables variable values profile power variables ( number of variables variable values Derived table variables int int real int int real ind(3) pro(3) _ _ Structure: PRotorTab ivar nvi(3) vi(ntablemax,3) ivar nvo(3) vo(ntablemax,3) 2 2 2 1 2 2 2 2 2 0.8 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0024 Default ) ) ) D ) D D ) D SC D SC SC ) SC = SC = = ) = ) D = D D/q Dscale ) _ D/q D/q SC D/q ) D ) SC ) D/q ) = D C D ; 3 scaled, squared-cubed; 4 scaled, square-root) ) = D D Units C C D D ( C C C C D/q / D/q /W ; 2 scaled, ) ; 2 scaled, ; 2 scaled, ; 2 scaled, ; 2 scaled, ; 2 scaled, ; 2 scaled, D/q D/q ( D/q D/q D/q D/q D/q D/q or / (based on pylon wetted area, (based on duct wetted area, (based on pylon wetted area, (based on rotor area, (based on duct wetted area, (based on spinner wetted area, (based on rotor area, 1000) hub pylon duct hub pylon duct spin W/ D D D D hub DV pylon DV duct DV ( hub pylon duct spin V V V C / C C C C C C ) ) ) ) ) ) ) ) D/q D/q D/q D/q D/q D/q D/q D/q ( ( ( ( ( ( ( = ( area coefficient k area coefficient area coefficient area coefficient area coefficient area coefficient area coefficient hub drag specification (1 fixed, pylon drag specification (1 fixed, duct drag specification (1 fixed, spinner drag specification (1 fixed, hub drag specification (1 fixed, pylon drag specification (1 fixed, duct drag specification (1 fixed, hub drag model (1 general, 2 quadratic) pylon drag model (1 general, 2 quadratic) forward flight drag vertical drag transition from forward flight drag to vertical drag Rotor Drag, Standard Model Description + + + + + + + + + + + + + + + + + + + + + + + + + + + + Type int real real real int real real int real real int real real int real real int real real int real real int int Dhub Dpylon _ _ hub hub pylon duct _ _ _ _ hub pylon duct hub pylon duct spin Dhub Dpylon Dduct Dspin Vhub Vpylon Vduct _ _ _ _ _ _ _ _ hub _ pylon _ duct _ spin _ _ _ _ _ _ _ Chapter 50 Structure: DRotor Variable SET DoQ CD kDrag SET DoQ CD SET DoQ CD SET DoQ CD SET DoQV CDV SET DoQV CDV SET DoQV CDV MODEL MODEL 2. 2. 2.

/ /kg hub _ or m (antitorque or aux thrust rotor) / kDrag , /lb hub _ ; ft Tdesign * / ); other parameter calculated CD , CD /Mg hub fThrust _ d d d X (derived) or m DoQ X (derived) X d (derived) d / d X X X /klb ) or scaled (use (main rotor) or DoQ are ft M T O W kDrag W (2/3 power) = 1.4 for typical hubs, 0.8 for current low drag hubs, 0.5 for faired hubs (English units) (1/2 power) = 0.074 for single rotor helicopters, 0.049 for tandem helicopters, f , hub drag: use one of 0.038 for hingeless rotors, 0.027 for faired hubs (English units) : fixed (use = 0.0040 for typical hubs, 0.0024 for current low drag hubs, 0.0015 for faired hubs = xxx Dhub tiltrotor with tilting engines use pylon drag (and no nacelle drag), since pylon connected to rotor shaft axes tiltrotor with nontilting engines: use nacelle drag as well rotor with a spinner (such as on a tiltrotor aircraft) likely not have hub drag units of CD kDrag kDrag W _ _ duct drag model (1 general, 2 quadratic) hub drag, transition exponent pylon drag, transition exponent duct drag, transition exponent SET component drag contributions must be consistent; pylon is rotor support, and nacelle is engine support SET hub drag, transition exponent pylon drag, transition exponent duct drag, transition exponent + + + + int real real real real real real Dduct _ hub pylon duct _ _ _ Structure: DRotor MODEL X X X Xh Xp Xd 3 3 1 1 6 2 0.2 0.2 1.0 1.0 1.0 1.4142 Default ∞ √ Xboundary= high _ ) ) Vint Ktwin= ≥ V Xdevelop=1.

points along span Xdevelop ) Xboundary Nint to 0 at for no interference at all Ktwin low _ int=0 Vint _ ; always immersed same as ≤ ; nominal same as V at MODEL Kint Xboundary=0 T Xdevelop=0 K (fraction contracted radius) s (0. for no interference) t int K : step function same as : step function same as : only for coaxial or tandem or side-by-side; nominal same as develop boundary twin _ _ _ to suppress interference at component; rate parameter far wake contraction (0 no, 1 yes) boundary transition velocity factor in overlap region development along wake axis (1 step function, 2 nominal, 3 input immersion in wake (1 step function, 2 always immersed, 3 input twin rotor interference (1 no correction, 2 nominal, 3 input number wing span stations number tail span stations at fuselage at wing at tail Kint=0 interference factor linearly transition from to account for wing or tail area in wake, interference averaged at MODEL MODEL MODEL model interference factors Rotor Interference, Standard Model Description + + + + + + + + + + + + + + + Type int real int int real int real int int real real real twin _ develop boundary contract int _ _ _ _ wing(nwingmax) tail(ntailmax) fus wing(nwingmax) tail(ntailmax) _ _ _ _ _ Chapter 51 Structure: IRotor Variable MODEL Xdevelop MODEL MODEL Xboundary MODEL Ktwin Nint Nint Kint Kint Kint 0.

− ∼ = int C (0. for no interference) int C For tiltrotors, typically the interference is wing-like, with kind (1 wing-like, 2 propeller-like) factor Induced power interference at wing + + + int real wing _ int _ wing(nwingmax) _ Structure: IRotor KIND Cint 1 1 1 0 1 2 1 1 1 0. 0. 0. 0. 1.

0.3 0.21 0.05 0.15 0.11 Default (at 70%R) (per-rev at hover tip speed) R ν .

τ (at 25%R) R .

(at 20%R) τ R .

) at τ at (ft or m) P T s /T R ) m f or fixed ) blade _ s/R (roll moment L flapfreq flapfreq hub weight equation depend on blade weight (for hub weight; 0 no, 1 yes) rotor kind (for blade weight; 1 tilting, 2 not) blade (0. to use hub (0. to use rotor tip clearance (for blade weight; 1 scaled, 2 fixed) design lift offset blade airfoil thickness-to-chord ratio tip clearance, scaled GARTEUR: blade airfoil thickness-to-chord ratio AFDD82: design maximum thrust AFDD10: design maximum power AFDD10: material factor model (1 rotor, 2 tail rotor, 3 auxiliary thrust) blade weight model (1 AFDD82, 2 AFDD00, 3 lift offset, 4 Boeing, 5 GARTEUR, 6 Tishchenko) hub and hinge weight model (1 AFDD82, 2 AFDD00, 3 lift offset, 4 Boeing, 5 GARTEUR, 6 Tishchenko) inter-rotor shaft weight (from lift offset; 0 not included) AFDD00 weight models AFDD00 and AFDD82: first flapwise natural frequency lift offset rotor Boeing: blade airfoil thickness-to-chord ratio tail rotor weight model (1 AFDD, 2 Boeing, 3 GARTEUR) auxiliary thrust weight model (1 AFDD10, 2 AFDD82, 3 Boeing, 4 GARTEUR, 5 Torenbeek) Rotor Group, NDARC Weight Model Description + + + + + + + + + + + + + + + + + + + + + + + Type int int int int int int real real int real real real real int real int real real real tip config Wblade Whub Wshaft type offset tr aux _ aux blade hub _ _ _ _ _ _ _ _ _ _ _ aux aux rotor _ _ _ Chapter 52 Structure: WRotor Variable MODEL MODEL MODEL MODEL MODEL KIND flapfreq flapfreq MODEL offset thick20 clearance thick25 MODEL thick70 MODEL thrust power material 0. 0. 0.

1.5 0.09 = 0) results in a lower average type _ MODEL should be the coning frequency blade _ flapfreq = 1) is best for advanced technology rotors with blades lighter than type _ ) MODEL or kg/m (lb/ft (fraction total blade weight) duct U fold f = 0.04 for manual fold, 0.28 for automatic fold , use design maximum thrust of rotor from sizing task , use design maximum power of rotor from sizing task : tail rotor and auxiliary thrust models use only rotor, support, and duct weights (not shaft, fold, or 1 for composite construction, 1.20 for wood, 1.31 for aluminum spar, 1.44 for aluminum construction is the reference rotor speed prop fWfold config aux=0 aux=0 aux= _ _ _ _ Ω thrust power MODEL separate blade and hub weights) duct weight only used for ducted fan configuration for teetering and gimballed rotors, the flap frequency The AFDD00 hub weight equation using the calculated blade weight ( error, and best represents legacy rotor systems. Using the actual actual blade weight ( trend. if if material default typically rotor support structure weight must be consistent with engine support and pylon support weights of engine section Boeing: blade attachment (fraction rotor radius) blade fold weight rotor support structure weight (fraction maximum takeoff weight) duct weight per area parameters Custom Weight Model + + + + + + real real real real real rotor(8) _ Structure: WRotor rattach fWfold fWsupt Uduct WtParam 1.0 Default ), ’panel’ or ’hub’ or ’width’ or input) ’radius’ wingload or SizeParam ’ratio’ *DGW/ or fDGW ’span’ ), ( ’aspect’ ), area = ’WL’ , aspect ratio wing=’WL+xx’ _ ’chord’ or wing loading SET ’area’ to calculate span from rotor radius to calculate span from rotor radius, fuselage width, and clearance (tiltrotor) to calculate span from wing panel widths /S to calculate span from span of another wing to calculate span from rotor hub position (tiltrotor) D W W (for wing loading) f W = f ’ratio+XX’ ’radius+XX’ ’width+XX’ ’hub+XX’ ’panel+XX’ = input two of ( = = = = = W/S R A wing wing wing wing wing wing _ _ _ _ _ _ c SET SET SET SET SET SET b S fraction DGW wing parameters: for each wing; input two quantities, other two derived ( if wing sized from wing loading ( title notes wing loading area span chord aspect ratio Wing wing number Geometry Description + + + + + + + + + + Type c*100 c*1000 int real real real real real real Chapter 53 Structure: Wing Variable title notes kWing wingload fDGW area span chord AspectRatio 0 0 0 0.

1.0 .23 0.45 d c = airfoil Vdrag _ SET = tiltrotor; (+ aft) A wing (+ up) ¯ x _ A ¯ z MODEL ; RotorForSpan or ’hub’ ’hub’ or or ’width’ ’width’ = = RotorOnWing panel wing _ _ SET SET (fraction wing chord) A ¯ c tb w used for fuel in wing calculated for entire wing from wing panel geometry gets radius from box _ w sweep required for required for τ ’radius’ used by tiltrotor weight equations = and fWidth box _ wing and taper _ number of rotors mounted on wing rotor numbers ratio wing span to span of other wing, or to rotor radius other wing number rotor number for span (if nRotorOnWing=0) rotor at wing panel edge RotorOnWing RotorOnPanel SET taper, sweep, thickness used by weight equations fWidth thick rotors span calculation thickness ratio wing torque box chord twist taper ratio sweep (+ aft, deg) dihedral (+ up, deg) mean aerodynamic chord mean aerodynamic center chordwise offset from root aero center mean aerodynamic center vertical offset from root aero center Geometry Geometry (for graphics) Geometry (derived) + + + + + + + + + + + + + int int real int int int real real real real real real real real real box _ Structure: Wing nRotorOnWing RotorOnWing(nrotormax) fSpan otherWing RotorForSpan RotorOnPanel(npanelmax) thick fWidth twist taper sweep dihedral MAC xAC zAC 1 1 1. 1. 1. 1. 1. 1. 0. 0. 0. 0.

0.5 0.5 0.7 0.25 0.25 ’span+taper’ 2) b/ ( / p b , panel _ (+ aft) y ) fspan Ip (+ up) x 2) Ip z b/ (fraction panel chord) /S ( (fraction panel span) p p / p b p S p npanelmax b /c /c , , else free p f /b c c f 2) b = panel b/ _ = ( f (fraction panel chord) ’bratio’ b ref (fraction panel span) not required with only one panel y/ p f ref or span /c p /c = /c O F /b c I E E c c F panel y η p b _ ’span’ = S = input ratio to wing span, p = = input b (deg, + up) F SET b (deg, + aft) p p f c δ p : specify consistent definition of panels (span, edge, area, chord) Λ ’span’ ’bratio’ panel panel span: _ outboard edge, outboard edge, sweep dihedral chordwise offset at panel inboard edge vertical offset at panel inboard edge flap chord flaperon/aileron chord flap span flaperon/aileron span aileron aerodynamic center lateral position wing panels: SET aerodynamic center location number of wing panels (maximum aero center offset (1 fixed, 2 fraction root chord, 3 fraction inboard chord) panel parameters span (one side), area (both sides), mean chord, ratio span to wing span (one side), ratio area to wing area (both sides), ratio mean chord to wing chord, panel edges inboard chord ratio, outboard chord ratio, aerodynamic center locus control surfaces Geometry Wing Panels + + + + + + + + + + + + + + + + + + + + + + + + + + + + Location int int c*24 real real real real real real real real real real real real real real real real real real real panel(npanelmax) _ panel(npanelmax) flap(npanelmax) flaperon(npanelmax) panel(npanelmax) ACoffset panel(npanelmax) panel(npanelmax) panel(npanelmax) panel(npanelmax) panel(npanelmax) _ _ _ flap(npanelmax) flaperon(npanelmax) panel(npanelmax) _ _ panel(npanelmax) panel(npanelmax) panel(npanelmax) _ _ _ panel(npanelmax) _ _ _ _ _ aileron(npanelmax) _ _ _ _ wing _ _ Structure: Wing loc nPanel KIND SET span area chord fspan farea fchord edge fedge lambdaI(npanelmax) lambdaO(npanelmax) sweep dihedral dxAC dzAC fchord fchord fspan fspan fAC wing _ SET 2) b/ ( y/ (not used for tip panel) = E , else free η , /c ’adjust’ p /S c from wing area p , ’taper’ , or panel S , _ ref , ) c panel ) ’hub’ _ , fedge panel lambdaO from wing area ’cratio’ _ adjust free , p , , S fchord and farea hub ’width’ , , taper ’chord’ , , p c width E lambdaI p , , y cratio S ’station’ , , , , ’Sratio’ ) , panel radius _ panel free , panel chord _ , ’edge’ _ , ’area’ chord edge area bratio station Sratio , , , , from centerline and fraction wing semispan = input fraction wing semispan = from rotor radius = from adjacent input panel span or span ratio = input ratio to wing area, = input ratio to wing chord, = from rotor radius, fuselage width, and clearance (tiltrotor) = input edge = from chord ratios span area = input _ = input _ _ = from rotor hub position (tiltrotor) one and only one input panel span or span ratio that not used to define edge Ip c : from panel inboard edge, fraction panel span panel panel panel ’edge’ ’station’ ’radius’ ’width’ ’hub’ ’adjust’ ’area’ ’Sratio’ ’chord’ ’cratio’ ’taper’ all edges known (from input edge or station, or from adjacent panel span or span ratio) resulting edges unique and sequential if wing span calculated from panel widths: if known span: no input panel span or span ratio that not used to define edge if one or more taper (and no free), calculate if one (and only one) free, calculate _ _ nPanel=1 _ panel outboard edge: panel area or chord: require consistent definition of panel spans and outboard edges, and consistent with panel area or chord: aileron for SET SET SET _ fAC span ( edge ( area ( kind area and chord solution (1 tapered panels, 2 free panel) inboard chord Derived geometry int int int int real span(npanelmax) edge(npanelmax) area(npanelmax) _ _ _ panel panel panel area _ _ _ _ Structure: Wing iSET iSET iSET kind chordI(npanelmax) 0 2 0 0 3 3 3 1 0.

(from centerline, fraction wing semispan) y ) ) f f D D , , f f M M , , f f X X , , f f L L ) ) X X b (both sides) (one side) S Op X X − c − b S b S flap (1 fraction root flap; 2 increment relative root flap; 3 independent) aileron (1 fraction root aileron; 2 increment relative root aileron; 3 independent) incidence (1 fraction root incidence; 2 increment relative root incidence; 3 independent) outboard chord aileron aerodynamic center lateral position flap area/panel area flaperon-aileron area/panel area trailing edge flap factors ( trailing edge flap factors ( total flap area/wing area total flaperon-aileron area/wing area consistent geometry (0 if calculated geometry not consistent) extension (0 for none) wing panel number wing extension as kit (0 not kit) extension area extension span inboard area ( inboard span ( inboard flap area inboard flaperon-aileron area inboard wing aspect ratio inboard wing sweep inboard wing taper wing as kit (0 not, 1 kit, 2 kit as fixed useful load) kit weight (fraction total wing weight) kind deflection kind flaperon deflection (1 fraction flap; 2 increment relative flap; 3 independent) Wing Extensions Wing Kit Controls (each panel) + + + + + + + + + + + + + real real real real real real real real int int int int real real real real real real real real real int real int int int int ext _ flap(npanelmax) aileron(npanelmax) incid(npanelmax) flaperon(npanelmax) flap(npanelmax) flaperon(npanelmax) Wflap Wflaperon _ _ _ _ _ _ flap(4,npanelmax) flaperon(4,npanelmax) _ _ ext aileron(npanelmax) flapI flaperonI ext wing _ _ _ _ _ _ _ _ Structure: Wing chordO(npanelmax) eAC rArea rArea Ktef Ktef rArea rArea isConsistent SET kPanel KIT areaX spanX areaI spanI area area AspectRatioI sweepI taperI KIT fWkit KIND KIND KIND KIND 1 0 1 0 1 0 1 0 ) ) ) ) nvelmax nvelmax nvelmax nvelmax matrix) matrix) matrix) matrix) T T T T 2 piecewise linear, maximum 2 piecewise linear, maximum 2 piecewise linear, maximum 2 piecewise linear, maximum ≥ ≥ ≥ ≥ p i f p δ ap δ F p δ connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) flap flaperon aileron incidence + + + + + + + + + + + + + + + + + + + + + + + + int real int real real int real int real real int real int real real int real int real real flap(npanelmax) flaperon(npanelmax) aileron(npanelmax) incid(npanelmax) _ _ _ _ flap(ncontmax,nstatemax,npanelmax) flaperon(ncontmax,nstatemax,npanelmax) aileron(ncontmax,nstatemax,npanelmax) incid(ncontmax,nstatemax,npanelmax) _ _ _ _ Structure: Wing INPUT T nVflap(npanelmax) flap(nvelmax,npanelmax) Vflap(nvelmax,npanelmax) INPUT T nVflaperon(npanelmax) flaperon(nvelmax,npanelmax) Vflaperon(nvelmax,npanelmax) INPUT T nVaileron(npanelmax) aileron(nvelmax,npanelmax) Vaileron(nvelmax,npanelmax) INPUT T nVincid(npanelmax) incid(nvelmax,npanelmax) Vincid(nvelmax,npanelmax) 0.

) c 0 nvelmax c + AC T c = c (for each control state) and value 2 piecewise linear, maximum T ≥ L quant _ C wb ) D/q ( wing Aircraft%trim ) wing ) wing ) D/q ( D/q ( D/q ( wet for helicopter nominal drag (deg; 0 for not tilt) S i flight state specifies control state, or that control state obtained from conversion schedule can be zero, constant, or function of flight speed (CAS or TAS, piecewise linear input) for each component control, define matrix c by connecting aircraft control to comp control, flight state can specify comp control value initial values if control is connected to trim variable; otherwise fixed for flight state can be fraction total aircraft lift, lift, or number of speeds (0 zero value; 1 constant; target speeds (CAS or TAS) aircraft controls connected to individual controls of component, target definition determined by Klift wing lift model (0 none, 1 standard) incidence angle standard model wing cruise drag, area wing helicopter drag, area wing vertical drag, area wing-body interference drag, area total wetted area Trim Target Aerodynamics Derived drag + + + + + + + + int real real int real AWing real real real real real aero _ wing wing wing _ _ _ wb _ Structure: Wing nVlift Klift(nvelmax) Vlift(nvelmax) MODEL Idrag AWing DoQC DoQH DoQV DoQ Swet 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 xx=0.

_ TECH or xx=0 _ MODEL prim χ ; for fixed (input) weight use dWxx + flap χ efold model _ χ ext Wxx χ * xx fold _ χ fit fair χ χ TECH = wing primary structure wing extension fairing fittings flaps and control surfaces wing fold wing extension fold increment for weight on wing tips Wxx model (0 input, 1 NDARC, 2 custom) weight increment NDARC model (except tiltrotor) NDARC tiltrotor model tiltrotor model weight model result multiplied by technology factor and increment added: weight statement (component) wing group wing weight wing extension weight wing kit weight weight on wing tips wing primary structure (torque box) weight wing extension weight fairing weight fittings weight flaps and control surfaces weight wing fold weight wing extension fold weight Weight Technology Factors + + + + + + + + + + + + + + + + + + + + + Weight int real real real real real real real WWing WWingTR real real real real real real real real real real real real weight _ total ext kit prim ext fair fit flap wfold efold _ _ _ _ _ _ _ _ _ _ total _ Structure: Wing Weight MODEL dWprim dWext dWfair dWfit dWflap dWwfold dWefold WWing WWingTR xWtip Wwing Wwing Wwing Wtip TECH TECH TECH TECH TECH TECH TECH , without changing weight statements total _ Wtip adjusts rotor group, engine section or nacelle group, air induction group, engine system, drive system (less drive shaft), rotary wing and conversion flight controls, hydraulic group, trapped fluids, wing tip extensions negative increment required when engine and transmission not at tip location with rotor tiltrotor model requires weight on wing tips: both sides; calculated as sum of xWtip Structure: Wing 0 2 2 2 0. 0. 2.

1.5 0.8 0.0 5.73 0.25 0.85 0.43 0.75 0.012 Default ) ) R ) πeA M ac d ) ( ) c C / D D τ ) SC L SC ) C = L = C − ) ; 3 airfoil | | (derived) /dα δ δ L D D | | D/q L C 1 1 D/q C C Lα η η dC C = ( − − (per rad) 0 0 div η η (deg) Di , , M ) C 0 1 ; 3 3D ( /dα D η η ; 2 scaled, ; 2 scaled, L e 3 max /dα (deg) dC L L /dα D/q D/q zl C L = α dC dC (based on wing area, Lα ( (based on wing area, / C DV D max C 0 V C L ) ) ) C R = D/q D/q ( ( πeA ( max / specification (2 2D lift curve slope lift curve slope non-elliptical loading correction Oswald or span efficiency lift coefficient for minimum induced drag incompressible 3D lift curve slope 1 α control effectiveness factor control effectiveness factor lift-divergence Mach number pitch moment coefficient about aerdynamic center specification (1 fixed, area coefficient specification (1 fixed, area coefficient, zero lift angle of attack maximum lift coefficient compressibility correction (0 none, 1 lift, 2 drag, 3 both) lift pitch moment forward flight drag vertical drag Wing Aerodynamics, Standard Model Wing Drag, Standard Model Description + + + + + + + + + + + + + + + + + + + + + + + + Type real real int int real real real real real real real real real real real int real real int real real zl compress lift max drag Vdrag _ _ _ _ _ _ Dmin _ Chapter 54 Structure: AWing Variable AoA CLmax SET SET dCLda Tind Eind CL dCLda3D fDind AoA eta0 eta1 Mdiv CMac SET DoQ CD SET DoQV CDV 2 3 1 0. 0. 2. 0. 2. 0. 0.

1.4 2.5 10. 25.

0.74 0.31 0.0011 should be L C s X ) s α | − e α | ( s K D C = D C Δ d X | e α | d K ) 0 (deg) D D d C min f SC D = = α D ); other parameter calculated C ) is just the span efficiency factor for the induced power (and D/q Δ D cc CD e (deg) represents the wing parasite drag variation with lift, as well as the induced drag. C M s e 0 1 α at cc cc (deg) t M M Dcc α (–90 deg) C 2 scaled, d c (derived) (derived) ) or scaled (use D/q d s d s d s K X K X X X DoQ (based on wing area, Dwb wb C ) : fixed (use D/q varies with angle-of-attack, then ( angle of attack for wing minimum drag drag increment drag increment angle of attack for separation drag increment drag increment drag exponent drag exponent xxx _ Dp airfoil drag coefficient airfoil drag coefficient flap effectiveness factor model (0 none, 1 general, 2 quadratic) separated flow model (0 none, 1 general, 2 quadratic, 3 cubic) angle of attack for transition specification (1 fixed, area coefficient C critical Mach number constant critical Mach number constant SET Conventionally the Oswald efficiency If zero).

compressibility drag increment drag variation with angle of attack transition from forward flight drag to vertical drag wing-body interference drag + + + + + + + + + + + + + + + + + + + + real real real real real int real real real int real real real real real real int real real drag sep _ _ wb Dmin sep tran wb _ _ _ _ _ wb _ Structure: AWing cd90 fd90 CDcc Mcc0 Mcc1 MODEL AoA Kdrag Xdrag MODEL AoA Ksep Xsep Xd Xs AoA SET DoQ CD 0. 0. 0. 0.

negative for favorable rotor _ Kintp (rad/rad) d /dα wing(otherwing)=0.7 _ = for aft-on-front interference at rotors (0. for no interference) E Kint for front-on-aft interference int K at other wings (0. for no interference) int K wing(aftwing)=2. wing(frontwing)=0.

_ _ Kint Kint normal (helicopter) inplane (propeller) for tandem wings, typically for biplane wings, typically with mutual interference (as for biplane), require trim or other iteration for convergence interference power: inplane (propeller) factor angle of attack change at tail, interference factor interference power factor Interference velocity + + + + + + real real real real rotor(nrotormax) rotor(nrotormax) wing(nwingmax) _ _ _ Structure: AWing Etail(ntailmax) Kint Kintn Kintp 5. 3. 0.

1.0 0.0 0.0 0.0 200. 0.10 0.12 0.10 Default ) or kg/m (0 to 1) ) fold b or kg/m (knots) dive V flap f (fraction wing extension weight) , wing primary structure (lb/ft prim efold , wing extension (lb/ft f U ext U fold f fit fair f f weight per area fairing fittings flaps and control surfaces wing fold model (1 area, 2 parametric, 3 tiltrotor, 4 other) model (1 Boeing, 2 GARTEUR, Torenbeek (3 light, 4 transport), Raymer (5 transport, 6 general aviation)) lift factor parametric method: fraction wing span that folds Boeing: maximum fuselage width (fraction wing span) Boeing or Raymer: design dive speed GARTEUR: ratio maximum lift with and without flaps area method weight per area weight factors (fraction total wing weight) wing extension fold parameters Wing Group, NDARC Weight Model Custom Weight Model Description + + + + + + + + + + + + + + + + + + + Type int int real real real real real real real real real real real real real wing(8) _ wing other _ _ Chapter 55 Structure: WWing Variable MODEL MODEL fLift bFold wfus Vdive rflaplift Uprim Uext fWfair fWfit fWflap fWfold fWefold WtParam 1 0 1 1. 0.

2.0 0.5 0.8 0.9 0.8 0.8 1.5 0.20 0.23 0.30 600. 0.75 0.50 750.0 0.6048 0.4874 1.6384 0.5018 Default ) tip V , 2 cruise tip V , 1 hover freq _ V H F ) Vtip /σ w tip τ T V C (fraction rotor radius) tb C B e C , 1 fraction fuselage width, 2 fraction wing span) B /R F F ω ω fus (equivalent stiffness) sp /w pylon (equivalent stiffness) (equivalent strength) m e r T t j T C F wAttach ω C C at SDGW attach w (ft or m) jump n attach w rotor maximum blade loading load factor rotor tip speed (0. to use hover definition (0 input fraction width width beam bending frequency chord bending frequency torsion frequency reference rotor speed (0 from input rotor tip speed torque box beam bending torque box chord bending torque box torsion spar caps vertical/horizontal bending weight correction weight correction strength correction jump takeoff condition wing airfoil thickness-to-chord ratio width of wing structural attachments to body pylon radius of gyration wing mode frequencies (per rev, fraction rotor speed) form factors (1 calculate, 2 input) spar structural efficiency torque box structural efficiency tapered spar cap correction factors Wing Group, NDARC Tiltrotor Weight Model Description + + + + + + + + + + + + + + + + + + + + + + + + + + + + Type real real real real int real real real real real real int real int real real real real real real real real real form _ beam chord tors spar jump Attach refrpm freq jump pylon beam chord tors _ _ _ _ _ _ _ _ _ _ _ _ _ spar box t j m jump _ _ _ _ _ _ Chapter 56 Structure: WWingTR Variable CTs n Vtip thickTR SET fAttach wAttach fRG freq freq freq SET Vtip MODEL form form form form eff eff C C C 2. 3. 3. 0.

0.0 0.0 0.01 0.01 0.06 0.06 0.01 10.E6 10.E6 4.0E6 cruise _ and radius Vtip tip V rotor , cruise ref(1) ) _ Vtip U RotorOnWing(1) , m/m, kg/m U tb flap ) G 2 U ; or N/m 3 obtained from tb (fraction wing extension weight) E tb rotor; hover tip speed tip ρ or kg/m sp V ext efold ρ f sp U (fraction total wing weight excluding fold) E (fraction maximum thrust of one rotor) fold , in/in, lb/in f : attachment width used for both torsion stiffness and fairing area fair fit f U only used for tiltrotor wing weight RotorOnWing(1) Attach _ spar modulus torque box modulus torque box shear modulus spar ultimate strain allowable torque box ultimate strain allowable density spar cap density torque box fairing flaps and control surfaces wing extension fittings wing fold jump takeoff: hover wing frequencies: reference rotor rotation speed from rotor from thickTR SET material (lb/in weight per area (lb/ft weight factor wing extension fold parameters Custom Weight Model + + + + + + + + + + + + + + + + + + real real real real real real real real real real real real real real wingtr(8) _ spar box spar box _ _ _ _ spar box box _ _ _ Structure: WWingTR E E G StrainU StrainU density density Ufair Uflap UextTR fWfitTR fWfoldTR fWefoldTR WtParam 1 2 1 Default ’vol+aspect’ ) ’chord’ or ’aspect’ or aspect ratio ’span’ ), ( ’vol’ (tailarea * taillength / (diskarea * radius)) (tailarea * taillength / (wingarea * wingchord)) w (tailarea * taillength / (wingarea * wingspan)) c w w S/RA b w = or tail volume S/S V = S/S V = ’area’ V = input two of ( R A V used for geometry, baseline orientation, tail volume, tail weight model tail _ tail _ c SET or horizontal tail volume or vertical tail volume b S tail volume reference (1 wing, 2 rotor) wing or rotor number for tail volume KIND tail parameters: input two quantities, others calculated tail volume reference: tail volume title notes kind (1 horizontal tail, 2 vertical tail, 3 V-tail horizontal, 4 V-tail vertical) specification area span chord aspect ratio tail volume other V-tail number Empennage horizontal tail (0 vertical) V-tail (0 not) tail number Geometry Description + + + + + + + + + + + + + + Type c*100 c*1000 int int int int c*16 real real real real real int int int tail TailVol _ _ tail _ Chapter 57 Structure: Tail Variable title notes KIND isHortail isVtail kTail SET area span chord AspectRatio TailVol KIND TailVolRef otherVtail 1 0 0. 0.

1.0 .12 0.0 1.0 0.25 ) nvelmax ) set to zero cant otherVtail calculated, matrix) dihedral T 2 piecewise linear, maximum ≥ ) f D , f M , ) f X , BF f (fraction tail chord) chord C (fraction tail span) , L /c /b ) f AR r c f , δ b vol , span _ area _ tail (deg) V V _ b or rudder S tail φ _ e δ SET separately sized, aerodynamic loads for each; weight only for second tail, based on V-tail area and aspect ratio SET connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) V-tail: modeled as pair of horizontal and vertical tails (identified by taper ratio sweep (+ aft, deg) dihedral (deg) thickness ratio area ( length ( tail length control surface area/tail area trailing edge flap factors ( tail axes relative airframe, V-tail area V-tail span V-tail aspect ratio aerodynamic center location cant angle control surface chord control surface span elevator Geometry (for graphics and weights) Derived geometry Geometry Controls + + + + + + + + + + + + + + + + + real real real real int int real real real real real real real Location real real real int real int real real area len cont _ _ tail cont _ _ control _ cont _ cont(4) _ tail tail _ _ _ tail _ cont(ncontmax,nstatemax) _ Structure: Tail taper sweep dihedral thick iSet iSet Length rArea Ktef CBF(3,3) areaVtail spanVtail AspectRatioVtail loc cant fchord fspan INPUT T nVcont cont(nvelmax) Vcont(nvelmax) 1 0 1 1 0.0 ) c nvelmax 0 c + AC T c = c ) ) cant = 90 matrix) cant = 90 (for each control state) and value T 2 piecewise linear, maximum T ≥ tail V ) D/q ( tail ) D/q ( flight state specifies control state, or that control state obtained from conversion schedule i can be zero, constant, or function of flight speed (CAS or TAS, piecewise linear input) basic for each component control, define matrix c by connecting aircraft control to comp control, flight state can specify comp control value initial values if control is connected to trim variable; otherwise fixed for flight state connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) model (0 input, 1 NDARC, 2 custom) weight increment horizontal tail cant angle: + to left (vertical tail for vertical tail cant angle: + to right (horizontal tail for aircraft controls connected to individual controls of component, incidence model (0 none, 1 standard) standard model tail drag, area tail vertical drag, area total wetted area weight statement (component) tail (empennage group) Aerodynamics Derived drag Weight + + + + + + + + + + + + + int real int real real int ATail real real real Weight int real aero weight incid _ _ _ tail _ tail _ incid(ncontmax,nstatemax) _ Structure: Tail INPUT T nVincid incid(nvelmax) Vincid(nvelmax) MODEL ATail DoQ DoQV Swet Weight MODEL dWtail 0.0 1.0 1.0 xx=0.

_ TECH or xx=0 _ MODEL ; for fixed (input) weight use dWxx + model _ Wxx * xx vt _ χ or fold TECH ht χ χ = fold Wxx NDARC model weight model result multiplied by technology factor and increment added: tail weight tail weight fold weight Technology Factors + + + + real WTail real real real tail tfold total _ _ _ Structure: Tail dWfold WTail Wtail TECH TECH 0 2 2 2 0. 1. 1.

0.8 0.0 5.73 0.25 0.85 0.43 0.75 0.011 Default ) ) ) ) τ R D D πeA SC ) ( L SC / C = ) = ) | | ) (derived) /dα δ δ 0 D D | | L L 1 1 C C Lα η η D/q C D/q dC C − − − (per rad) 0 0 div L η η (deg) , , C M ) 0 1 ; 3 3D /dα D η η ; 2 scaled, ; 2 scaled, L 3 = ( max /dα (deg) dC L Di L /dα D/q D/q zl C C L = ( α dC e dC Lα ( (based on tail area, (based on tail area, / C D DV max 0 V C C L ) ) ) C R = D/q D/q ( ( πeA ( max / specification (2 2D lift curve slope lift curve slope non-elliptical loading correction Oswald efficiency lift coefficient for minimum induced drag incompressible 3D lift curve slope 1 α control effectiveness factor control effectiveness factor lift-divergence Mach number specification (1 fixed, area coefficient specification (1 fixed, area coefficient zero lift angle of attack maximum lift coefficient compressibility correction (0 none, 1 lift, 2 drag, 3 both) lift forward flight drag vertical drag Tail Aerodynamics, Standard Model Tail Drag, Standard Model Description + + + + + + + + + + + + + + + + + + + + + + Type real real int int real real real real real real real real real real int real real int real real zl compress lift max drag Vdrag _ _ _ _ _ _ Dmin _ Chapter 58 Structure: ATail Variable AoA CLmax SET SET dCLda Tind Eind CL dCLda3D fDind AoA eta0 eta1 Mdiv SET DoQ CD SET DoQV CDV 0. 0. 2.

25.

0.74 0.31 0.0011 d X | e α | d K D (deg) C min = D D α ); other parameter calculated C Δ cc CD M 0 1 at cc cc (deg) t M M Dcc α C ) or scaled (use d d K X DoQ (derived) d X : fixed (use angle of attack for tail minimum drag drag increment drag increment xxx _ model (0 none, 1 general, 2 quadratic) exponent angle of attack for transition critical Mach number constant critical Mach number constant SET compressibility drag increment drag variation with angle of attack transition from forward flight drag to vertical drag + + + + + + + + + + real real real int real real real real real drag _ Dmin tran _ _ Structure: ATail CDcc Mcc0 Mcc1 MODEL AoA Kdrag Xdrag Xd AoA 3 1 1 1 1 1 3. 0.

0.2 0.8 0.0 200.

Default ) tail _ (fraction span) max V ht KIND /b 25 .

f ) w vt = 1 b vt ) S ( Vdive / ) ht h or kg/m ht S ( (lb/ft (knots) (fraction total tail weight excluding fold) tail U dive fold = SLS max speed, f V max V model (1 helicopter or compound, 2 tiltrotor or tiltwing, 3 area, 4 other) model (1 GARTEUR, Torenbeek (2 low speed, 3 transport), Raymer (4 transport, 5 general aviation)) Torenbeek or Raymer: kind (1 fixed, 2 variable incidence) Raymer: fuselage width at horizontal tail model (1 helicopter or compound, 2 tiltrotor or tiltwing, 3 area, 4 other) model (1 GARTEUR, Torenbeek (2 low speed, 3 transport), Raymer (4 transport, 5 general aviation)) AFDD: antitorque placement (0 none, 1 on tail boom, 2 on vertical tail) Torenbeek or Raymer: kind (1 conventional, 2 T-tail) Torenbeek: T-tail factor weight per area weight models can use taper ratio, sweep, and thickness ratio dive speed: model (1 horizontal tail, 2 vertical tail, 3 based on horizontal tail vertical tail design dive speed area method fold weight factor parameters Tail, NDARC Weight Model Custom Weight Model Description + + + + + + + + + + + + + + + + + + + Type int int int int real int int int int real real real real real tail(8) _ tail Htail Hother Vtail Vother _ _ _ _ _ Htail Vtail AntiQ _ _ _ Chapter 59 Structure: WTail Variable MODEL MODEL MODEL KIND wfus MODEL MODEL place KIND fTtail Vdive Utail fTfold WtParam ’ ’ 6.5 1.0 1.0 0.0 42.8 Default , other weight parameters , other weight parameters fWauxtank , density _ fWtank ; energy , fFuelWing , eWauxtank , energy _ eWtank ; specific cap , − fuel battery f _ (lb/gallon or kg/liter) (MJ/kg) density _ cap fuel fuel (weight, lb or kg) − ρ (energy, MJ) fuel e , IDENT , fuel cap cap d cap − − _ cap _ fuel fuel W E Waux Eaux , , cap cap _ _ Wfuel Efuel fuel weight per volume fuel energy per weight fraction wing torque box filled by fuel tanks fuel capacity fuel capacity ratio capacity to mission fuel capacity increment store and use weight: energy calculated from weight use store and use energy: fuel weight zero use title notes fuel quantity stored and used (1 weight, 2 energy) fuel weight properties fuel tank sizing battery identification Fuel Tank System tank number Configuration Description + + + + + + + + + + + + + + + Type c*100 c*1000 int int real real real real real real real c*16 energy battery _ _ cap cap cap _ cap burn _ _ _ density _ _ Chapter 60 Structure: FuelTank Variable title notes kTank SET fuel specific fFuelWing(nwingmax) Wfuel Efuel fFuel dFuel IDENT 1 1 0 0. 0.

0.0 0.0 0.0 0.0 0.0 1000.

20000.

(energy) ) cap _ Efuel *((maximum mission fuel)+(reserve fuel)) BatteryModel of cap _ nauxtankmax *(maximum mission fuel), (maximum mission fuel)+(reserve fuel)) cap ident _ fFuel (weight) or cap + _ Efuel cap _ cap _ or fFuel cap Wfuel _ dFuel = = max( Wfuel = calculate from mission fuel used and mission battery discharge power cap cap _ _ = function of mission fuel used Efuel Efuel : input : calculate from mission fuel used (weight) or or (energy) (fraction auxiliary fuel weight) (MJ/kg) cap i cap cap cap (each tank) _ − eq _ − P , scale with density aux , constant , scale with temperature aux d t auxtank auxtank W E tank=’input’ tank=’miss’ Wfuel tank=’miss+power’ tank=’f(miss)’ Wfuel eq0 eq eq f e _ _ _ _ auxtank P P P ) SET SET SET SET D/q ( fuel tank sizing: usable fuel capacity battery identification: energy storage only, match specific fuel consumption: weight (lb/hp-hr or kg/kW-hr) or energy (MJ/hp-hr or MJ/kW-hr) placement (1 internal, 2 sponson, 3 wing, 4 combination) number of auxiliary tank sizes (minimum 1, maximum fuel capacity fuel capacity tank weight tank weight drag location model (0 for none) specific fuel consumption power loss power loss power loss deice power loss Geometry (for graphics) Auxiliary Fuel Tank Equipment power + + + + + + + + + + + + + + + + + int int real real real real real Location int real real real real real Peq _ cap(nauxtankmax) cap(nauxtankmax) _ auxtank(nauxtankmax) _ 0 d t deice _ auxtank(nauxtankmax) _ _ _ _ _ Structure: FuelTank place Mauxtanksize Waux Eaux fWauxtank(nauxtankmax) eWauxtank(nauxtankmax) DoQ loc MODEL sfc Peq Peq Peq Peq 0. 0.

1.0 1.0 xx=0.

_ TECH or xx=0 _ MODEL ) battery _ ; for fixed (input) weight use IDENT (kW/kg) ) dWxx .

3 + , from cap / _ tank e model Wfuel _ wing mbd − x Wxx (volume) * BatteryModel fuel = xx cap _ W plumb tank − χ at DGW takeoff conditions (lb/hr or kg/hr) χ tank π fuel F V TECH = tanks and support; battery management system plumbing; power distribution (wiring) Wxx model (0 input, 1 NDARC, 2 custom) weight increment NDARC model weight model result multiplied by technology factor and increment added: fuel capacity wing fuel capacity wing fuel capacity (fraction number of components in fuel tank system battery identification ( specific power weight statement (component, not including auxiliary tanks) fuel system (propulsion group) number of main engines total fuel flow fuel tank weight plumbing weight Derived Weight Technology Factors + + + + + + + + + real real real int int real Weight int real real WTank int real real real tank _ weight power _ wing in _ tank plumb wing _ _ cap _ _ _ _ Structure: FuelTank Vfuel Wfuel rWfuel ncomp kBatteryModel specific Weight MODEL dWtank dWplumb WTank Neng fuelflow TECH TECH 2 4 2 2 4 2.5 3.2 1.0 1.0 2.0 0.4 0.2 0.2 0.09 150.

0.727 Default crash _ KIND , Ktoler , int _ plumb N ntank (1.0 to 2.5) bt f ; parametric method uses tank X tank fWtank K (lb) int N (fraction total fuel system weight) (lb) (MJ/liter) 0plumb plumb (fraction fuel capacity weight) K tank f ρ 1plumb (MJ/kg) K tank f tank e : fraction method uses tank _ model (1 fraction, 2 parametric, Torenbeek (3 integral, 4 generic), Raymer (5 transport, 6 general aviation)) number of internal tanks tank weight parametric: ballistic tolerance factor parametric: survivability (1 baseline, 2 UTTAS/AAH level of survivability) Torenbeek (generic): factor Torenbeek (generic): exponent Raymer: integral tank capacity (fraction total) Raymer: protected tank capacity (fraction total) model (1 fraction, 2 parametric) total number of fuel tanks (internal and auxiliary) for plumbing weight increment weight factor plumbing weight fuel tank plumbing tank weight tank volume density battery management system (fraction basic tank weight) power distribution (wiring) weight (fraction basic tank weight) MODEL weight storage energy storage Fuel System, NDARC Weight Model Description + + + + + + + + + + + + + + + + + + + + + + + Type int int real real int real real real real int int real real real real real real real tank plumb _ _ density int crash _ _ _ plumb plumb _ _ Chapter 61 Structure: WTank Variable MODEL ntank fWtank Ktoler KIND Ktank Xtank fint fprot MODEL nplumb K0 K1 fWplumb eWtank energy fBMS fwire 0.

= 2.

plumb _ K1 is a crashworthiness and survivability factor; typically is the sum of weights for auxiliary fuel, in-flight refueling, pressure refueling, inerting system, etc.; typically = 50 to 250 lb plumb plumb plumb _ _ _ K1 K0 K0 parameters Custom Weight Model + + real tank(8) _ Structure: WTank WtParam 1 0 Default ) ngearmax number of rotors in group number of main rotors number of engine groups first engine group propulsion group is set of components and engine groups, connected by drive system components (rotors) define power required, engine groups define power available drive system defines ratio of rotational speeds of components (relative primary rotor speed) title notes primary rotor rotors in group (0 no, 1 main rotor, 2 other) primary engine group engine groups in propulsion group (0 no, 1 only produce power, 2 can consume power) engine group generator or compressor, can consume shaft power (0 only produce power) number of states (maximum drive system state for variable speed transmisson (0 for none) Propulsion Group propulsion group number Specification Drive system Description + + + + + + Type c*100 c*1000 int int int int int int int int int int int int var _ group(nengmax) _ prim group(nrotormax) gear prim _ main _ in _ _ _ _ in _ Chapter 62 Structure: Propulsion Variable title notes kPropulsion kRotor rotor nRotor nRotor kEngine engine nEngineGroup firstEngineGroup canConsumePower nGear STATE 2 2 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.9 1.0 0.04 (control) included gear f ), others dependent (specify gear ratio) tip V , gear ratio factor ) var _ ds _ ) to use gear ratio #n gear _ nGear fLength STATE (1 to state=n ) or calculated (from ds _ limit windage (fraction component+transmission power) P DS P Length acc i (fraction total component power required) (fraction total engine power) acc , scale with density and rpm P , constant , scale with density d n xmsn to use conversion schedule, IRfan acc0 acc acc : input (use DS P P P length specify primary engine group only if no rotors in propulsion group state=0 when evaluate rotational speed of dependent rotors and engines _ length factor drive system branches: one primary rotor per propulsion group (specify drive system state: identifies gear ratio set for multiple speed transmissions variable speed transmission: for drive system state SET model (1 fraction component power required; 2 with function drive shaft limit) gear box loss power loss due to windage power loss power loss power loss deice power loss ECU (etc.) power loss IRS fan loss drive shaft length (1 input; 2 calculated) drive system power limit drive system power limit factor Transmission losses Accessory losses Geometry Drive system torque limit + + + + + + + + + + + + + + + + + + int real real real real real real real real int real real real real Xloss ds _ ds ds _ xmsn _ ds _ ECU IRfan windage _ _ 0 d n deice _ _ length _ _ _ _ _ _ Structure: Propulsion MODEL fPloss Ploss Pacc Pacc Pacc Pacc fPacc fPacc SET Length fLength Plimit fPlimit ’ ’ 1.0 ) ) xmsn xmsn _ _ DESIGN DESIGN ) ) ds _ ) ’maxPQ’ xx limit _ _ or ) SET fPlimit Peng × GW=’maxQ’ _ ) nEngine SET , depending on ) eng req P P ) or calculate (from eng or ) xx ) N ( _ av P G P =on) ∑ ds or eng _ Plimit =on) /P Plimit nratemax quant=’Q margin’ eng _ _ P ) EG Plimit fPlimit ) ) _ max eng av req P ( P P = input (use ’input’ eng eng limit ds = N N _ input ( ( FltState%SET ds DS _ (engine group at primary rotor reference speed) ds ∑ ∑ P (if limit _ ) ) FltState%SET × × _ quant=’Q margin’ limit av _ _ limit es es (if es P prim prim _ _ SET _ ) Plimit Ω Ω DS av trim : / / SET : from takeoff power, : from engine power required at transmission sizing conditions and missions ( P P : from engine power available at transmission sizing conditions and missions ( ref ref M CQ Plimit fPlimit fPlimit : : : (Ω (Ω f /f ds ds _ _ ds=’input’ ds=’ratio’ ds=’Pav’ ds=’Preq’ es=0 es=1 es=2 _ _ _ _ _ _ _ fPlimit fPlimit limit limit limit limit limit limit limit _ _ _ _ _ _ _ SET SET SET SET SET SET SET can be used for trim ( used for drive system weight, tail rotor weight, transmission losses limits propulsion group limits engine group can be used for max effort in flight state ( can be used for max gross weight in flight condition or mission ( always check and print whether exceed torque limit drive system torque limits: engine shaft: options for drive system power limit: corresponds to power of all engines of propulsion group (all engine groups) engine shaft power limit: corresponds to all engines of engine group ( rotor shaft power limit: corresponds to one rotor all limits number of ratings (1 for only continuous, maximum drive system rating designation torque limit factor drive system torque limit ( MCQ or MCP rating number drive system rating designation torque limit factor ( Drive system ratings Derived drive system limit + + + + int c*12 real real int c*12 real ds(nratemax) ds(nratemax) ds ds(nratemax) ds ds(nratemax) _ _ _ _ ds _ _ _ Structure: Propulsion nrate rating frating Qlimit krateC arating xrating 0 0 1 0.

) c nvelmax 0 c + ) is the rating factor x AC T c ’MCP’ = or c =off), first engine shaft limit and then drive system limit ’MCQ’ Plimit matrix) _ (for each control state) and value T 2 piecewise linear, maximum FltState%fPower T ≥ is the rotor speed ratio and r FltState%SET , where limit , primary rotor or primary engine (rpm) N rxP Δ , drive system rating not used ≤ ds _ flight state specifies control state, or that control state obtained from conversion schedule can be zero, constant, or function of flight speed (CAS or TAS, piecewise linear input) nrate gear box blank to use engine ratings of first engine group limit at flight state is if for each component control, define matrix c by connecting aircraft control to comp control, flight state can specify comp control value initial values if control is connected to trim variable; otherwise fixed for flight state connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) model (0 input, 1 NDARC, 2 custom) weight increment the engine model gives the power available, accounting for installation losses and mechanical limits then the power available is reduced by the factor next torque limits are applied (unless drive system ratings: must include maximum continuous ( aircraft controls connected to individual controls of component, rotational speed increment weight statement (component, not including EngineGroup) drive system (propulsion group) Control Weight + + + + + + + + + + + + int real int real real Weight int real DS DN _ _ DN(ncontmax,nstatemax) _ Structure: Propulsion INPUT T nVDN DN(nvelmax) VDN(nvelmax) Weight MODEL dWgb 1 1 0. 0. 0. 0. 0.

1.0 1.0 1.0 1.0 1.0 1.0 xx=0.

_ TECH or xx=0 _ MODEL ; for fixed (input) weight use dWxx + model _ Wxx * rb xx ds rs _ χ χ gd gb χ χ χ cl χ TECH = rotor shaft drive shaft rotor brake clutch gas drive Wxx NDARC model weight model result multiplied by technology factor and increment added: drive system weight = gear box (including rotor shaft) + drive shaft + rotor brake + clutch + gas drive tiltrotor wing weight model requires weight on wing tip (drive system, without rotor shaft) drive system state for weight EngineGroup for weight weight on wing tip weight gear box and rotor shaft gear box weight rotor shaft weight drive shaft weight rotor brake weight clutch weight gas drive weight Technology Factors + + + + + + + + + + + + + + real real real real real WDrive int int real real real real real real real real wt _ wt _ gear _ gb rs ds rb cl gd _ _ _ _ _ _ Structure: Propulsion dWrs dWds dWrb dWcl dWgd WDrive STATE kEngineGroup Wtip Wgbrs TECH TECH TECH TECH TECH TECH 1 7 4 0 6 0.

1.0 5.0 0.13 0.03 0.15 0.15 Default = 0.15 (0.18 for 2-bladed teeter) (fraction total drive system rated torque) Q f fPower trgb K (fraction total drive system rated power) (excluding rotor shafts) = 0.03, P f ds N gb N fTorque (fraction gear box and rotor shaft weight) =0.6 for twin main rotors (tandem, coaxial, tiltrotor) rs f fPower = *(otherrotor RPM)/(mainrotor RPM) = 0.13 (data range 0.06 to 0.20) fTorque fTorque fShaft = tail rotor drive weight increment factor tail rotor power (fraction total drive system rated power) tail rotor gear ratio model (1 AFDD83, 2 AFDD00, 3 other) model (1 Boeing, 2 Boeing (alternate), GARTEUR (3 helicopter, 4 tiltrotor), 5 Tishchenko) rotor shaft weight AFDD83: number of gear boxes AFDD83: second (main or tail) rotor rated torque Boeing: number of stages in main-rotor drive other: separate tail rotor drive weight increment (0 none) number of intermediate drive shafts second (main or tail) rotor rated power fPower typically for single main rotor and tail rotor, typically gear box (including rotor shafts) drive shaft (AFDD82) parameters Drive System, NDARC Weight Model Custom Weight Model Description + + + + + + + + + + + + + + + + + Type int int real int real int int real real real int real real drive(8) _ gbrs other _ _ tr other _ _ tr _ Chapter 63 Structure: WDrive Variable MODEL MODEL fShaft ngearbox fTorque nstage KIND Ktrgb fPower gear ndriveshaft fPower WtParam 1 1 1 1 0 1 0.

’ ’ 1.0 ’MCP’ ’MCP’ Default ’Engine’ ’RPTEM’ ) to) _ ) motor ref(rating , _ P0 comp , recip , table (ratio rpm to rpm of primary rotor in propulsion group) , ) prim ) gear Ω / xxx RPTEM _ _ , 2 spec Nspec engine engine = Ω _ _ r eng (SLS static at takeoff rating, 0. for N eng MODEL MODEL P second system identification number of main engines takeoff power rating idle power rating engine gear ratio title notes engine model engine identification number of engines engine power fuel tank system number rotor number for reaction drive group number drive system branch (1 primary, 0 dependent) gear ratio input (1 from engine model ( engine model ( can consume shaft power (0 only produce power), generator or compressor can produce shaft power (0 only consume power) convertible engine, reaction drive (0 not) convertible engine, turbojet/fan (0 not) Engine Group engine group number Description Propulsion Group Derived Description + + + + + + + + + + + + + + + + + + + Type c*100 c*1000 int c*32 c*16 c*16 int int real c*12 c*12 int int int int int real int int int int int int engine engine main _ _ engine system2 gear _ react _ _ _ to idle _ xmsn engine _ _ _ _ Chapter 64 Structure: EngineGroup Variable title notes kEngineGroup MODEL IDENT IDENT nEngine nEngine Peng rating rating kFuelTank kRotor kPropulsion KIND INPUT gear(ngearmax) iMODEL KIND canConsumePower canProducePower isConvertReact isConvertJet ) engine _ MotorModel or IDENT ; fuel is energy ) , from ; not use fuel ; fuel is weight CompressorModel MotorModel or BatteryModel MotorModel → or EngineModel engine EngineModel _ → ; fuel is weight CompressorModel → RecipModel ) ; fuel is weight → to or EngineModel _ ; fuel is weight IDENT ; not use fuel → system2 engine ; fuel is energy (generated, not burned) _ _ ; fuel is energy engine CompressorModel rating _ ( EngineTable ; not use weight or MotorModel → IDENT system2 IDENT RecipModel EngineTable for motor); → _ ); IDENT or → ’takeoff’ MotorModel ) ≥ MotorModel = 1 engine _ ) → engine IDENT B ) or B EngineModel CompressorModel → _ RecipModel engine ); of _ idle or → _ system2 IDENT = 1 _ engine engine _ IDENT EngineModel _ B ident IDENT rating engine ( of _ BatteryModel IDENT and no size task (or engine power not sized) IDENT .

(0 for none) IDENT ’idle’ EngineTable ident IDENT = 0 or , from eng = turboshaft engine (RPTEM); P one eng kFuelTank = compressor for reaction drive; = turbojet/turbofan (mode = motor + generator (mode W = motor + fuel cell; can be set to : engine model ’shaft’ : convertible engine; only with turboshaft , = reaction drive (mode EngineModel EngineModel = electric motor; = compressor; ’+fan’ = turboshaft engine (table); = reciprocating engine; , = electric generator; engine engine _ _ ’RPTEM’ ’table’ ’recip’ ’comp’ ’comp+react’ ’motor’ ’gen’ ’motor+gen’ ’motor+cell’ ’+react’ ’+jet’ or energy (motor, may have MODEL MODEL engine identification: match second system identification: match number of main engines: for fuel tank weight for fixed engine: use takeoff power rating: for engine scaling, aircraft power loading, fuel tank weight FltState%rating fuel tank system identified for burn must store and use weight (turboshaft, reciprocating, fuel cell) fuel tank system identified for generation must store and use energy (may have identification ( identification ( battery model, from number of ratings rating designations (lowercase) MCP rating number takeoff power rating number weight one engine reference engine speed (at drive state #1) int int int int c*12 int int real real eng sys2 _ _ to _ Structure: EngineGroup kModel kModel kBattery nrate rating(nratemax) krateC krate WOneEng Nref 1.0 (control) included gear f , gear ratio factor var values not used for first group _ gear _ fPsize .

and of primary rotor of propulsion group, or specify gear ratio > STATE ref _ power _ Vtip ) SET ) C and ) , to distribute power required among engine groups ) ˙ C ) peak m C C P C 0 ) sfc Nspec C or opt0 ) ST N /SP P ) C = C spec 0 mech 0 C P P N perf=’engine’ 0 sfc _ g = F = sfc C for first engine group, must be specifies distribution of power required for flight state SIZE C ˙ ) 0 m ˙ n w f Preq fPsize SP _ : calculate gear from : used if ) gear _ P power when evaluate rotational speed of engine must size at least first engine group; calculate _ drive system branch: primary engine group only designated if no rotors for propulsion group INPUT variable speed transmission: for drive system state SET FltState%SET specification (0 sized, 1 fixed) sized power ratio power ( specific power ( mechanical limit of power ( specific fuel consumption at MCP ( gross jet thrust at MCP ( specification engine speed ( optimum engine speed at MCP ( mass flow at MCP ( fuel flow at MCP ( specific fuel consumption ( specific fuel consumption at MCP gross jet thrust at MCP, jet/fan only fuel flow at MCP Sizing Engine model performance parameters (one engine) Engine model performance parameters (one engine), ratio converted to base + + + int real real real real real real real real real real real real real real conv _ conv conv _ _ power _ Structure: EngineGroup SET fPsize P0(nratemax) SP0(nratemax) Pmech(nratemax) sfc0C Fg0C Nspec Nopt0C mdot0C wdot0C sfc0(nratemax) rsfc0C rFg0C rwdot0C 1.0 ) ) xmsn xmsn _ _ DESIGN DESIGN ) ) ds _ ) es ’maxPQ’ _ limit _ or ) SET fPlimit Peng × GW=’maxQ’ _ ) nEngine SET , depending on ) eng req P P ) or calculated (from eng or ) es N ( _ av P G P =0 ∑ ds or eng _ Plimit =on) Fg0C /P quant=’Q margin’ eng _ P EG sfc0 Plimit fPlimit ) ) _ max , with eng av req used P ( P P = input (use sfc0C ’input’ eng eng , limit used, and ds = N N Nspec _ input ( ( , ds DS _ (engine group ds at engine reference speed) ∑ ∑ P limit _ ) ) FltState%SET × × _ Nspec wdot0C limit , , _ Pmech es es (if limit es prim prim _ _ , limit SET _ Plimit Ω Ω av ES : P0 / / SET : from takeoff power, : from engine power required at transmission sizing conditions and missions ( ES P P : from engine power available at transmission sizing conditions and missions ( Pmech P , mdot0C ref ref Plimit fPlimit fPlimit : : : P0 (Ω (Ω ds ds _ _ ds=’input’ ds=’ratio’ ds=’Pav’ ds=’Preq’ es=0 es=1 es=2 _ _ _ _ _ _ _ fPlimit fPlimit limit limit limit limit limit limit limit _ _ _ _ _ _ _ SET SET SET SET SET SET SET limits engine group can be used for max effort in flight state ( can be used for max gross weight in flight condition or mission ( always check and print whether exceed torque limit reciprocating: only motor or generator: only motor + fuel cell: and drive system torque limits: engine shaft: options for engine shaft power limit: corresponds to all engines of engine group ( engine shaft (0 input, 1 fraction power, 2 fraction drive system limit) engine shaft power limit engine shaft power limit factor engine shaft torque limit ( Drive system torque limit Derived engine shaft limit + + + + int real real real es _ es es es _ _ _ limit _ Structure: EngineGroup SET Plimit fPlimit Qlimit 0 1 2 ’x’ 0.0 0.0 0.0 0.8 1.05 0.98 0.75 0.01 0.75 1.05 0.015 0.007 0.030 0.015 ) gbrs W and ES auxair _ W eta , auxair _ ; 3 scaled, ES fMF , W f f d f f d K K Kffd ) ) ) ) ; 2 scaled, a a ’–z’ in operating condition , /P /P Swet ) Dscale IRS ’+z’ _ loss loss _ M , P P δ ’–y’ Units , ( ; motor + fuel cell: STATE 3 in / 2 ’+y’ Kffd ) , aux aux η η (IRS off) eng ’–x’ (fraction, for ex ex ex , d η w/N in ( ’+x’ / (fraction engine mass flow) (fraction engine mass flow) wet aux aux S f f (per engine) = k wet S engine inlet loss inlet particle separator loss engine exhaust loss mass flow ram recovery efficiency engine exhaust loss mass flow ram recovery efficiency engine exhaust loss rotor number area factor, installation power losses = inlet + particle separator + exhaust (including IRS) IR suppressor state specified by motor or generator: only use deterioration factor on engine fuel flow or performance engine inlet efficiency power losses (fraction power available, auxiliary air momentum drag (IRS off) power losses (IRS on) auxiliary air momentum drag (IRS on) deterioration factor on engine fuel flow or performance power losses (fraction power available, location nominal orientation ( position (0 standard, 1 tiltrotor) nacelle/cowling wetted area (1 fixed, input Installation IR suppressor Convertible Geometry + + + + + + + + + + + + + + + + + + + + + + + + + + + + real real real real real real real real real real real real Location c*16 int int int real real IRon IRon _ conv IRon _ _ _ inlet ps exh exh exh _ _ _ _ _ conv geom Swet auxair auxair _ _ _ d _ auxair _ auxair engine _ _ _ _ Structure: EngineGroup Kffd eta fPloss fPloss fPloss fMF eta fPloss fMF eta Kffd fPloss loc direction SET RotorForEngine SET Swet kSwet 1 0 1 0 ) ) nvelmax nvelmax ) RotorForEngine kSwet (engine system and drive system) EG matrix) matrix) /N T T 2 piecewise linear, maximum 2 piecewise linear, maximum ≥ ≥ gbrs W / 2 + ) or calculated (from /kg ES W Swet (fixed) or m 3 BF / C /lb nac S are ft (fixed) f f (engine system) or 123) e e =tiltrotor: calculate lateral position (BL) from ± kSwet ES (fixed engine group power) : calculation override part of location input 123) geom , wetted area: input (use W _ A ± = geom Swet SET units of w B _ _ connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) SET SET nacelle wetted area used for nacelle drag, and for cowling weight engine group nacelle must be consistent with rotor pylon nacelle/cowling area total wetted area nominal orientation (1, –1, 2, –2, 3, –3) axis incidence ( axis yaw ( orientation (1 fixed) engine axes relative airframe, engine direction, engine direction, amplitude mode Derived geometry Controls + + + + + + + + + + + + + real real int int int int real real real int real int real real int real int real real amp mode _ _ nac _ incid yaw _ _ amp(ncontmax,nstatemax) mode(ncontmax,nstatemax) _ _ Structure: EngineGroup Snac Swet iDirection axis axis isFixed CBF(3,3) ef0(3) ef(3) INPUT T nVamp amp(nvelmax) Vamp(nvelmax) INPUT T nVmode mode(nvelmax) Vmode(nvelmax) 1 0 1 0 1 0 1 0.

) ) ) c nvelmax nvelmax nvelmax 0 c + AC T c = c matrix) matrix) matrix) (for each control state) and value T T T 2 piecewise linear, maximum 2 piecewise linear, maximum 2 piecewise linear, maximum T ≥ ≥ ≥ (variable speed transmission only) gear for helicopter nominal drag (deg; 0 for not tilt) f i (tilt) flight state specifies control state, or that control state obtained from conversion schedule i can be zero, constant, or function of flight speed (CAS or TAS, piecewise linear input) for each component control, define matrix c by connecting aircraft control to comp control, flight state can specify comp control value initial values if control is connected to trim variable; otherwise fixed for flight state ψ connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) aircraft controls connected to individual controls of component, incidence yaw gear ratio factor model (0 none, 1 standard) incidence angle standard model Nacelle Drag + + + + + + + + + + + + + + + + + + + + + int real int real real int real int real real int real int real real int real DEngSys drag incid yaw fgear _ _ _ _ incid(ncontmax,nstatemax) yaw(ncontmax,nstatemax) fgear(ncontmax,nstatemax) _ _ _ Structure: EngineGroup INPUT T nVincid incid(nvelmax) Vincid(nvelmax) INPUT T nVyaw yaw(nvelmax) Vyaw(nvelmax) INPUT T nVfgear fgear(nvelmax) Vfgear(nvelmax) MODEL Idrag DEngSys 1 1 1 1 0.0 0.0 0.0 0.0 0.0 0.0 0.0 nac ) nac ) nac ) D/q ( (engine, exhaust, accessories) D/q ( D/q ( ES W since pylon connected to rotor shaft axes engine system engine section or nacelle air induction exhaust accessories engine support engine cowling pylon support air induction tiltrotor with tilting engines use pylon drag (and no nacelle drag), tiltrotor with nontilting engines, use nacelle drag as well model (0 input, 1 RPTEM or NASA, 2 custom) weight increment (all engines) model (0 input, 1 NDARC, 2 custom) weight increment NDARC model component drag contributions must be consistent pylon is rotor support, and nacelle is engine support nacelle cruise drag, area nacelle helicopter drag, area nacelle vertical drag, area weight statement (component, including engine weight) engine weight engine system (except engine), engine section or nacelle group, air induction group engine weight engine system weight weight on wing tip Derived drag Weight + + + + + + + + + + + + + + + + real real real Weight int real int int int real real real real real real WEngSys real real real weight sys nac air _ _ _ _ nac nac nac total _ _ _ _ Structure: EngineGroup DoQC DoQH DoQV Weight MODEL dWEng MODEL MODEL MODEL dWexh dWacc dWsupt dWcowl dWpylon dWair WEngSys Weng WES Wtip 1.0 1.0 1.0 1.0 1.0 1.0 1.0 xx=0.

_ TECH or xx=0 _ MODEL used for rotor pylon wetted area, engine nacelle wetted WES ; for fixed (input) weight use dWxx + supt χ EngineModel airind model χ acc _ χ cowl exh pylon Wxx χ χ χ * xx _ eng χ TECH = Wxx weight model result multiplied by technology factor and increment added: engine system weight = engine + exhaust + accessory ( area, rotor moving weight) nacelle weight = support + cowl + pylon engine weight parameters in tiltrotor wing weight model requires weight on wing tip: engine section or nacelle group, air induction group, engine system engine weight engine cowling weight pylon structure weight engine support structure weight air induction system weight exhaust system weight engine accessories weight Technology Factors + + + + + + + + real real real real real real real eng cowl pylon supt air exh acc _ _ _ _ _ _ _ Structure: EngineGroup TECH TECH TECH TECH TECH TECH TECH 2 2 2.0 Default ) ) D D SC SC = = ); other parameter calculated ) ) D/q D D/q D CD C C ; 2 scaled, ; 2 scaled, ) or scaled (use D/q D/q DoQ (based on wetted area, (based on wetted area, D DV d 0 V C C ) ) X : fixed (use D/q D/q ( ( xxx _ specification (1 fixed, area coefficient specification (1 fixed, area coefficient exponent SET forward flight drag vertical drag transition from forward flight drag to vertical drag Nacelle Drag, Standard Model Description + + + + + + + + + + + Type int real real int real real real drag Vdrag _ _ Chapter 65 Structure: DEngSys Variable SET DoQ CD SET DoQV CDV Xdrag 1 1 1 0.

0.0 20. 0.2 0.3 0.0 0.002 Default C P vs exh W ) (fraction maximum takeoff weight) or kg/m pylon (lb/ft f C P (fraction engine support plus air induction weight) vs airind U W airind f = 0.3 (data range 0.1 to 0.6) nac X nac fWair K 0exh 1exh factor exponent K K lubrication system weight (1 in engine weight, 2 in accessory weight) typically engine support and pylon support weights must be consistent with rotor support structure weight model (1 parametric, 2 scale with power, 3 Boeing, 4 Raymer (transport)) pylon support structure weight nacelle group weight, Boeing: crash load factor Raymer: nacelle width (fraction nacelle length) model (1 parametric, 2 area) air induction weight weight per nacelle area exhaust system weight, per engine, including IR suppressor; engine accessories parameters Engine Section or Nacelle Group, NDARC Weight Model Air Induction Group, NDARC Weight Model Engine System, NDARC Model Custom Weight Model Description + + + + + + + + + + + + + + + + + + + + Type int real real real real real int real real real real int real engsys(8) _ nacelle airind lub _ _ _ nac _ exh exh _ _ clf _ Chapter 66 Structure: WEngSys Variable MODEL fWpylon Knac Xnac n fWidth MODEL fWair Uair Kwt0 Kwt1 MODEL WtParam 1 1 0.

’ ’ ’Jet’ ’MCT’ ’MCT’ Default ’RPJEM’ ) to) _ ref(rating _ T0 ) ) ) jet system2 _ _ simple , IDENT IDENT ) , from , from xxx RPJEM _ _ jet jet _ _ jet one jet N W JetModel JetModel (SLS static at takeoff rating, 0. for jet MODEL MODEL T second system identification takeoff thrust rating idle thrust rating title notes jet model jet identification number of jets jet thrust fuel tank system number rotor number for reaction drive jet model ( jet model ( convertible engine (0 not) identification ( identification ( number of ratings rating designations (lowercase) MCT rating number takeoff thrust rating number weight one jet Jet Group jet group number Description Derived Description + + + + + + + + + + + + + Type c*100 c*1000 int c*32 c*16 c*16 int real c*12 c*12 int int int int int int int int c*12 int int real jet jet _ jet sys2 _ jet system2 react _ _ _ _ to idle _ jet to _ _ _ _ Chapter 67 Structure: JetGroup Variable title notes kJetGroup MODEL IDENT IDENT nJet Tjet rating rating kFuelTank kRotor iMODEL KIND isConvertReact kModel kModel nrate rating(nratemax) krateC krate WOneJet 0.0 1.05 0.98 0.01 0.75 0.007 ; fuel is weight JetModel → jet _ JetModel ; fuel is weight IDENT → JetModel ; not use weight system2 _ → jet _ ) IDENT JetModel a ); of /T ) IDENT = 1 loss C ) ) T 0 ) B ident C f f d C ) M sfc K T δ C JetModel C ) /ST T of C C 0 aux mech and no size task (or jet thrust not sized) η T (IRS off) . T = sfc = turbojet/turbofan engine (RPJEM); ident = ex C = sfc = 0 0 C ˙ 0 in C w (fraction, for ’fan’ ˙ 0 , jet m d (fraction engine mass flow) ) ˙ w T 0 η ’jet’ aux ST , = reaction drive (mode f = simple force generator; no model identified; fuel is weight or energy : jet model = reaction drive (RPJEM)); : convertible engine; only with turbojet/turbofan jet jet ) _ _ T ’RPJEM’ ’react’ ’simple’ ’+react’ engine inlet loss engine exhaust loss mass flow ram recovery efficiency MODEL MODEL jet identification: match second system identification: match for fixed jet: use thrust ( specific thrust ( mechanical limit of thrust ( specific fuel consumption at MCT ( mass flow at MCT ( fuel flow at MCT ( energy flow at MCT ( specific fuel consumption at MCT fuel flow at MCT deterioration factor on jet fuel flow jet inlet efficiency power losses (fraction thrust available, auxiliary air momentum drag (IRS off) Jet model performance parameters (one jet) Jet model performance parameters (one jet), ratio converted to base Installation + + + + + + + + + real real real real real real real real real real real real real real real conv _ conv inlet exh _ _ _ auxair d _ auxair _ _ Structure: JetGroup T0(nratemax) ST0(nratemax) Tmech(nratemax) sfc0C mdot0C wdot0C Edot0C rsfc0C rwdot0C Kffd eta fTloss fTloss fMF eta 1 1 2 ’x’ 0.0 1.0 0.0 0.8 0.03 0.01 0.75 1.05 0.01 burn _ SET ) ) in operating condition ) ; 2 scaled) ’–z’ a jet , _ Dscale _ /T Swet IRS ’+z’ _ loss , T Units ( f f d ’–y’ 3 K / , STATE ) ’+y’ jet , aux η /N ’–x’ max ex ex , ES T W ( ’+x’ / (fraction engine mass flow) (per jet) wet aux S S f (per jet) = k wet S engine exhaust loss mass flow ram recovery efficiency engine exhaust loss area factor, installation power losses = inlet + exhaust (including IRS) IR suppressor state specified by fuel tank system identified must be consistent with power losses (IRS on) auxiliary air momentum drag (IRS on) deterioration factor on jet fuel flow power losses (fraction thrust available, design maximum thrust fuel quantity used (1 weight, 2 energy) thrust specific fuel consumption (weight or energy) specific weight weight group (1 engine system, 2 propeller/fan installation, 3 tail rotor) location nominal orientation ( nacelle/cowling wetted area (1 fixed, input IR suppressor Convertible Simple force generator Geometry + + + + + + + + + + + + + + + + + + + + + + real real real real real real int real real int Location c*16 int real real IRon IRon _ conv IRon _ _ _ exh exh simple _ _ _ conv burn Swet auxair _ _ _ _ auxair jet _ _ Structure: JetGroup fTloss fMF eta Kffd fTloss Tmax SET sfc SW KIND loc direction SET Swet kSwet 1 0 1 0 ) ) nvelmax nvelmax ) kSwet matrix) matrix) T T 2 piecewise linear, maximum 2 piecewise linear, maximum ≥ ≥ / ) or calculated (from /kg Swet or m / /lb (fixed) BF nac S are ft C 123) (fixed) ± 0 kSwet f f e e 123) , wetted area: input (use A ± Swet units of B _ connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) SET nacelle wetted area used for nacelle drag, and for cowling weight nacelle/cowling area total wetted area nominal orientation (1, –1, 2, –2, 3, –3) axis incidence ( axis yaw ( orientation (1 fixed) jet relative airframe, jet direction, jet direction, amplitude mode Derived geometry Controls + + + + + + + + + + + + + real real int int int int real real real int real int real real int real int real real amp mode _ _ nac _ incid yaw _ _ amp(ncontmax,nstatemax) mode(ncontmax,nstatemax) _ _ Structure: JetGroup Snac Swet iDirection axis axis isFixed CBF(3,3) ef0(3) ef(3) INPUT T nVamp amp(nvelmax) Vamp(nvelmax) INPUT T nVmode mode(nvelmax) Vmode(nvelmax) 1 0 1 0 1 0.

) ) c nvelmax nvelmax 0 c + AC T c = c matrix) matrix) (for each control state) and value T T 2 piecewise linear, maximum 2 piecewise linear, maximum T ≥ ≥ nac ) nac ) nac ) D/q ( D/q ( D/q ( for helicopter nominal drag (deg; 0 for not tilt) i (tilt) flight state specifies control state, or that control state obtained from conversion schedule i can be zero, constant, or function of flight speed (CAS or TAS, piecewise linear input) for each component control, define matrix c by connecting aircraft control to comp control, flight state can specify comp control value initial values if control is connected to trim variable; otherwise fixed for flight state ψ connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) aircraft controls connected to individual controls of component, incidence yaw model (0 none, 1 standard) incidence angle standard model nacelle cruise drag, area nacelle helicopter drag, area nacelle vertical drag, area weight statement (component, including jet weight) Nacelle Drag Derived drag Weight + + + + + + + + + + + + + + + + int real int real real int real int real real int real DJetSys real real real Weight drag incid yaw _ nac _ _ nac nac _ _ _ incid(ncontmax,nstatemax) yaw(ncontmax,nstatemax) _ _ Structure: JetGroup INPUT T nVincid incid(nvelmax) Vincid(nvelmax) INPUT T nVyaw yaw(nvelmax) Vyaw(nvelmax) MODEL Idrag DJetSys DoQC DoQH DoQV Weight 1 1 1 1 0.0 0.0 0.0 0.0 0.0 0.0 0.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 xx=0.

_ TECH or xx=0 _ MODEL used for nacelle wetted area) WES ; for fixed (input) weight use dWxx + supt χ airind model χ (engine, exhaust, accessories) acc _ χ JetModel cowl exh ES pylon Wxx χ χ χ * W xx _ TECH jet = χ engine system engine section or nacelle air induction exhaust accessories engine support engine cowling pylon support air induction Wxx model (0 input, 1 RPJEM, 2 custom) weight increment (all jets) model (0 input, 1 NDARC, 2 custom) weight increment NDARC model weight model result multiplied by technology factor and increment added: engine system weight = engine + exhaust + accessory ( nacelle weight = support + cowl + pylon jet weight parameters in jet weight engine system (except jet), engine section or nacelle group, air induction group jet weight engine system weight jet weight engine cowling weight pylon structure weight engine support structure weight air induction system weight exhaust system weight engine accessories weight Technology Factors + + + + + + + + + + + + + + + + + + + + + + + int real int int int real real real real real real WJetSys real real real real real real real real real weight sys nac air _ _ _ _ jet jetcowl jetpylon jetsupt jetair jetexh jetacc _ _ _ _ _ _ _ total _ Structure: JetGroup MODEL dWJet MODEL MODEL MODEL dWexh dWacc dWsupt dWcowl dWpylon dWair WJetSys Wjet WES TECH TECH TECH TECH TECH TECH TECH 2 2 2.0 Default ) ) D D SC SC = = ); other parameter calculated ) ) D/q D D/q D CD C C ; 2 scaled, ; 2 scaled, ) or scaled (use D/q D/q DoQ (based on wetted area, (based on wetted area, D DV d 0 V C C ) ) X : fixed (use D/q D/q ( ( xxx _ specification (1 fixed, area coefficient specification (1 fixed, area coefficient exponent SET forward flight drag vertical drag transition from forward flight drag to vertical drag Nacelle Drag, Standard Model Description + + + + + + + + + + + Type int real real int real real real drag Vdrag _ _ Chapter 68 Structure: DJetSys Variable SET DoQ CD SET DoQV CDV Xdrag 1 1 1 0.

0.0 20. 0.2 0.3 0.0 0.002 Default ) C T (fraction maximum takeoff weight) or kg/m vs exh pylon (lb/ft W f C T (fraction engine support plus air induction weight) vs airind U W airind f nac X nac K 0exh 1exh factor exponent K K lubrication system weight (1 in jet weight, 2 in accessory weight) model (1 parametric, 2 scale with thrust, 3 Boeing, 4 Raymer (transport)) pylon support structure weight nacelle group weight, Boeing: crash load factor Raymer: nacelle width (fraction nacelle length) model (1 parametric, 2 area) air induction weight weight per nacelle area exhaust system weight, per jet; engine accessories parameters Engine Section or Nacelle Group, NDARC Weight Model Air Induction Group, NDARC Weight Model Engine System, NDARC Model Custom Weight Model Description + + + + + + + + + + + + + + + + + + + + Type int real real real real real int real real real real int real jetsys(8) _ nacelle airind lub _ _ _ nac _ exh exh _ _ clf _ Chapter 69 Structure: WJetSys Variable MODEL fWpylon Knac Xnac n fWidth MODEL fWair Uair Kwt0 Kwt1 MODEL WtParam 1 1 0.

’ ’ ’MCP’ ’MCP’ Default ’Charge’ ) to) _ ) charge ref(rating _ _ P0 IDENT ) , from solarcell , ) xxx fuelcell SolarCellModel _ _ (0 for none) or charge charge _ _ chrg one chrg (SLS static at takeoff rating, 0. for kFuelTank N W chrg MODEL MODEL P FuelCellModel takeoff thrust rating idle thrust rating title notes charger model charger identification number of chargers charger power fuel tank system number (generated) fuel tank system number (burned) charger model ( charger model ( identification ( battery model, from number of ratings rating designations (lowercase) MCP rating number takeoff thrust rating number weight one charger Charge Group charge group number Description Derived Description + + + + + + + + + + + + Type c*100 c*1000 int c*32 c*16 int real c*12 c*12 int int int int int int int c*12 int int real burn _ charge charge _ chrg _ charge _ _ to idle charge to _ _ _ _ Chapter 70 Structure: ChargeGroup Variable title notes kChargeGroup MODEL IDENT nCharge Pchrg rating rating kFuelTank kFuelTank iMODEL KIND kModel kBattery nrate rating(nratemax) krateC krate WOneChrg ’x’ 0.0 0.8 1.05 0.75 0.007 ) BatteryModel f f d ; fuel generated is energy K SolarCellModel ; fuel generated is energy or ) ) ) burn ; 2 scaled) _ ’–z’ , Dscale Swet SolarCellModel _ ’+z’ FuelCellModel ) , → C FuelCellModel → Units kFuelTank ( of ’–y’ ) sfc , and no size task (or charger power not sized) 3 C / . 0 charge 2 ) _ P charge ident ’+y’ C _ = 0 , aux η chrg ’–x’ chrg IDENT , ) = sfc P IDENT C W/N nac ( C ’+x’ S ˙ 0 / m (fraction charger mass flow) ˙ w wet : charger model aux S f fuel burned is weight ( (per charger) = = solar cell; = fuel cell; k charge ) wet _ S P ’fuel’ ’solar’ mass flow ram recovery efficiency area factor, MODEL charger identification: match for fixed charger: use fuel tank system identified for generation must store and use energy (may have fuel tank system identified for burn must store and use weight power ( specific fuel consumption at MCP ( mass flow at MCP ( fuel flow at MCP ( solar cell total area deterioration factor on charger fuel flow or performance auxiliary air momentum drag location nominal orientation ( nacelle/cowling wetted area (1 fixed, input nacelle/cowling area total wetted area Charger model performance parameters (one charger) Installation Geometry + + + + + + + + + + + real real real real real real real real Location c*16 int real real real real nac Swet auxair _ _ _ auxair charger _ _ Structure: ChargeGroup P0(nratemax) sfc0C mdot0C wdot0C solararea Kffd fMF eta loc direction SET Swet kSwet Snac Swet 1 0 1 0 1 0 ) ) ) nvelmax nvelmax nvelmax ) kSwet matrix) matrix) matrix) T T T 2 piecewise linear, maximum 2 piecewise linear, maximum 2 piecewise linear, maximum ≥ ≥ ≥ / ) or calculated (from /kg Swet or m (fixed) / BF /lb C (fixed) are ft f f e e 123) ± kSwet 123) , wetted area: input (use (tilt) A ± i Swet units of B _ connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) SET nacelle wetted area used for nacelle drag nominal orientation (1, –1, 2, –2, 3, –3) axis incidence ( axis yaw ( orientation (1 fixed) charger relative airframe, charger direction, charger direction, amplitude mode incidence Derived geometry Controls + + + + + + + + + + + + + + + + + + + int int int int real real real int real int real real int real int real real int real int real real amp mode incid _ _ _ incid yaw _ _ amp(ncontmax,nstatemax) mode(ncontmax,nstatemax) incid(ncontmax,nstatemax) _ _ _ Structure: ChargeGroup iDirection axis axis isFixed CBF(3,3) ef0(3) ef(3) INPUT T nVamp amp(nvelmax) Vamp(nvelmax) INPUT T nVmode mode(nvelmax) Vmode(nvelmax) INPUT T nVincid incid(nvelmax) Vincid(nvelmax) 1 0 1 1 0.

0.0 ) c nvelmax 0 c + AC T c = c matrix) (for each control state) and value T 2 piecewise linear, maximum T ≥ nac ) nac ) nac ) D/q ( (engine, exhaust, accessories) D/q ( D/q ( ES W for helicopter nominal drag (deg; 0 for not tilt) i flight state specifies control state, or that control state obtained from conversion schedule can be zero, constant, or function of flight speed (CAS or TAS, piecewise linear input) for each component control, define matrix c by connecting aircraft control to comp control, flight state can specify comp control value initial values if control is connected to trim variable; otherwise fixed for flight state ψ connection to aircraft controls (0 none, 1 input control matrix number of speeds (0 zero value; 1 constant; values speeds (CAS or TAS) model (0 input, 1 NDARC, 2 custom) weight increment (all chargers) NDARC model aircraft controls connected to individual controls of component, yaw model (0 none, 1 standard) incidence angle standard model nacelle cruise drag, area nacelle helicopter drag, area nacelle vertical drag, area weight statement (component, including charger weight) charger weight charge group weight engine system weight Nacelle Drag Derived drag Weight + + + + + + + + + + + + + int real int real real int real DChrgSys real real real Weight int real WChrgSys real real drag weight yaw _ _ total nac _ nac nac _ _ _ _ yaw(ncontmax,nstatemax) _ Structure: ChargeGroup INPUT T nVyaw yaw(nvelmax) Vyaw(nvelmax) MODEL Idrag DChrgSys DoQC DoQH DoQV Weight MODEL dWChrg WChrgSys Wchrg WES 1.0 xx=0.

_ TECH or used for nacelle wetted area) xx=0 _ WES MODEL SolarCellModel or ; for fixed (input) weight use dWxx + FuelCellModel model _ Wxx * xx _ chrg χ TECH = Wxx weight model result multiplied by technology factor and increment added: engine system weight = engine + exhaust + accessory = charge group weight ( charger weight parameters in charger weight Technology Factors + + real chrg _ Structure: ChargeGroup TECH 2 2 2.0 Default ) ) D D SC SC = = ); other parameter calculated ) ) D/q D D/q D CD C C ; 2 scaled, ; 2 scaled, ) or scaled (use D/q D/q DoQ (based on wetted area, (based on wetted area, D DV d 0 V C C ) ) X : fixed (use D/q D/q ( ( xxx _ specification (1 fixed, area coefficient specification (1 fixed, area coefficient exponent SET forward flight drag vertical drag transition from forward flight drag to vertical drag Nacelle Drag, Standard Model Description + + + + + + + + + + + Type int real real int real real real drag Vdrag _ _ Chapter 71 Structure: DChrgSys Variable SET DoQ CD SET DoQV CDV Xdrag 0.

Default parameters Custom Weight Model Description + + Type real chrgsys(8) _ Chapter 72 Structure: WChrgSys Variable WtParam 0. 0. 0. 0.

0.25 Default ’Engine’ ’Default’ ) w/P sfc = ˙ N n F F ) ) , net jet thrust eng , net jet thrust ’T800’ G X g F P F 2eng K input (eg + P 1eng , gross jet thrust , gross jet thrust K q req ); fuel flow = specific fuel consumption * power ( P + EngineGroup P ˙ m of P/ ) 0eng ) K = m ( ˙ = engine _ SP W SW , and parameters for optimum speed, thrust available, and performance , power required , 2 a ) IDENT , power required ref _ P P av ( model ( P C W sfc0C required, but not used; weight, ratings, technology, and scaling variables not used P , ” = MCP ref ref _ vs _ C eng SP0 eng sfc0C W 0eng 1eng 2eng X and K K K ref _ only use P0 constant constant constant exponent ” = SLS static; “ engine identification: used by installed: power available uninstalled: power available “ mass flow = power / specific power ( engine model can be used by more than one engine group, so all parameters fixed as model for turbojet or reaction drive of convertible engine: title notes identification RPTEM model (0 fixed, 1 engine weight (fixed) engine weight, Engine Model engine model number Weight Description + + + + + + + + + + + + Type c*100 c*1000 c*16 int int real real real real real weight _ eng eng eng _ _ _ eng _ Chapter 73 Structure: EngineModel Variable title notes ident kEngineModel MODEL Weng Kwt0 Kwt1 Kwt2 Xwt 4. 5. 0.

10.

150. 0.45 2000. 2500.

’MCP’ 20000. 20000.

; typically designated as ) tech ) FltState m lim ) ) model = ˙ ) C m C ) C ˙ 0 C 0 ) m ( = ˙ C ˙ sfc 0 /P m opt0 ˙ ) ˙ ) ref R m m N ( vs C spec nratemax SW mech mech lim ) eng N P P C SF SW = P/W C /SP = ) g R 0 0 F ) SP SP C SW = Maximum Rated Power (5 or 10 min) = Maximum Continuous Power (normal operations) = Emergency Rated Power (OEI power) = Contigency Rated Power (2.5 min) /P = Intermediate Rated Power (30 min) ) R 0 0 P P if MCP scaled, ratios to MCP values kept constant ’ERP’ ’CRP’ ’MRP’ ’IRP’ ’MCP’ engine model being used may not contain data for all ratings specific weight reference specific weight limit Reference Engine Rating: SLS, static engine rating: match rating designation in engine weight, parameters number of ratings (maximum rating designations MCP rating number power ( specific power ( mechanical limit of power ( specific fuel consumption at MCP ( specific jet thrust ( specification turbine speed ( optimum turbine speed at MCP ( power ( specific power ( mechanical limit of power ( Custom Weight Model Parameters Engine Ratings Reference Derived ratios + + + + + + + + + + + + + + + + + real real real int c*12 int real real real real real real real real real real engine(8) _ ref ref(nratemax) _ ref _ ref ref _ _ _ ref(nratemax) ref limit _ ref(nratemax) _ _ _ Structure: EngineModel SW SW WtParam nrate rating(nratemax) krateC P0 SP0 Pmech sfc0C SF0C Nspec Nopt0C rP0(nratemax) rSP0(nratemax) rPmech(nratemax) 0. 0. 0. 0.

30.

200. 0.34 tech _ ) ref/SP0C _ ref _ tech _ P0C = sfc0C limit=0.

_ tech = ˙ m MF C , tech=Nspec _ or C C sfc tech ˙ _ m ˙ m Nspec and vs , size=1 ) vs _ sfc0C C ref 0 , ref C _ tech _ 0 FIX _ sfc ) tech SP ), _ ) sfc0C ), SP0C weight ref spec _ _ = spec SP0C limit N _ tech=sfc0C ) C ref(MCP) N _ 2 0 _ (0. for Nspec MF C N s 0 , SP MODEL with scale, use SP0 C ˙ sfc0C tech K m 0 , ) ˙ m limit sfc lim vs _ (0. for SW lim vs ˙ sfc m (0. for ( C spec tech sfc0C eng , and N , ref(MCP) N sfc tech opt0 _ sfc N (otherwise get , lim SP limit P/W and _ lim SP = SP ˙ m SP SP0C tech=SP0 < SW _ functions are defined by values tech sfc SP0C ˙ 0 1 m 1 2 0 1 0 1 sp sp sf c sf c N s N s N o sw sw and limits and power limit K K K K K K K K K SP defaults require for no variation of engine weight scaling determined by specific power at MCP specific fuel consumption at MCP specification turbine speed engine size (0 scaled, 1 fixed) mass flow at limit specific power limit specific fuel consumption limit specification turbine speed variation ( specific power available (SLS static, MCP, specific fuel consumption (SLS static, MCP, specification turbine speed, optimum turbine speed, engine weight, Technology Scaling Derived scaling + + + + + + + + + + real real real int real real real real real real real real real real real real real real tech tech limit tech limit _ _ _ _ _ limit size limit _ _ _ Structure: EngineModel SP0C sfc0C Nspec FIX MF SP0C sfc0C KNspec P0C Ksp0 Ksp1 Ksfc0 Ksfc1 KNs1 KNs2 KNo Ksw0 Ksw1 1 1 1 1. 0. 1. 0. 0. 0. 0. 1.

2.0 N η X and opt N OptN=0 _ MODEL ) ) spec N ( t nspeedmax /η ) N ( t η C = single set; no variation if spec /P q /P q P N/N param P _ unique and sequential vs A B vs C opt opt opt opt0 opt1 opt2 opt3 N N η INPUT N N N N N N rNeng spec opt0 X X K K K K K K /N /N opt opt N N constant constant constant constant constant constant exponent exponent number of engine speeds (maximum engine speed ratio, parameter sets engine power and performance variation with power turbine speed determined by used only for interpolated: model (0 none, 1 linear, 2 cubic) linear, cubic, power turbine efficiency function, parameter input form (1 single set; 2 interpolated) single set interpolated Optimum Power Turbine Speed Power Available and Power Required Parameters + + + + + + + + + + + + + + + + + + int real real real real real real real real int EngineParam int real EngineParam OptN _ param _ Structure: EngineModel MODEL KNoptA KNoptB KNopt0 KNopt1 KNopt2 KNopt3 XNopt XNeta INPUT Param nspeed rNeng(nspeedmax) ParamN(nspeedmax) 1 0 0 0. .3 1. 0.

-.2 -.3 3.5 -2.5 Default ) ) ) ) ) b θ θ θ θ ) ) ) ) 1 1 1 1 K vs temperature θ θ θ θ ) ) , 1 1 1 1 0 K K X X b vs temperature θ ˙ K K X X m + + 0 + + / 0 0 0 0 + + + + a 0 0 0 0 ˙ K K X X m /SP K K X X a = = = = nengkmax-1 = = nengkmax-1 = = ; 2 values SP mf a mf a mf a mf a spa spa spa spa K K K X X , K K X X ) K ) ParamN ParamN (piecewise linear (piecewise linear b (piecewise linear (piecewise linear b , 2 (piecewise linear (piecewise linear (piecewise linear (piecewise linear b b 0 1 − 0 1 − 0 1 − 0 1 − spa spa spa mf a mf a mf a spa spa spa mf a mf a mf a Param b b b b number of regions (maximum K K θ K X X θ X number of regions (maximum K K θ K X X θ X input form (1 coefficients referred specific power available, referred mass flow at power available, Parameters parent (1 engine model number engine param number ( Power Available Description + + + + + + + + + + + + + + + + + + + + + + + Type int int int int int real real real real real real real real int real real real real real real real real lin _ Chapter 74 Structure: EngineParam Variable parent kEngineModel kEngineParam INPUT Nspa(nratemax) Kspa0(nengkmax,nratemax) Kspa1(nengkmax,nratemax) Tspak(nengkmax,nratemax) Kspab(nengkmax,nratemax) Xspa0(nengkmax,nratemax) Xspa1(nengkmax,nratemax) Tspax(nengkmax,nratemax) Xspab(nengkmax,nratemax) Nmfa(nratemax) Kmfa0(nengkmax,nratemax) Kmfa1(nengkmax,nratemax) Tmfak(nengkmax,nratemax) Kmfab(nengkmax,nratemax) Xmfa0(nengkmax,nratemax) Xmfa1(nengkmax,nratemax) Tmfax(nengkmax,nratemax) Xmfab(nengkmax,nratemax) .2 .8 0. 0. .6 .1 .2 .8 0. 0.

1.3 .78 3.5 2.0 -.48 ) N ( a P give ) N ( t η , so N insensitive to t values) , SLS static /η a P spec spec N+1 ( N N = b G K , P b ) C 1 0 θ C unique and sequential K /P , b q θ /P C P q 0 K P vs /P N q vs C P or 0 C ˙ b m vs ˙ sets) or values / K w , / C N b 0 req ( g θ req ˙ m ˙ w /F K g N-1 , different, resulting F = optimum power turbine speed K K opt N unique and sequential b θ : adjacent 0 1 2 3 1 : 0 1 2 3 0 1 2 3 b K f f q mf q f gq K , f f q f f q f f q f f q mf q mf q mf q mf q f gq f gq f gq f gq , X X X b K K K K K K K K K K K K K θ = referred specific power and mass flow available, at = referred specific power and mass flow available, at a = specification power turbine speed ˙ ˙ m m input form = coefficients form not input is calculated ( input input , , a 0 = power turbine speed, = power turbine efficiency; assume gas power available constant constant constant constant exponent constant constant constant constant exponent constant constant constant constant exponent spec t piecewise linear function: N SP SP N η referred fuel flow at power required, referred mass flow at power required, gross jet thrust at power required, Performance at Power Required + + + + + + + + + + + + + + + + + + + real real real real real real real real real real real real real real real Structure: EngineParam Kffq0 Kffq1 Kffq2 Kffq3 Xffq Kmfq0 Kmfq1 Kmfq2 Kmfq3 Xmfq Kfgq0 Kfgq1 Kfgq2 Kfgq3 Xfgq .8 .6 0. 0.

ex (installed thrust loss) vs g /F G F 0 1 2 3 f gr f gr f gr f gr K K K K constant constant constant constant installed net jet thrust at power required, + + + + + real real real real Structure: EngineParam Kfgr0 Kfgr1 Kfgr2 Kfgr3 1 0 0 ’MCP’ Default ’Engine’ ’Default’ input EngineGroup of ) engine ) _ ) IDENT nengtmax ) nratemax nengtmax ) spec N ) h, V, R ) ( a P h, V, R ( h, V, R N ( ˙ F w (TAS) h V engine identification: used by engine table can be used by more than one engine group, so all parameters fixed title notes identification number of ratings (maximum rating designations MCP rating number number of altitudes (maximum number of speeds (maximum altitude speed power available fuel flow net thrust Engine Table engine table number Engine ratings Specification turbine speed ( Table Description + + + + + + + + + + + + + + + + Type c*100 c*1000 c*16 int int c*12 int real int int real real real real real Chapter 75 Structure: EngineTable Variable title notes ident kEngineTable nrate rating(nratemax) krateC Nspec nalt nspeed alt(nengtmax) speed(nengtmax) Tp(nengtmax,nengtmax,nratemax) Tw(nengtmax,nengtmax,nratemax) Tf(nengtmax,nengtmax,nratemax) 1.0 1.0 1.0 exh=0 _ fPloss , not used) ps=0 _ auxair _ fPloss , eta , inlet=0 auxair _ _ fMF fPloss not used d _ eta ) weight=0 not used); _ fPsize , MODEL ( , accounting for deterioration of engine efficiency power Kffd _ dWEng SET from table; mechanical limits included in power available data Peng engine not scaled ( fixed engine weight no mass flow value, so no momentum drag of auxillary air flow ( obtain tables intended for installed engine, including losses, so expect fuel flow multiplied by power available fuel flow net thrust Factors + + + + real real real Structure: EngineTable Kp Kw Kf 0. 0. 0. 0. 0.

0.25 Default ’Engine’ ’Default’ N n F F ) , net jet thrust , net jet thrust G eng g X F F P 2eng input K ); mass flow = fuel flow / fuel-air ratio + P w/P , gross jet thrust , gross jet thrust 1eng q K req P EngineGroup P sfc = ˙ + of 0eng K engine _ = W , power required a IDENT , power required P av P model ( P ) ) P vs ( W eng eng W 0eng 1eng 2eng X K K K constant constant constant exponent engine identification: used by installed: power available uninstalled: power available fuel flow = specific fuel consumption * power ( reciprocating engine model can be used by more than one engine group, so all parameters fixed title notes identification model (0 fixed, 1 engine weight (fixed) engine weight, parameters Reciprocating Engine Model reciprocating engine model number Weight Custom Weight Model Description + + + + + + + + + + + + + + Type c*100 c*1000 c*16 int int real real real real real real recip(8) _ weight _ eng eng eng _ _ _ eng _ Chapter 76 Structure: RecipModel Variable title notes ident kRecipModel MODEL Weng Kwt0 Kwt1 Kwt2 Xwt WtParam 0.

0.60 0.08 1000. 1000. 1000. 2000. 2000.

’MCP’ ) R ; typically designated as = 0. for limit independent of engine speed /N Xcrit spec FltState N * R ) ) ) /P R ) spec mep ) ) P nratemax 0 ˙ spec mech m /N 0 ) N P ) R sfc ) 0 0 R crit SF N N 0 P = /P g R F ) crit ) 0 C P F = Maximum Rated Power (5 or 10 min) = Maximum Continuous Power (normal operations) /P : zero for no mechanical (mep) limit ) R : zero for no critical (throttle) limit; 0 0 ref P P ref _ if MCP scaled, ratios to MCP values kept constant ’MRP’ ’MCP’ _ Reference Engine Rating: SLS, static engine rating: match rating designation in ratings encompass mixture settings and supercharger speeds Pmep Pcrit number of ratings (maximum rating designations MCP rating number power ( specific fuel consumption ( fuel-air ratio ( specific jet thrust ( mean effective pressure limit ( critical (throttle) limit ( reference engine speed ( specification engine speed ( power ( reference engine speed ( critical power ( mechanical limit of power ( Parameters Engine Ratings Reference Derived ratios + + + + + + + + + + + + + int c*12 int real real real real real real real real real real real real ref ref(nratemax) ref(nratemax) _ _ ref(nratemax) ref(nratemax) _ ref(nratemax) _ _ ref(nratemax) ref(nratemax) _ _ _ Structure: RecipModel nrate rating(nratemax) krateC P0 sfc0 F0 SF0 Pmep Pcrit N0 Nspec rP0(nratemax) rN0(nratemax) rcrit(nratemax) rmep(nratemax) 0 1 0 1 1. 1. 1. 0. 1. 0. 0. 0. 1. 0. 0. 0.

0.2 0.3 0.1 0.5 3.0 ) f nengrmax X s ) o X /P X q o 0 P − /P X ) q o vs /P 0 1 2 3 (2 P q 0 1 2 3 X / ) P vs f f q f f q f f q f f q F q F q F q F q − s /F K K K K K K K K X ˙ f f q (2 w f f q req / / K f X F pN pθ crit req X (1 + ram ˙ w − − X X X K p constant constant constant constant number of values (maximum power ratio factor constant constant constant constant K model (1 polynomial, 2 piecewise linear) polynomial piecewise linear exponent model (1 polynomial, 2 piecewise linear) polynomial engine size (0 scaled, 1 fixed) specific output exponent mean piston speed exponent specific fuel consumption exponent exponent exponent factor constant exponent exponent exponent fuel flow, fuel-air ratio, Scaling Derived scaling Power Available Performance at Power Required + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + + int real real real real real real real real real real int real real real real int real real real int real real real real Kffq KFq _ _ size _ Structure: RecipModel FIX Xo Xs Xf Xsfc XN Kp(nratemax) Kram(nratemax) XpN(nratemax) Xpt(nratemax) Xcrit(nratemax) MODEL Kffq0(nratemax) Kffq1(nratemax) Kffq2(nratemax) Kffq3(nratemax) Nffq(nratemax) Pffq(nengrmax,nratemax) Kffq(nengrmax,nratemax) Xffq(nratemax) MODEL KFq0(nratemax) KFq1(nratemax) KFq2(nratemax) KFq3(nratemax) 0. 1.

) (installed thrust loss) nengrmax g /F G F = f gr /P K q P F q F q f gr K X K number of values (maximum power ratio factor piecewise linear exponent constant installed net jet thrust, + + + + + + + int real real real real Structure: RecipModel NFq(nratemax) PFq(nengrmax,nratemax) KFq(nengrmax,nratemax) XFq(nratemax) Kfgr(nratemax) 0. 0. 0. 0. 0.

0.2 ’Comp’ Default ’Default’ ) ˙ m T / = ST ) comp X P 2comp K + P input 1comp K + EngineGroup of 0comp ); gross thrust = specific thrust * mass flow ( K ˙ m = P/ engine _ W = SP IDENT model ( C P vs ) ) comp ” = MCP P C ( W W comp 0comp 1comp 2comp X K K K constant constant constant exponent ” = SLS static; “ compressor identification: used by “ mass flow = power / specific power ( compressor model can be used by more than one engine group, so all parameters fixed title notes identification model (0 fixed, 1 compressor weight (fixed) compressor weight, parameters Compressor Model compressor model number Weight Custom Weight Model Description + + + + + + + + + + + + + + Type c*100 c*1000 c*16 int int real real real real real real comp(8) _ weight _ comp comp comp _ _ _ comp _ Chapter 77 Structure: CompressorModel Variable title notes ident kCompressorModel MODEL Wcomp Kwt0 Kwt1 Kwt2 Xwt WtParam 1. 1. 1. 1.

’MCP’ C /P q P vs FltState /ST C ˙ req m ˙ m / / ST ) a req C ˙ ) ˙ ) m 0 /SP ) m a C /P spec ) ˙ R SP m N C nratemax mech mech ) P P C ST = C /SP ) g R 0 0 0 1 2 3 F ) SP SP spa mf a mf q stq C mf q mf q mf q mf q X X X X K K K K /P ) R 0 0 P P if MCP scaled, ratios to MCP values kept constant exponent exponent constant constant constant constant exponent exponent Reference Compressor Rating: SLS, static compressor rating: match rating designation in number of ratings (maximum rating designations MCP rating number power ( specific power ( mechanical limit of power ( specific jet thrust ( specification compressor speed ( power ( specific power ( mechanical limit of power ( referred specific power available, referred mass flow at power available, referred mass flow at power required, referred specific thrust at power required, Parameters Compressor Ratings Reference Derived ratios Power Available Performance at Power Required + + + + + + + + + + + + + + + + + + + + + + + + int c*12 int real real real real real real real real real real real real real real real real ref(nratemax) ref _ ref _ _ ref(nratemax) _ ref(nratemax) _ Structure: CompressorModel nrate rating(nratemax) krateC P0 SP0 Pmech ST0C Nspec rP0(nratemax) rSP0(nratemax) rPmech(nratemax) Xspa Xmfa Kmfq0 Kmfq1 Kmfq2 Kmfq3 Xmfq Xstq 2 0 0. 0. 0. 0. 0. 0.

’Motor’ Default ’Default’ ) motor X P factor) 2motor K Q/W + P input , 2 low 1motor K Q/W + EngineGroup ; 1 high 0motor of K = ) Q/W ) engine Q W _ ( W model , 2 IDENT ) model ( P C ( peak P Q W vs vs ” = MCP C motor motor motor 0motor 1motor 2motor W W X K K K constant constant constant exponent torque-to-weight design (0 only high ” = SLS static; “ motor identification: used by “ motor model can be used by more than one engine group, so all parameters fixed title notes identification NASA model (0 fixed, 1 motor weight (fixed) motor weight, motor weight, parameters Motor Model motor model number Weight Custom Weight Model Description + + + + + + + + + + + + + + + + Type c*100 c*1000 c*16 int int real real real real real int real motor(8) _ weight _ design motor motor motor _ _ _ _ motor _ Chapter 78 Structure: MotorModel Variable title notes ident kMotorModel MODEL Wmotor Kwt0 Kwt1 Kwt2 Xwt KIND WtParam 1 2 2 0. 0.

86.

1.00 0.00 0.00 1.00 1.00 1.00 0.00 ’MCP’ ) j n i t ij C =0 FltState 3 j ) ) ∑ ij C ) C 0 C =0 3 i sfc = /P ∑ ) R ) ) loss nratemax peak peak ) eng f spec ) P P P N eng eng ) eng P P P eng ) = ˙ Closs(i+1,j+1) P w loss ) ˙ m/ P C /P loss f ) R 0 0 P P if MCP scaled, ratios to MCP values kept constant reference efficiency (at power loss (fraction loss coefficients factor controller efficiency Reference Motor Rating: SLS, static motor rating: match rating designation in number of ratings (maximum rating designations MCP rating number power ( mechanical limit of power ( specification motor speed ( power ( mechanical limit of power ( kind (1 fixed, 2 function power, 3 map) fixed or function power efficiency map ( specific fuel consumption at MCP ( mass flow ratio ( efficiency kind (1 fixed, 2 function power) reference efficiency (at power loss (fraction Parameters Motor Ratings Reference Derived ratios Performance Motor/Generator Efficiency Fuel Cell + + + + + + + + + + + + + + + + + + + + + + + + int c*12 int real real real real real int real real real real real real real int real real cell _ ref(nratemax) ref eff eff _ _ _ _ motor cell motor cont cell ref(nratemax) _ _ _ _ _ _ Structure: MotorModel nrate rating(nratemax) krateC P0 Ppeak Nspec rP0(nratemax) rPpeak(nratemax) KIND eta loss Closs(4,4) floss eta sfc0C Kmf KIND eta loss 0. 0.

= 0.

KNbase = scaling KNspec peak P ) affects N s K base ) N N b K used by efficiency map; spec N for no variation of motor speeds with scale, use specification motor speed variation ( base motor speed variation ( Scaling + + + real real Structure: MotorModel KNspec KNbase 0. 0.

0.2 ’Jet’ Default ’Default’ ) w/T sfc = ˙ ) jet X T 2jet K + T input q req T ); fuel flow = specific fuel consumption * thrust ( 1jet T ˙ m K T / + JetGroup = 0jet of K ST jet = _ , thrust required ) W ) a , thrust required T T av ( IDENT T W required, but not used; weight, ratings, technology, and scaling variables not used and parameters for thrust available and performance at thrust required model ( ref ” = MCT C ref _ _ C T ST0 vs sfc0C 0jet 1jet jet and K K W ref _ only use T0 constant constant ” = SLS static; “ jet identification: used by installed: thrust available uninstalled: thrust available “ mass flow = thrust / specific thrust ( jet model can be used by more than one jet group, so all parameters fixed as model for reaction drive of convertible engine: title notes identification RPJEM model (0 fixed, 1 jet weight (fixed) jet weight, Jet Model jet model number Weight Description + + + + + + + + + + Type c*100 c*1000 c*16 int int real real real weight _ jet jet _ _ Chapter 79 Structure: JetModel Variable title notes ident kJetModel MODEL Wjet Kwt0 Kwt1 1 0 0. 0. 0. 0. 0. 0. 0. 0.

’MCT’ ) ref _ ) sfc0C ref(MCT) (0. for _ ) ) FltState C ) 0 C ST0 tech ) sfc /T sfc lim ) R lim ˙ sfc m (0. for ( nratemax mech mech ) T T C sfc tech ST lim and /ST ST ) R 0 0 ST ) jet ST ST 2jet C X 0 K /T ) R 0 0 T T if MCT scaled, ratios to MCT values kept constant constant exponent Reference Jet Rating: SLS, static jet rating: match rating designation in parameters number of ratings (maximum rating designations MCT rating number thrust ( specific thrust ( mechanical limit of thrust ( specific fuel consumption at MCT ( thrust ( specific thrust ( mechanical limit of thrust ( specific thrust at MCT specific fuel consumption at MCT engine size (0 scaled, 1 fixed) mass flow at limit specific thrust limit specific fuel consumption limit Custom Weight Model Parameters Jet Ratings Reference Derived ratios Technology Scaling + + + + + + + + + + + + + + + + + + + + + real real real int c*12 int real real real real real real real real real int real real real jet(8) _ ref(nratemax) _ tech limit jet ref tech limit _ _ _ _ _ _ jet ref(nratemax) size limit _ _ ref(nratemax) _ _ _ Structure: JetModel Kwt2 Xwt WtParam nrate rating(nratemax) krateC T0 ST0 Tmech sfc0C rT0(nratemax) rST0(nratemax) rTmech(nratemax) ST0C sfc0C FIX MF ST0C sfc0C 0. 1. 1. 0. 1. 0. 1. 1. 1.

tech _ ) ref/ST0C _ tech _ T0C = sfc0C tech = ˙ m C , limit=0.

_ sfc tech _ MF and or C C sfc0C ref C C ˙ , _ 0 m tech /T ˙ _ q m /T vs T size=1 q tech _ _ T vs C vs ST0C C FIX vs C sfc = limit C ST0C 0 ˙ tech=sfc0C _ 0 ST m _ ˙ C ˙ m / w / / MF a , req ST ˙ sfc0C req ˙ m /ST m ˙ , a w limit _ ST with scale, use sfc0C sfc , ref(MCT) _ and (otherwise get limit _ lim ST (0, 1, or 1/2) ˙ m 0 1 2 ST0C tech=ST0 < sta mf a f f q mf q _ f f q f f q f f q mf q X X X X functions are defined by values K K K K tech ST0C ˙ sfc 0 1 m 0 1 st st sf c sf c and limits and thrust limit K K K K exponent exponent constant constant constant exponent constant exponent ST defaults require for no variation of specific thrust available (SLS static, MCT), specific fuel consumption (SLS static, MCT), referred specific thrust available, referred mass flow at thrust available, referred fuel flow at thrust required, referred mass flow at thrust required, Derived scaling Turbofan bypass ratio (0. for turbojet) Thrust Available Performance at Thrust Required + + + + + + + + + + + + + + + real real real real real real real real real real real real real real limit _ Structure: JetModel T0C Kst0 Kst1 Ksfc0 Ksfc1 bypass Xsta Xmfa Kffq0 Kffq1 Kffq2 Xffq Kmfq Xmfq 0. 0. 0. 0. 0. 0.

’Cell’ Default ’Default’ ) cell X P 2cell input K + P 1cell K ChargerGroup + of 0cell K = charge _ W IDENT model ( C P ) ) vs ” = MCP P C ( cell W W cell 0cell 1cell 2cell X K K K constant constant constant exponent ” = SLS static; “ fuel cell identification: used by “ fuel cell model can be used by more than one charger group, so all parameters fixed title notes identification model (0 fixed, 1 fuel cell weight (fixed) fuel cell weight, parameters Fuel Cell Model fuel cell model number Weight Custom Weight Model Description + + + + + + + + + + + + + + Type c*100 c*1000 c*16 int int real real real real real real fuelcell(8) _ weight _ cell cell cell _ _ _ cell _ Chapter 80 Structure: FuelCellModel Variable title notes ident kFuelCellModel MODEL Wcell Kwt0 Kwt1 Kwt2 Xwt WtParam 1 2 0. 0.

86.

1.00 0.00 ’MCP’ FltState ) C ) sfc nratemax ) ) chrg P chrg ) ˙ P w ) ˙ m/ C /P ) R 0 0 P P if MCP scaled, ratios to MCP values kept constant Reference Fuel Cell Rating: SLS, static fuel cell rating: match rating designation in number of ratings (maximum rating designations MCP rating number power ( specific fuel consumption at MCP ( power ( mass flow ratio ( kind (1 fixed, 2 function power) reference efficiency (at power loss (fraction Parameters Fuel Cell Ratings Reference Derived ratios Performance Efficiency + + + + + + + + + + + + + int c*12 int real real real real int real real eff ref _ _ cell cell ref(nratemax) _ _ _ Structure: FuelCellModel nrate rating(nratemax) krateC P0 sfc0C rP0(nratemax) Kmf KIND eta loss 1 1 0. 0.

’Cell’ ’MCP’ Default ’Default’ input ChargerGroup of charge _ ) IDENT nratemax ) ) ” = MCP ) A C ( W ” = SLS static; “ solar cell identification: used by “ solar cell model can be used by more than one charge group, so all parameters fixed title notes identification model (0 fixed, 1 solar cell weight (fixed) weight density (kg/m parameters number of ratings (maximum rating designations MCP rating number Solar Cell Model solar cell model number Weight Custom Weight Model Parameters Solar Cell Ratings Description + + + + + + + + + + + + + + Type c*100 c*1000 c*16 int int real real real int c*12 int solarcell(8) _ weight _ Chapter 81 Structure: SolarCellModel Variable title notes ident kSolarCellModel MODEL Wsolar ssolar WtParam nrate rating(nratemax) krateC 0.

1.00 0.00 FltState ) ) chrg P chrg ) P ) C /P ) R 0 0 P P if MCP scaled, ratios to MCP values kept constant Reference Solar Cell Rating: SLS, static solar cell rating: match rating designation in power ( power ( power density (W/m kind (1 fixed, 2 function power) reference efficiency (at power loss (fraction Reference Derived ratios Performance Efficiency + + + + + + + + real real real int real real eff _ cell cell ref(nratemax) _ _ _ Structure: SolarCellModel P0 rP0(nratemax) esolar KIND eta loss 1 2 4.

4.2 20.

1.00 0.00 Default ’Default’ ’Battery’ dE _ Vref )/ input / xmdb/Econv FuelTank of battery (1/hr) _ mbd (1/hr) x IDENT ) max ref ) P CC ref x P ref (A-hr) from energy capacity (MJ); ( (hp or kW) from energy capacity (MJ); V C P reference efficiency (at power loss (fraction battery identification: used by battery model can be used by more than one fuel tank system, so all parameters fixed title notes identification model (1 equivalent circuit, 2 lithium-ion) reference voltage maximum burst discharge current maximum charge current charge capacity power capacity kind (1 fixed, 2 function power) discharge Battery Model battery model number Performance Derived performance Equivalent Circuit Model Description + + + + + + + + + + + + + + Type c*100 c*1000 c*16 int int real real real real real int real real battery _ eff _ dischrg dischrg _ _ Chapter 82 Structure: BatteryModel Variable title notes ident kBatteryModel MODEL Vref xmbd xCCmax CfromE PfromE KIND eta loss 0. 0.

0.6 1.0 20. 0.1 1.0 0.1 0.2 1.00 0.00 0.01 0.05 0.005 0.000005 nFV=0 ) ) ) d cV ( ref k ) V (fraction component power) V F d ref is capacity) ref d f T C V V C c f = crit f = f d mbd CR/V = ) V o (deg C) x = V c (fraction) ref ) dI dT V V mbd ref V T P d k k F x k T σ ref k P = 0.,.1,.2,.3,.4,.5,.6,.7,.8,.9,.91,.92,.93,.94,.95,.96,.97,.98,.99,1.,1.01,1.02 = 1.,.97,.95,.93,.915,.90,.89,.88,.87,.85,.847,.842,.835,.826,.815,.8,.78,.75,.7,.6,.4,0.

V DoD FV DoD FV number of points (maximum 40) depth-of-discharge F internal resistance depth-of-discharge voltage increment depth-of-discharge reference efficiency (at power loss (fraction critical voltage factor ( nominal discharge voltage ( open circuit voltage ratio ( reference temperature temperature control power loss current influence on discharge voltage temperature influence on discharge voltage nominal charge voltage ( CC phase starting voltage decrement CV phase parameter open circuit voltage ratio: monotonically decreasing; default used if default default charge discharge charge reversed reversed Lithium-Ion Model Derived lithium-ion discharge + + + + + + + + + + + + + + + + + + + + + + + real real real real int real real real real real real real real real real real real real chrg chrg _ _ Structure: BatteryModel eta loss fcrit fd nFV DoD(40) FV(40) Tref fTC R kdI kVT kdT fc kcV ks DoDrev(40) FVrev(40) 0 1 ’ ’ Default ) ) scale kScale _ , scale KIND ) _ geom KIND _ geom=2 _ INPUT WL / , or position of identified component INPUT BL / Ref _ SL XX in Location override global scale _ in Location override ): divided by reference length ( ): dimensional KIND geom _ FIX ; input dimensional geom = 2 + up; from reference point (or combination) to override _ geom = 1 _ kRef ’z’ , ZoL , ’y’ Ref INPUT , _ INPUT ’x’ + right, = KIND YoL geom _ + aft, FIX stationline + aft, buttline + right, waterline + up; arbitary origin; units = ft or m XoL option to fix some geometry ( option to specify reference length ( component reference must be fixed input stationline buttline waterline x/L y/L z/L kind (0 global, 1 rotor radius, 2 wing span, 3 fuselage length) identification (component number) fixed geometry input ( scaled geometry input ( Reference point: Locations can be calculated from other parameters (configuration specific) fixed (dimensional, arbitrary origin) scaled (based on reference length, from reference point) reference length Fixed input: Geometry: Location for each component input Location Description + + + + + + + + + + + + + + Type c*8 real real real real real real int int scale _ geom _ Chapter 83 Structure: Location Variable FIX SL BL WL XoL YoL ZoL KIND kScale XoLloc and SLloc (1 fixed; 2 scaled) (1 rotor radius, 2 wing span, 3 fuselage length) geom (0 calculated, 1 fixed, 2 scaled) _ scale _ FIX geom (component number) _ and KIND FIX SLloc and kScale XoLloc geom and _ and scale ; calc _ ; calc geom _ SLloc XoLloc calculation depends on component/configuraton; calc from from Aircraft%INPUT x x x Aircraft%KIND Aircraft%kScale : : : Aircraft%INPUT x y z x y z all fixed (0 not, some scaled or calculated) stationline buttline waterline x/L y/L z/L from from reference length (+ forward) (+ right) (+ down) FIX = 0 FIX = 1 FIX = 2 x y z input, from from fixed (dimensional, arbitrary origin) scaled (based on reference length, from reference point) reference length Derived Geometry (dimensional, body axes, relative reference point) int int int int int int int real real real real real real int int real real real real x y z _ _ _ loc _ geom geom geom loc _ _ _ scale _ _ x y z _ _ _ Structure: Location INPUT INPUT INPUT FIX FIX FIX isFixed SLloc BLloc WLloc XoLloc YoLloc ZoLloc KIND kScale scale x y z Default fairings (not RP8A) fittings (not RP8A) fold/tilt (not RP8A) basic (not RP8A) inter-rotor shaft (not RP8A) fairing/spinner (not RP8A) blade fold (not RP8A) basic (not RP8A) fold (not RP8A) basic (not RP8A) fold (not RP8A) blades hub & hinge basic structure secondary structure control surfaces blade assembly hub & hinge rotor support structure (not RP8A) duct (not RP8A) horizontal tail (not RP8A) vertical tail (not RP8A) tail rotor (not RP8A) wing group rotor group empennage group STRUCTURE WEIGHT EMPTY Description Type real real real real real real real real real real real real real real real real real real real real real real real real real real real real blade hub basic shaft fair fold supt duct basic fold basic secondary fair fit fold control basic fold _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ blade hub _ _ structure wing wing wing wing wing wing wing rotor rotor rotor rotor rotor rotor rotor rotor rotor tail Htail Htail Htail Vtail Vtail Vtail tailrotor tr tr _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ Chapter 84 Structure: Weight Variable WE W W W W W W W W W W W W W W W W W W W W W W W W W W W rotor supports rotor/fan duct basic (not RP8A) wing & rotor fold/retraction (not RP8A) tail fold/tilt (not RP8A) marinization (not RP8A) pressurization (not RP8A) crashworthiness (not RP8A) basic (not RP8A) retraction (not RP8A) crashworthiness (not RP8A) engine support (not RP8A) engine cowling (not RP8A) pylon support (not RP8A) engine exhaust system accessories (not RP8A) blades (not RP8A) hub & hinge (not RP8A) rotor supports (not RP8A) rotor/fan duct (not RP8A) tanks and support plumbing gear boxes transmission drive fuselage group alighting gear group engine section or nacelle group air induction group engine system propeller/fan installation fuel system drive system PROPULSION GROUP real real real real real real real real real real real real real real real real real real real real real real real real real real real real real real real real real real box xmsn blade hub supt duct basic retract crash tank plumb _ _ engsupt cowling pylon _ _ _ _ _ _ _ basic wingfold tailfold mar press crash _ _ _ _ _ supt duct _ _ _ _ _ _ _ _ tr tr fuselage fus fus fus fus fus fus gear gear gear gear nacelle nac nac nac airind propulsion engsys engine exhaust acc propeller prop prop prop prop fuelsys fuel fuel drive drive drive _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ Structure: Weight W W W W W W W W W W W W W W W W W W W W W W W W W W W W W W W W W W non-boosted (not RP8A) boost mechanisms (not RP8A) non-boosted (not RP8A) boost mechanisms (not RP8A) boosted (not RP8A) non-boosted (not RP8A) boost mechanisms (not RP8A) fixed wing systems rotary wing systems conversion systems rotor shaft rotor brake (not RP8A) clutch (not RP8A) gas drive cockpit controls automatic flight control system system controls fixed wing (not RP8A) rotary wing (not RP8A) conversion (not RP8A) equipment (not RP8A) aircraft (not RP8A) anti-icing (not RP8A) armament provisions (not RP8A) armor (not RP8A) flight controls group auxiliary power group instruments group hydraulic group pneumatic group electrical group avionics group (mission equipment) armament group SYSTEMS AND EQUIPMENT real real real real real real real real real real real real real real real real real real real real real real real real real real real real real real real real real real nonboost mech boost nonboost mech nonboost mech rtrsft brake clutch gas aircraft deice _ _ _ _ _ _ _ _ _ _ _ fw rw cv eq _ _ _ _ _ _ cockpit afcs system fw fw fw rw rw rw rw cv cv cv _ _ _ _ _ _ _ _ _ _ _ _ _ drive drive drive drive equip fltcont fc fc fc fc fc fc fc fc fc fc fc fc fc auxpower instrument hydraulic hyd hyd hyd hyd pneumatic electrical elect elect avionics arm armprov armor _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ Structure: Weight W W W W W W W W W W W W W W W W W W W W W W W W W W W W W W W W W W =2 or 3) at any level, ignore child weights if not lowest level k can define weights ( furnishings & equipment group environmental control group anti-icing group load & handling group follows SAWE RP8A Group Weight Statement, except as noted typical only lowest elements of hierarchy specified, others obtained by summation set status flag when define weight when print weight statement, designate all fixed (ie input) quantities VIBRATION (not RP8A) CONTINGENCY crew fluids (oil, unusable fuel) (not RP8A) auxilary fuel tanks other fixed useful load (not RP8A) equpment increment (not RP8A) folding kit (not RP8A) wing extension kit (not RP8A) wing kit (not RP8A) other kit (not RP8A) standard tanks (not RP8A) auxiliary tanks (not RP8A) FIXED USEFUL LOAD PAYLOAD USABLE FUEL scaled weight (sum all K=3 in operating weight) fixed weight (sum all K=2 in operating weight) military features in empty weight OPERATING WEIGHT = weight empty + fixed useful load USEFUL LOAD = fixed useful load + payload + usable fuel GROSS WEIGHT = weight empty + useful load = operating weight + payload + usable fuel real real real real real real real real real real real real real real real real real real real real real real real real real real crew fluid auxtank other equip foldkit extkit wingkit otherkit _ _ _ _ _ _ _ _ _ std aux _ _ furnish environ deice load vib cont fixUL fixUL fixUL fixUL fixUL fixUL fixUL fixUL fixUL fixUL _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ Structure: Weight W W W W W W W W W W W W W W W W Wpayload Wfuel Wfuel Wfuel Wscaled Wfixed Wfeature WO WUL GW K=1 3 3 3 3 3 and 2 2 3 2 3 (child) ∑ 1 1 1 3 3 W= of total as follows 0 0 1 2 3 Kt = Ka = 0 = 1 = 2 = 3 Kb Kb Kb Kb =2 or 3 K and some child defined/sum, then , with K=0 W ; put W=K=0 fairings (not RP8A) fittings (not RP8A) fold/tilt (not RP8A) basic (not RP8A) inter-rotor shaft (not RP8A) fairing/spinner (not RP8A) blade fold (not RP8A) basic structure secondary structure control surfaces blade assembly hub & hinge rotor support structure (not RP8A) duct (not RP8A) horizontal tail (not RP8A) set all then fill structure: if addition or increment sums all elements, with status wing group rotor group empennage group usage: STRUCTURE Status (0 none; 1 sum of child; 2 defined, fixed (input); 3 defined, not fixed (scaled, wt eq; or composite)) WEIGHT EMPTY int int int int int int int int int int int int int int int int int int int int basic secondary fair fit fold control blade hub basic shaft fair fold supt duct _ _ _ _ _ _ _ _ _ _ _ _ _ _ structure wing wing wing wing wing wing wing rotor rotor rotor rotor rotor rotor rotor rotor rotor tail Htail _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ Structure: Weight KE K K K K K K K K K K K K K K K K K K K basic (not RP8A) fold (not RP8A) basic (not RP8A) fold (not RP8A) blades hub & hinge rotor supports rotor/fan duct vertical tail (not RP8A) tail rotor (not RP8A) basic (not RP8A) wing & rotor fold/retraction (not RP8A) tail fold/tilt (not RP8A) marinization (not RP8A) pressurization (not RP8A) crashworthiness (not RP8A) basic (not RP8A) retraction (not RP8A) crashworthiness (not RP8A) engine support (not RP8A) engine cowling (not RP8A) pylon support (not RP8A) engine exhaust system accessories (not RP8A) blades (not RP8A) hub & hinge (not RP8A) fuselage group alighting gear group engine section or nacelle group air induction group engine system propeller/fan installation PROPULSION GROUP int int int int int int int int int int int int int int int int int int int int int int int int int int int int int int int int int int basic fold basic fold blade hub basic retract crash _ _ _ _ engsupt cowling pylon _ _ _ _ _ basic wingfold tailfold mar press crash _ _ _ _ _ _ _ _ _ blade hub supt duct _ _ _ _ Htail Htail Vtail Vtail Vtail tailrotor tr tr tr tr fuselage fus fus fus fus fus fus gear gear gear gear nacelle nac nac nac airind propulsion engsys engine exhaust acc propeller prop prop _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ Structure: Weight K K K K K K K K K K K K K K K K K K K K K K K K K K K K K K K K K K non-boosted (not RP8A) boost mechanisms (not RP8A) non-boosted (not RP8A) boost mechanisms (not RP8A) boosted (not RP8A) non-boosted (not RP8A) boost mechanisms (not RP8A) fixed wing systems rotary wing systems conversion systems rotor supports (not RP8A) rotor/fan duct (not RP8A) tanks and support plumbing gear boxes transmission drive rotor shaft rotor brake (not RP8A) clutch (not RP8A) gas drive cockpit controls automatic flight control system system controls fixed wing (not RP8A) rotary wing (not RP8A) conversion (not RP8A) equipment (not RP8A) fuel system drive system flight controls group auxiliary power group instruments group hydraulic group SYSTEMS AND EQUIPMENT int int int int int int int int int int int int int int int int int int int int int int int int int int int int int int int int int int nonboost mech nonboost mech boost nonboost mech box xmsn rtrsft brake clutch gas supt duct _ _ _ _ _ _ _ tank plumb _ _ _ _ _ _ fw rw cv eq _ _ _ _ _ _ _ _ cockpit afcs system fw fw fw rw rw rw rw cv cv cv _ _ _ _ _ _ _ _ _ _ _ _ _ prop prop fuelsys fuel fuel drive drive drive drive drive drive drive equip fltcont fc fc fc fc fc fc fc fc fc fc fc fc fc auxpower instrument hydraulic hyd hyd hyd hyd _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ Structure: Weight K K K K K K K K K K K K K K K K K K K K K K K K K K K K K K K K K K aircraft (not RP8A) anti-icing (not RP8A) armament provisions (not RP8A) armor (not RP8A) pneumatic group electrical group avionics group (mission equipment) armament group furnishings & equipment group environmental control group anti-icing group load & handling group VIBRATION (not RP8A) CONTINGENCY crew fluids (oil, unusable fuel) (not RP8A) auxilary fuel tanks other fixed useful load (not RP8A) equipment increment (not RP8A) folding kit (not RP8A) wing extension kit (not RP8A) wing kit (not RP8A) other kit (not RP8A) standard tanks (not RP8A) auxiliary tanks (not RP8A) FIXED USEFUL LOAD PAYLOAD USABLE FUEL OPERATING WEIGHT = weight empty + fixed useful load USEFUL LOAD = fixed useful load + payload + usable fuel GROSS WEIGHT = weight empty + useful load = operating weight + payload + usable fuel int int int int int int int int int int int int int int int int int int int int int int int int int int int int int int int crew fluid auxtank other equip foldkit extkit wingkit otherkit aircraft deice _ _ _ _ _ _ _ _ _ _ _ std aux _ _ pneumatic electrical elect elect avionics arm armprov armor furnish environ deice load vib cont fixUL fixUL fixUL fixUL fixUL fixUL fixUL fixUL fixUL fixUL _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ Structure: Weight K K K K K K K K K K K K K K K K K K K K K K K K Kpayload Kfuel Kfuel Kfuel KO KUL KGW

NDARC NASA Design and Analysis of Rotorcraft - Input, Appendix 2

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NASA/TP – 2015 - 21 8751

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NASA Design and Analysis of Rotorcraft

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NASA/TP – 2015 - 21 8751

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NASA Design and Analysis of Rotorcraft

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National Aeronautics and

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Ames Research Center

Moffett Field, California 94035 - 1000

March 201 6

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