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AAA 3.6 User Manual

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1. oonccnnncnicnonnconnconcconccnnos 114 4 3 4 Trailing Edge Flap oooncnnnninnnnnnnonnnonnconocnnos Error Bookmark not defined 4 3 5 Krueger Flap ooonnonnnnnnnnnncnncnnonnconnconoconcconocnnos Error Bookmark not defined 436 Sla eatin este Rt eee eee SN Error Bookmark not defined 43 7 Aero abi chs his iets Sais Riis nee eee 115 cI Ae O eee eee 115 4 3 9 Chord Length 2 tsteied ctieitivcintin ii 115 A310 Atoll anal 115 43T Epose denan rn Gh E N ee 115 4 4 Horizontal Tail Geometry iccse eeni eee iea oer E EE SES 115 4 4 1 A AN 116 4 42 Cranked Horizontal Tall oonnnnnnnnninnnonnconononcnnncnnnnnnnnannnnn conc cono co noc nocnnccnnos 116 4 43 Volume Coefficients neinni a A e conan cn necn cocoa 116 AAA A O 116 4 45 SA NS 116 446 Airfoil A Meese bakes ee nile E 116 AAT EXPOSE tdi iia 116 4 5 Vertical E A AR 116 A A caesiectctssescetsseessaatep EE rE erii h 117 4 32 Cranked Vertical Talni nnn E E S 117 Table of Contents VI 4 6 4 7 4 8 4 9 4 10 4 11 4 12 4 13 4 14 4 15 4 16 4 17 4 18 4 19 4 20 4 21 4 22 4 23 4 5 3 Vols Cosina deidad 117 434 Rudder abri a SAU ANS 117 ADS Chord Len aths rta ae e ene ae no rE EnS EERE EREN 117 AO ATOI ori 118 AS O NO 118 Canard Geometry senene a E a a S N E ES 118 4 6 1 Straight Tapered oooonncconccnoncnnenonornncrnenonorno nono a A r EE EEE E T E 118 4 6 2 Cranked Canard inciso raone dni aa Ea sons besa EErEE aE aiie
2. gyration method using airplane moment of inertia methods 1 4 1 Weight Fractions This option allows the user to calculate component weights by using the weight fractions method The data can be provided by the user or the AAA database of comparable airplanes can be used After selecting an airplane category the user may select airplanes available in that particular II 66 Aerodynamics category One or more airplanes can be selected Because the Class I component weight estimation is based on the selected aircraft the user should choose the most comparable aircraft To deselect an airplane select the airplane from the current selection list and remove it The Class I Weight Fractions component contains the options Add Airplane Select Airplane Fractions and Weights which have the following functions e Add Airplane e Select Airplane e Fractions e Weights Aerodynamics Allows the user to add new airplanes to the already available airplanes appearing in the table The user has to supply the weight fraction data This data will automatically be added to the Select Airplane menu under the Airplane Category labeled User Defined Allows the user to select existing airplanes from the program database for use in the Class I component weight estimation All the available airplanes in all categories can be displayed The user may select any combination of these to use in the weight estimation by adding them to the curre
3. Propulsion Engines gt pNacelles Propulsion Yes C No Number 0 C Yes C No C Jet C Propeller C Propfan Ducted Fan Turboprop Engine Location Nacelle Location Wing Fuselage Other Wing Fuselage Other Cc c c a E Fuel Tank Aspiration C Integral C Normally Aspirated C Self Sealing Bladder C Supercharged C Non Self Sealing Turbocharged Type of Inlet Prop Type C Plenum C Composite C Straight Through C Metal Prop Pitch C Fixed Variable 1 2 3 4 5 6 7 8 9 elele tele eeel e YD eee e e elelee e 1010 o ala uni elelee 110 eleele oT ee Ra e SNe NCH iG ee Ned elie o gt Engine Air Flow Location C Not Buried no airflow from fuselage base Number of Fans C Buried in Fuselage airflow from fuselage base X Cancel 7 Help Figures 2 17e Powerplant Selection Dialog Box for a Ducted Fan Piston Propulsion Engines Nacelles Propulsion Engine Type Yes No Number 0 C Yes C No C Jet C Piston C Propeller Ducted Fan C Turboprop Engine Location Nacelle Location Wing Fuselage Other Wing Fuselage Other E E C C C C Fuel Tank C Integral C Self Sealing Bladder C Non Self Sealing Type of Inlet Prop Type C Plenum C Composite C Straight Through C Metal Prop Pitch C Fixed C Variable oon on a uni 1 OL BY 0 YL 11211 0116211 e 1111121101160 11 0 o wN omo an ES YL OL italie eiee ie Tied ey e teite See te tea Te eTe _ o a Engine Air Flow Loc
4. e Elevon Speed For the Speed to Stabilizer Transfer Function e Elevon Angle of Attack For the Angle of Attack to Stabilizer Transfer Function e Elevon Pitch Angle For the Pitch Angle to Stabilizer Transfer Function e Aileron Sideslip Angle For the Sideslip Angle to Aileron Transfer Function e Aileron Bank Angle For the Bank Angle to Aileron Transfer Function e Aileron Heading Angle For the Heading Angle to Aileron Transfer Function e Rudder Sideslip Angle For the Sideslip Angle to Rudder Transfer Function e Rudder Bank Angle For the Bank Angle to Rudder Transfer Function e Rudder Heading Angle For the Heading Angle to Rudder Transfer Function e Human Pilot For the Human Pilot Transfer Function e Time Delay For the Time Delay Transfer Function e User Defined For a Transfer Function defined by the user The transfer functions can be selected from the standard airplane transfer functions or defined by the user If the longitudinal and lateral directional stability derivatives of the airplane are known the user may use the Dynamics module prior to using the Control analysis module to generate the longitudinal and lateral directional dynamic transfer functions of the airplane These transfer functions are transferred into the Control analysis module and can only be generated in the Dynamics module The User Defined option allows the user to define a desired transfer function An example of when this option might be used is
5. 2 3 1 1 Maximum Lift Coefficient The following options are presented once the CZ nax submodule is selected e Airfoil Climax e Surface C7 The methodology used to calculate the lifting surface root and tip airfoil maximum lift Aerodynamics T 81 coefficient is described in Chapter 7 of Reference 2 To calculate the surface root or tip airfoil maximum lift coefficient select the Airfoil Cl nax button An input output window appears containing the surface root and tip airfoil type inputs Each is a pull down list with six different symmetric or cambered airfoil types as follows e NACA 4 amp 5 Digit Symm For NACA four and five digit symmetrical series airfoils e NACA 6 Digit Symmetric For NACA six digit series symmetrical airfoils e MS 1 Cambered For NASA medium speed airfoils e NACA 4 amp 5 Digit Camb For NACA four and five digit series cambered airfoils e NACA 6 Digit Cambered For NACA six digit series cambered airfoils e Other User Defined For a user defined airfoil To calculate the lifting surface planform maximum lift coefficient select the Surface C Diak button An input output window appears showing the surface planform lift coefficient input and output parameters The user can supply the data of other airfoils by choosing user defined and defining the root and tip airfoil maximum lift coefficients in the input menu for the surface planform maximun lift coefficient 2 3 1 2 Lift Distribution The
6. Category A B or C C A Airto Air Combat C B Loiter C Ground Attack B In Flight Refueling Tanker C A Weapon Delivery Launch C B Descent C amp Aerial Recovery C B Emergency Descent C amp Reconnaissance B Emergency Deceleration C A In Flight Refueling Receiver C B Aerial Delivery C amp Terrain Following C Take off C A Antisubmarine Search C C Catapult Take off C A Close Formation Flying C Approach C B Climb C C Wave Off Go Around C B Cruise C Landing Flight Condition Flying Qualities Category Notes X Cancel 7 Help Figure 2 11b Flight Condition Dialog Box Contd I 17 E Flight Condition Definition A Flight Phase Name Flight Condition 1 El L New j Edit ti Delete Gi Move EX Copy Flight Condition Flying Qualities Category Notes f X Cancel 7 Help Figure 2 11c Flight Condition Dialog Box Contd The options available to the user in the Flight Condition Definition dialog box are described as follows Flight Phase Name The user can select the name of the flight phase for which the present analysis is to be performed i e take off climb cruise etc The defined phases appear in the drop down list to be selected by the user Only one flight phase can be analyzed at one time The program can handle up to 95 different flight conditions New The user can define a new fligh
7. Pitch Angle For the Pitch Angle to Stabilizer Transfer Function e Canardvator Speed For the Speed to Canardvator Transfer Function e Canardvator Angle of Attack For the Angle of Attack to Canardvator Transfer Function e Canardvator Pitch Angle For the Pitch Angle to Canardvator Transfer Function e Canard Speed For the Speed to Canard Transfer Function e Canard Angle of Attack For the Angle of Attack to Canard Transfer Function e Canard Pitch Angle For the Pitch Angle to Canard Transfer Function e V Tail Speed For the Speed to V Tail Transfer Function e V Tail Angle of Attack For the Angle of Attack to V Tail Transfer Function e V Tail Pitch Angle For the Pitch Angle to V Tail Transfer Function e Ruddervator Speed For the Speed to Ruddervator Transfer Function H 164 Dynamics e Ruddervator Angle of Attack For the Angle of Attack to Ruddervator Transfer Function e Ruddervator Pitch Angle For the Pitch Angle to Ruddervator Transfer Function e Ruddervator Sideslip Angle For the Sideslip Angle to Ruddervator Transfer Function e Ruddervator Bank Angle For the Bank Angle to Ruddervator Transfer Function e Ruddervator Heading Angle For the Heading Angle to Ruddervator Transfer Function e Flying Wing Speed For the Speed to Elevator Transfer Function e Flying Wing Angle of Attack For the Angle of Attack to Elevator Transfer Function e Flying Wing Pitch Angle For the Pitch Angle to Elevator Transfer Function
8. The user can define the basic configuration of the airplane using this dialog box For vertical tail nacelles stores tailbooms floats pylons and ventral fins the number of each device can also be defined using this dialog PART I I 25 box 7 Airplane Configuration Wing C Yes No Vertical Tail C Yes No Horizontal Tail C Yes No Stores C Yes No Tailbooms C Yes No Tail Canard Floats C Yes No C Yes No C Yes No Windshield Spoiler Pylons C Yes No C Plug C Yes No Canopy C Flap Ventral Fins C Yes C No C Yes No C Slot Speed Brake Dorsal Fin C Yes No C None C Yes No sgt x Cancel 7 Help Figures 2 16 Airplane Configuration Dialog Box Propulsion Selection Dialog Box When the Engine button is selected the Powerplant Selection dialog box is displayed Figure 2 17 shows the Powerplant Selection dialog box Different options will show up on the dialog box depending on the engine selected Descriptions of the dialog box functions follow I 26 PARTI Propulsion Engines Nacelles Yes No Number jo Yes C No Number jo 3 Jet C Propeller C Ducted Fan Engine Location Nacelle Location Wing Fuselage Other Wing Fuselage Other T C T C E C Fuel Tank C Integral C Self Sealing Bladder C Non Self Sealing Type of Inlet C Plenum C Straight Through 1 2 3 4 5 6 7 8 9 Welte e ee ae gee e ee eE Delw eee ele o wN omn a
9. These four submodules are 6 4 1 Sideslip For estimation of the angle of sideslip derivatives 6 4 2 Sideslip Rate For estimation of the rate of angle of sideslip derivatives 6 4 3 Roll Rate For estimation of the roll rate derivatives 6 4 4 Yaw Rate For estimation of the yaw rate derivatives For certain submodules the user must define the longitudinal control surface in the Airplane Configuration dialog box See Section 2 2 3 Figure 2 16 H 134 Dynamics In addition the user must also define the vertical tail and the number of vertical tail s in the Airplane Configuration dialog box See Section 2 2 3 Figure 2 16 6 4 1 Angle of Sideslip Derivatives The methodology used to calculate the sideslip related derivatives can be found in Section 10 2 4 of Reference 7 Once the Sideslip button has been selected a new menu appears with the following angle of sideslip related derivative options Sy This submodule can be used to estimate the side force coefficient due to sideslip derivative Cts This submodule can be used to estimate the rolling moment coefficient due to sideslip derivative Eng This submodule can be used to estimate the yawing moment coefficient due to sideslip derivative Ors This submodule can be used to estimate thrust side force coefficient due to sideslip Derivative eC This submodule can be used to estimate the thrust yawing moment coefficient due to sideslip derivative 6 4 2
10. This take off weight is used in the first step of the take off weight iteration process and is defined as Wro Wstructure Wop Wfix Wpy Wet Wr We Wifo MyoWro Where efuel Wifo Wr 1 M p 1 M r Wro Aerodynamics Wro Airplane take off weight in Ib or N W structure Airplane total structure weight in 1b or N Wop Airplane total powerplant weight in Ib or N Wfix Airplane total fixed equipment weight in Ib or N Wer Airplane payload weight in Ib or N includes weight of crew WPlexp Expended payload weight in Ib or N Wr Airplane mission fuel weight in Ib or N Wr fiel Refueled fuel weight in lb or N Wifo Weight of trapped fuel and oil in Ib or N My Mission fuel fraction Pe Reserve fuel weight as fraction of fuel weight used in the mission in Mo Trapped fuel and oil weight as a fraction of take off weight in Instead of reentering the calculated take off weight in the individual component weight calculations and estimating the take off weight within a given accuracy the user can perform the iteration in the weight iteration submodule as discussed in Section 1 5 5 1 5 6 Weight Iteration Process The Weight Iteration submodule allows the user to perform the take off weight iteration with an accuracy of 0 05 lb Once all required input is defined the iteration table will be displayed If not all the input parameters are defined a message will be displayed indicating
11. V n Input window the user can use the Plot option A detailed description of the standard plot options is described in Section 2 1 4 of Part II of the manual 8 4 Structural Loads After invoking the Structural button the following options will be available to calculate the loads acting on a structural component 8 4 1 Fuselage For estimation of the structural loads on the fuselage 8 4 1 Wing For estimation of the structural loads on the wing 8 4 1 Horizontal Tail For estimation of the structural loads on the horizontal tail 8 4 1 Canard For estimation of the structural loads on the canard 8 4 1 Vertical Tail For estimation of the structural loads on the vertical tail 8 4 1 V Tail For estimation of the structural loads on the v tail 8 4 2 Load Factors For estimation of the accelerations and rates acting on the airplane Before the total internal loads for any structural component can be calculated the linear accelerations angular accelerations and angular rates must be calculated using this option H 172 Loads 8 4 1 Fuselage Wing Horizontal Tail Canard Vertical Tail V Tail When the Fuselage Wing Horizontal Tail Canard or Vertical Tail option is chosen the following options will be available for calculation of the respective structural component loads e Conc Weights For calculation of loads due to a mass placed at one concentrated point on the structural component Section 8 4 1 1 e Dist Weights For calculation
12. e seeesseeseseersreeresreerrstrrrsresrenresrerrsseerrnserrenresrentes 103 3 3 7 IN ON 104 3 4 Perform nce AnalySiS sonas dadas 104 3 4 1 Thrust Speed Performance Curve cocooococcnncconnnonnnonnnnnnonoconoconoconccnncnn nono ncnncnnos 105 3 4 2 Power Speed Performance Curve cooocococcnccconnnonnnonnnonnonononoconocn nono nono nonnninncnnss 105 3 4 3 Stall Sperdsocoiiia nina tibia lec isiici 105 3 4 4 Take off Distance vocoocosiitrncoo sicario tease ereen r rE rE E E E 105 Table of Contents A A A 106 A A REO 107 3 4 6 1 Maximum Cruise Speed oooonccncnnoncconoconccononnninncnononononanonnnc noc nocnncnnnos 107 34 632 Range eii la 107 3 4 6 3 Endurance eiir r a EEE rE EEES EEES ES 108 3 4 6 4 Payload Range Diagram eesssesseeeessereeesesrrereseeerssrerrsserrresesreeresee 108 34 7 Dive Descent sonr nt A e Laine eben ae ee 108 3 48 Maneuvers EE EET 109 AI E E sa 110 3 4 10 Landing Distances minene iere a AE E Ee ERE AREN E ENES 110 30 VExport to APP iia 110 4 Geometry Module iss sess cesses iaa cit diia da aiii 111 AY General Deseriptroniiss ssc s sssssscoesssthesstbcessesdsvesststanestits vehostebapesces dgsvansebigesdes ssebsacebarags 111 4 2 Airplane Geometry Categories eee ee eeeeeeeeeeeeeeeseceseeesecesecssecsaecnaecsaecaaeenevenevees 111 4 3 Wing Geometry pei lia 113 4 3 1 A cave eeren EEEE i 113 43 2 Cranked WINE isos rta ias 113 4 3 3 Fuel Volume Class I Class Il and Cranked WiN8
13. in this module are discussed in Section 4 5 For calculation of the canard geometry The options provided in this module are discussed in Section 4 6 For calculation of the v tail geometry The options provided in this module are discussed in Section 4 7 For calculation of the ventral fin geometry The options provided in this module are discussed in Section 4 8 For calculation of the fuselage geometry The options provided in this module are discussed in Section 4 9 For calculation of the nacelle geometry The options provided in this module are discussed in Section 4 10 For calculation of the tailboom geometry The options provided in this module are discussed in Section 4 11 Il 111 e Float e Landing Gear e Canopy e Store e Pylon e Propeller e Airplane 3 View e Angles e Scale e Translate e Airfoil Folder e AeroPack Tr 112 For calculation of the float geometry The options provided in this module are discussed in Section 4 12 For calculation of the landing gear geometry The options provided in this module are discussed in Section 4 13 For calculation of the canopy geometry The options provided in this module are discussed in Section 4 14 For calculation of the store geometry The options provided in this module are discussed in Section 4 15 For calculation of the pylon geometry The options provided in this module are discussed in Section 4 16 For calculation of the propeller geo
14. is first specified An empty table will appear with one column each for the airplane name take off weight and empty weight When all input data are entered and the user selects Calculate this will calculate the intercept A and slope B for a logarithmic relation between take off and empty weight using a least squares method Increasing the number of input parameters adds an empty row at the end of the table which allows for the addition of a data point to be used in calculating A and B The maximum number of rows is 50 This command can also be used to delete table rows from the bottom of the table After creating or loading a table use the Import option in the File menu the data can be plotted with a logarithmic scale A and B must be calculated first The standard plot options are described in Chapter 2 of Part II of this manual If the take off weight and empty weight of a design have been calculated using the Take Off Weight option they will be plotted with the weight table elements and be designated by a Design Point label 1 3 5 Sensitivity This option allows the user to calculate the sensitivity of take off weight to aerodynamic propulsion and mission parameters Once all data from Mission Profile and Take Off Weight are available selecting Calculate will display take off weight sensitivities with respect to payload weight and empty weight The results are shown in a table which will appear with the flight segments and the corre
15. the following rate of angle of attack related derivative options e C D In preliminary design the drag coefficient due to rate of angle of attack derivative is assumed to be negligible and will therefore not be calculated The program automatically sets this derivative equal to zero and no additional menus appear on the screen Dynamics TI 133 e Ci This submodule can be used to estimate the lift coefficient due to rate of angle of attack derivative e Cm This submodule can be used to estimate the pitching moment coefficient due to rate of angle of attack derivative 6 3 5 Pitch Rate Derivatives The methodology used to calculate the pitch rate related derivatives can be found in Section 10 2 7 of Reference 7 Once the Pitch Rate button is selected a new menu appears with the following pitch rate related derivative options e C Dg In preliminary design the drag coefficient due to pitch rate derivative is assumed to be negligible and will therefore not be calculated The program automatically sets this derivative equal to zero and no additional menus appear on the screen e C Es For estimation of the lift coefficient due to pitch rate derivative e Cn For estimation of the pitching moment coefficient due to pitch rate derivative 6 4 Lateral Directional Stability Derivatives Once the Lat Dir Stability option has been selected another menu appears with four lateral directional stability derivative submodules
16. 6 6 Rudder Tab Related Derivatives The methodology used to calculate the rudder tab related derivatives can be found in Section 10 3 8 of Reference 7 Once the Rudder Tab button is selected and the vertical tail H 144 Dynamics configuration has been chosen a new menu appears with the following three rudder tab related derivative options ec ys This submodule can be used to estimate the side force coefficient rt due to rudder tab derivative e Cs This submodule can be used to estimate the rolling moment rt coefficient due to rudder tab derivative e Crs This submodule can be used to estimate the yawing moment rt coefficient due to rudder tab derivative 6 6 7 V Tail Related Derivatives Once the V Tail button is selected a new menu appears with the following three v tail related derivative options ec yi This submodule can be used to estimate the sideforce coefficient vee due to v tail incidence derivative e C This submodule can be used to estimate the rolling moment coefficient due to v tail incidence derivative e C This submodule can be used to estimate the yawing moment coefficient due to v tail incidence derivative 6 6 8 Ruddervator Related Derivatives The methodology used to calculate the ruddervator related derivatives can be found in Section 10 3 7 of Reference 7 Once the Ruddervator button is selected a new menu appears with the following three ruddervator related derivative options
17. A logarithmic plot of the entered data is also provided After performing Mission Profile and Take off Weight sensitivities of take off weight for changes in aerodynamic propulsion and mission parameters can be calculated This module calculates the range for different payload scenarios starting with a minimum of 1 pilot and ending with the maximum payload specified in the Take off Weight module Aerodynamics 1 3 1 Mission Profile This option shows the Segment Input and Fuel Fraction Calculation menu This menu consists of four command buttons e New Segment e Delete Segment e Insert Segment e Move Segment Aerodynamics This command is used to enter a new segment of the mission specification A selection can be constructed from the twelve mission segments The newly specified flight segment is added at the end of the previously defined segments After choosing the desired segment an input output window is displayed For some flight segments the Calculate button must be selected to get the fuel fraction For other segments the input will be the fuel fraction The selected mission segment will be displayed in a table together with the fuel fraction for this segment With this command a segment can be deleted from the flight profile The user selects a segment that is to be deleted All data related to this segment will be erased If the user wishes to exit without deleting a segment any menu can be selected to exit t
18. Bar Buttons Table 2 3 Plot Window Command Bar Buttons Read Off Graph Allows the user to read a value from the plot When selected the corresponding X and Y coordinates are displayed when the user positions the cross hairs on the plot by holding down the left mouse button The cross hairs will be set at the instant that the mouse button is released Edit Allows the user to modify the font of all text displayed on the plot window Default Recalculates the axes so that the whole graph will show up on the plot window Export Export input and output data to a text file ASCID or to an Excel Spreadsheet Close Closes the plot window The window minimize button can be used to iconize the window if desired I 14 PART I 2 2 Toolbars The program main window contains five toolbars located above the status bar Figure 2 1 The main toolbar is located on the right and is always visible The four remaining toolbars can be displayed by clicking on the corresponding tab underneath the currently displayed toolbar on the left side of the main window The five toolbars are described in the following subsections 2 2 1 The Main toolbar see Figure 2 1 2 2 2 The File toolbar displayed by clicking on the File tab see Figure 2 1 2 2 3 The Configuration toolbar displayed by clicking on the Configuration tab 2 2 4 The Certification toolbar displayed by clicking on the Certification tab 2 2 5 The Setup toolbar displayed by cli
19. Control submodule has six lateral directional control surface and trim tab derivative submodules These eight submodules are 6 6 1 Aileron For estimation of the aileron derivatives 6 6 2 Aileron Tab For estimation of the aileron tab derivatives 6 6 3 Spoiler For estimation of the spoiler derivatives 6 6 4 Vert Tail For estimation of the vertical tail derivatives 6 6 5 Rudder For estimation of the rudder derivatives 6 6 6 Rudder Tab For estimation of the rudder tab derivatives 6 6 7 V Tail For estimation of the v tail derivatives 6 6 8 Ruddervator For estimation of the ruddervator derivatives 6 6 9 Ruddervator Tab For estimation of the ruddervator tab derivatives 6 6 10 Diff Stabilizer For estimation of the differential stabilizer derivatives 6 6 1 Aileron Related Derivatives The methodology used to calculate the aileron related derivatives can be found in Section 10 3 5 of Reference 7 Once the Aileron button is selected a new menu appears with the following three aileron related derivative options ec 5 In preliminary design the side force coefficient due to aileron a derivative is assumed to be negligible and will therefore not be calculated The program sets this derivative equal to zero and no additional menus appear on the screen e Cs This submodule can be used to estimate the rolling moment a coefficient due to aileron derivative e Cs This submodule can be used to estimate the yawing moment a coefficient d
20. ETT 127 7 Installed Dita EE E EEN E SA E A EE EE 127 6 Stability amp Control Module oooonoonnnnnnnnccnnonconnconoconoconocnnonnnonnnonn crac cnn nnnn conc nn cono cn neon nena ncnnninns 129 6 1 General Descrip As 129 6 2 Stability amp Control Main Window oooooccnoccconcoonconoconccononononnnonnnonnnran conc conc nnncc nena necnnos 129 6 3 Longitudinal Stability DerivativVesS ooonnncnnonnonnconnconncononononanonnnonn crac cnn nnnncnnccnnocnnocnnos 130 6 3 1 Steady State Coefficients seirene seve dll di 131 6 3 2 Speed Derivatives octal 131 6 3 3 Angle of Attack Derivatives ooooooonnonnncnnccnocnconoconoconocn nono nono nonn nono crnncrnncnnnono 132 6 3 4 Rate of Angle of Attack Derivatives ooonnonnnnnonnnocnconnconccnncnanonnnonncrancrnonnnono 133 6 3 5 Pitch Rate Derivatives orenera een a nono nncnnn nr nn nncnncnnnns 134 6 4 Lateral Directional Stability Derivatives ooonncnnnnninnnoncnoncnnncononanonancnnnonn nc noconocnnccnnos 134 6 4 1 Angle of Sideslip Derivatives oooooncnncnnncnnccnonnconnconoconocnnorononn nono crancrnonnnono 135 6 4 2 Rate of Angle of Sideslip Derivatives ooooonnnnonnnonnconnnonccnnccanonnnonnccnncnancnnnono 135 043 RollRate Derivatives rges i RE OE 136 6 44 Yaw Rate Derivatives cee ecececesccesneeceseceecceceseeeeeeecaceeeaeecsaeeeeeeecaeeeeeeeess 136 6 5 Longitudinal Control Derivatives ooonnocnnocnconnonnconoconccnnonononnnonnnonn cnn conc conc cnn c nono necnno
21. For calculating the aerodynamic shift due to the v tail 2 6 13 Ventral Fin For calculating the aerodynamic shift due to the ventral fin 2 6 14 Airplane For calculating the aerodynamic of the airplane 2 6 1 Aerodynamic Center Shift due to the Fuselage When the user selects the Fuselage button an input output window appears allowing the user to calculate the aerodynamic center shift based on the fuselage geometry defined in Section 4 3 NOTE The calculated value of the aerodynamic center shift due to the fuselage in terms of the wing mean geometric chord will be transferred into the C and C submodule input groups g g My mg put group 2 6 2 Aerodynamic Center Shift due to the Nacelles When the user selects the Nacelles button the number of nacelles must be specified in the Airplane Configuration dialog box A maximum of ten nacelles may be defined After the number of nacelles is defined an input output window appears In this window X jose is the X coordinate of each nacelle inlet Ynose is the Y coordinate of each nacelle centerline and mts is the number of segments into which each nacelle is to be divided The maximum number of segments that can be entered for each nacelle is 50 The first input group consists of a number of general aerodynamic center shift parameters The second group consists of a table listing the number of nacelle segments in the nacelle along with input parameters for each segment The Next Nacelle but
22. O O O O Q O O FAR 23 Federal Aviation Regulations Part 23 JAR 23 Joint Airworthiness Requirements Part 23 FAR 25 Federal Aviation Regulations Part 25 VLA Very Light Aircraft Mil Specs Military Specifications MIL F 8785C and MIL STD 1797A AS Specs Naval Air Systems Command Specifications Light Sport Light Sport Requirements Category Under FAR 23 or JAR 23 one of the following airplane categories can be defined O Normal PART I o Utility o Acrobatic o Commuter e Base Under Mil Specs or AS Specs not shown in Figure 2 14 the airplane base can be selected as land carrier or both Classification Dialog Box When the Classification button is selected the Classification dialog box is displayed Figure 2 24 shows the Classification dialog box The airplane type classification for military flying quality evaluation can be specified The regulations are used to evaluate flying qualities for both civilian and military airplanes Classification Classification C Class Small light airplanes C Class Il Medium weight low to medium maneuverability airplanes C Class Ill Large heavy low to medium maneuverability airplanes C Class I High maneuverability airplanes dL x Cancel 7 Help Figure 2 24 Classification Dialog Box PART I I 39 2 2 5 System Setup Toolbar The System Setup toolbar Figure 2 25 consists of seven bitmap buttons at the bottom of the main wind
23. Print and Print Setup under the File menu are used to print the output directly with no user interaction Using Print and Print Setup under the File menu the user can manipulate the print style see Figure 2 13 before the print command is sent to the printer r Print Print Options Print Parameters Options C Screendump Symbol Value Unit C Description Value Unit Print Parameters C Description Symbol Yalue Unit T Print Page Numbers M Print Date l Print File Name V Print Time av x Cancel 7 Help amp Setup Figure 2 13 The Print Dialog I 22 PART I The Print dialog box options are described in Chapter 4 In brief the Print dialog box options are e Screendump Prints a bitmap representation of the main window and any other open and visible windows e Active Window Prints a graphic representation of the active application input output or plot window e Print Parameters Prints a list of the input and output parameters in an input output window The Print Parameters option has three options o Symbol Value Unit Prints the parameter symbol the value and the unit o Description Value Unit Prints the description of the parameter the value and the unit o Description Symbol Value Unit Prints the parameter description symbol value and unit e The user can also choose whether to show the date time page number and file name on the print out 2 2 2 The File Toolbar The Fil
24. To deselect the calculator close the calculator window using the system menu or choose the Cancel button A number will be entered in the software database when the OK button on the calculator or lt Enter gt on the keyboard is selected The calculator is shown in Figure 3 1 PART I I 51 ES Calculator AR Ww are sin Ea an _sart ele EEE J OK X Cancel Info Figure 3 1 The Calculator E i The calculator is equipped with special buttons The functions of these special buttons are Info Undefine The Info button shows information about the current variable The Info button always provides the user with the definition of the variable and whenever possible with a typical value or an approximate range for that variable The user may select the typical value or range by moving the cursor to the line containing that value and pressing the left mouse button The typical value will be entered into the calculator display or a dialog box will appear allowing the user to choose a number within the suggested range Some variable descriptions include a graph or diagram which can be displayed by choosing See Graphic The Undefine button allows the user to make a variable temporarily undefined This button is designed for those cases where the user has the intention of removing the parameter for the calculation but not from the database To restore the value of the parameter invoke the calculator and pre
25. To remove objects from the list repeat the same procedure using the Current Selection list and then clicking the Remove lt lt button 4 1 2 3 Copy Distributed Weight Table This option is used to copy a distributed weight table from another flight condition A list of all of the flight conditions will be displayed By choosing one of the flight conditions all of the items in the Current Selection list from that flight condition will replace all of those in the current flight condition Loads II 175 8 4 1 2 4 Distributed Weights This option is provided to calculate and display the distributed weights from masses chosen in the Select Compon option There are two choices in this section e Y Seg If the distributed weights have been defined in the Weight module Chapter 1 those definitions will appear in the distributed weights table and the program will distribute the weight over the entire structural component to result in the specified center of gravity If this method is used the program will ignore the beginning and ending locations and weights per unit length e W L S1 S2 If the beginning and ending locations and weights per unit length are defined the program assumes that the weight per unit length varies linearly from the beginning location to the end location Fuselage distributed weights are assumed to be located along the centerline When the Calculate button is selected the program will check the input of the dis
26. Trim Analysis siinon raice seriene so cuashovscy tune ce O V E E 151 6 10 1 Trim Diagram Analysis 0 ee ee cesecesecese cece caeecaeseeeseeseeeeeaeseeeeeeseeeeeseens 151 6 102 Tongridmal TM tii 152 6 10 3 Lateral Directional TMs aeaaaee i ea a i ione Es 153 6 104 IN A ON 153 6 10 5 Trimmed Littl trom Desarme k 154 6 10 6 Trimmed Lift T Const eee cece csecseecseecseeeeeeeeeeeeeeeeseeeeeeeeeeenerens 154 6 11 Stick Free Static Margins vs ssscccctsscvssssssveciet isvesscedacessessasestcosntasced sevens codasestenasevstovesag e 154 O12 SUCK Forcen eevee cheuh ba scele he a cede wel Sue cb baste a Gncbss ses a aes 154 6 13 Aleron Forceen E E e iii case tends 154 6 14 R udder ForC Scion 154 6 15 Wing Location cintia dnd badness Aiea 154 T Dynamics Module virsotnes serere ereere oeat ESSEE ESPER Ses EEEE shah EERE Erein ERRER Sie 155 Tel General Desc iphon irnn nori deus cet EEE EE ES 155 7 2 Dynamics Main WindoW seesseseeesseeseseesssrserssesresresresrssrerrserrenresrenreserrrnsesrenereentns 155 7 3 Longitudinal Dynamics sseni senres oeereres eesin erer rs rier Erri ii 156 7 3 1 Calculate Transfer Function cooconnnccconccnnonnconnnononaconnnoonnononnnconnnnon cnn nnnnnnccno 156 7 3 2 Longitudinal Flying Qualities eee ee ceeecesecsse cree caeeeneeeeeeeeeeeeeeeeeeeeees 157 7 3 3 Longitudinal Stability Derivative Sensitivity Analysis eee 158 7 4 Lateral Directional DyNamicS oonoonnnniccnnonnonnconnconoconcc
27. and then the name weight and center of gravity positions can be entered for each mass 8 4 1 1 2 Select Concentrated Weight Component This option is used to select components to be used as a concentrated weight to act on the structural component The dialog box contains all of the weight items used in the weight and balance tables of the Weight module see Chapter 1 Components that are defined using the Add Component option subsection 8 4 1 1 1 can be found in the User Defined category The user can select objects by using the mouse button and then dragging down until all desired objects are highlighted When the gt gt Add button is pressed the objects will be added to the Current Selection list To remove objects from the list repeat the same procedure using the Current Selection list and then clicking the Remove lt lt button 4 1 1 3 Copy Concentrated Weight Table This option is used to copy a concentrated weight table from another flight condition A list of all of the flight conditions will be displayed By choosing one of the flight conditions all of the items in the Current Selection list from that flight condition will replace all of those in the current flight condition 8 4 1 1 4 Concentrated Weights This option is used to show a list of the selected concentrated weights and the corresponding center of gravity locations If the weight and center of gravity data have not already been entered into the Center of Gravity s
28. determined 2 The user leaves the thrust power undefined and the performance curves are used to determine the stall speed The performance curves should then be defined for the throttle setting in the stall maneuver The methodology used in the stall speed analysis can be found in Section 5 1 of Reference 8 3 4 4 Take off Distance To run the Take off Distance submodule the thrust or power versus speed performance curve must first be defined The take off gear down drag polar must also be known See Chapter 2 Part III of the user s manual Once the Take off Distance option has been selected another Performance II 105 window appears showing the Class II take off field length input and output parameters There are two default settings in the take off distance option These default settings described below are used if the listed variable is left undefined in the input 1 L D OEI The lift to drag ratio at take off with one engine inoperative will be found using the maximum lift coefficient at take off and the take off gear down drag polar 2 Tset The thrust setting will be determined using the defined performance curves at the lift off velocity See Part IH Section 3 4 1 in the user s manual for a description of the performance curves The methodology used in the take off distance analysis can be found in Section 5 2 of Reference 8 3 4 5 Climb To run the Climb submodule the thrust or power versus speed perf
29. development manufacturing and flight testing of prototype airplanes This submodule is to be used only for those airplane programs that are not aimed at eventual mass production Estimation of the prototype cost should be applied only to prototype programs where 1 4 airplanes are constructed If more than 4 prototypes are to be constructed the R D T E Cost option should be used for calculation of the prototype cost 10 6 Acquisition Cost After invoking the Acquisition Cost button a menu is displayed with the following four options e Engr 8 Design For estimation of the airframe engineering and design cost e Prog Production For estimation of the airplane program production cost e Test Operations For estimation of program production flight test operations cost e Total Cost For estimation of the total acquisition cost the cost to finance the manufacturing phase and the profit 10 7 Operating Cost for Military Airplanes Before selecting the Operating Cost button the user must specify the airplane civil military designation For military aircraft a menu is displayed for estimating the operating cost of military airplanes in peacetime The menu shows the following options e Fuel Oil amp Lub For estimation of the program cost of fuel oil and lubricants e Dir Personnel For estimation of the program cost of direct personnel e Consum Mat For estimation of the program cost of consumable materials e Total Operation For e
30. found in 7 4 Lateral Directional Dynamics After selecting the Lateral Direct option the following options will be displayed 7 4 1 Transfer Function 7 4 2 Flying Qualities 7 4 3 Sensitivity Dynamics For transfer function computation For flying qualities checking For lateral directional derivative sensitivity analysis H 159 7 4 1 Transfer Function Transfer Function is used to calculate the lateral directional characteristic equations associated modes including frequencies and damping ratios and the associated transfer functions After choosing Transfer Function the user is presented with the following options e Aileron For calculating the aileron open loop transfer function e Spoiler For calculating the spoiler open loop transfer function e Vert Tail For calculating the vertical tail open loop transfer function e Rudder For calculating the rudder open loop transfer function e V Tail For calculating the v tail open loop transfer function e Ruddervator For calculating the ruddervator open loop transfer function e Spoileron For calculating the spoileron open loop transfer function NOTE The corresponding control surface must be defined by the user in the Control Surfaces Configuration dialog box before the dynamic analysis can be conducted After all input parameters are defined and the Calculate option is selected the lateral directional transfer function output parameters will be displayed and the
31. has been selected another window appears showing the Class II Landing Distance input and output parameters If the approach speed is known rather than the landing stall speed the stall speed is left undefined and the approach velocity is entered in the output The methodology used in landing distance analysis can be found in Section 5 9 of Reference 8 3 5 Export to APP This module exports aerodynamic and other pertinent data into Aircraft Performance Program APP native files These files can be used in APP T 110 Performance 4 Geometry Module 4 1 General Description The purpose of the Geometry module is to help the user define the geometry of the fuselage wing horizontal tail vertical tail canard and v tail and calculate related parameters The methodology used to calculate the airplane components geometry is described in Reference 2 Use of the Geometry module options will be described in the following sections 4 2 Airplane Geometry Categories Twelve options represent the following submodules e Wing e Horizontal Tail e Vertical Tail e Canard e V Tail e Ventral Fin e Fuselage e Nacelle e Tailboom Geometry For calculation of the wing geometry The options provided in this module are discussed in Section 4 3 For calculation of the horizontal tail geometry The options provided in this module are discussed in Section 4 4 For calculation of the vertical tail geometry The options provided
32. if the user wishes to insert a filter into the feedback system Dynamics I 165 The user would select the User Defined option to input the filter transfer function The required transfer functions must be defined in the input output window that appears The procedure for entering a transfer function into a Control loop is presented in the following steps 1 2 3 II 166 Select the desired transfer function from the presented list and the numerator and denominator polynomial input window will appear on the screen Boxes Ao As are the coefficients of the numerator polynomial and boxes Bo Be are the coefficients of the denominator polynomial for respective powers of s The polynomial may have a maximum degree of six The A and B coefficients are extracted from the database if the user selected one of the standard airplane transfer function options 1 9 If the user selected the User Defined option the A and B coefficients must be entered as input by the user If the user wishes to change a user defined transfer function which has already been defined in a T F box the User Defined option must be selected after which the polynomial coefficients can be changed This procedure will load the new transfer function into the program database replacing the existing one If the user wishes to transform an airplane transfer function in terms of an angle for example roll angle to aileron transfer function to a transfer function i
33. is made that flap deployment will occur only in the subsonic speed regime The flap drag estimation is therefore only valid in the subsonic speed regime The methodology used to calculate the flap drag coefficient is described in Section 4 6 of Reference 7 2 4 2 13 Landing Gear Drag In this module the drag coefficient for different landing gear types can be estimated The landing gear drag estimation is only valid in the low subsonic speed regime The methodology used to calculate the gear drag coefficient is described in Section 4 7 of Reference 7 Landing gear estimation methods are available for the following gear types e Fixed For estimation of the fixed gear drag coefficient Aerodynamics Il 89 e Retractable For estimation of the retractable gear drag coefficient The number of gears must be entered before the input output window will appear 2 4 2 14 Canopy Drag In this module the canopy drag coefficient can be calculated The methods used in this module are only valid in the subsonic speed regime In the transonic and supersonic speed regimes the wave drag generated by the canopy can be significant In these speed ranges it will be necessary to employ area ruling to reduce the wave drag to a minimum The methodology used to calculate the canopy drag coefficient is described in Section 4 8 of Reference 7 2 4 2 15 Windshield Drag In this module the windshield drag coefficient can be calculated The methods used in this
34. methodology used to determine the lifting surface spanwise lift distribution is based on lifting line theory To determine the spanwise lift distribution select the Lift Distribution button An input output window appears showing the surface planform input parameters The user can display the lift distribution by selecting the Plot button 2 3 1 3 Airplane Lift Coefficient and Downwash Zero angle of attack This submodule can be used to determine the lift coefficients downwash angle and upwash angle at zero airplane angle of attack and the zero lift angle of attack This module includes flap effects Before the Cz submodule can be selected the user must define one of three types of airplane configurations in the Airplane Configuration dialog box See Section 2 2 3 Figure 2 16 II 82 Aerodynamics 2 3 2 High Lift Devices The function of the High Lift Device submodule is to determine the type and size of high lift devices needed to meet the maximum lift requirement for take off and landing conditions To determine the type and size of the high lift devices needed select the High Lift Device button The user must first select the desired flap type from Flap Slat dialog box See Section 2 2 3 Figure 2 19 Two options are presented in the High Lift Device submodule These options are as follows e Sizing For calculating the required high lift device size the desired values of the trimmed lift coefficients can be determined from th
35. moment equation e Stick force equation The four equations contain six variables Therefore the user must define two of the following variables as control variables ea Airplane angle of attack ey Flight path angle Elevator deflection angle Oot oani Additional elevator trim tab deflection angle required to trim e i Stabilizer incidence angle e F Stick force If the angle of attack is selected as the first control variable the flight path angle cannot be IT 152 Dynamics selected as the second variable The remaining four variables not defined by the user as control variables are calculated and displayed in the output along with various gradients important to flying qualities 6 10 3 Lateral Directional Trim The Lateral Directional Trim submodule is used to solve the lateral directional equations for trim Solving the following equations simultaneously completes the lateral directional trim calculations for either flight condition e Side force equation e Rolling moment equation e Yawing moment equation e Uudder pedal force equation e Aileron stick or wheel force equation The five equations contain eight variables so three of the following variables must be defined Bank angle e p Sideslip angle Aileron deflection angle e Rudder deflection angle Oat rim Additional aileron trim tab deflection angle required to trim On An Additional rudder trim tab deflection angle require
36. of an airplane may vary if one of the derivatives weight or inertia parameters is varied H 158 Dynamics Once the input parameters have been entered in the Transfer Function submodule the user can enter the Sensitivity submodule The following sensitivity analysis options will be displayed e Angle of Attack a e Pitch Rate q e Speed u e Thrust T e Weight e Inertia e Steady State The parameters that can be varied are C Da gt C La Cing E Le and Ci The parameters that can be varied are C Lo and Cing The parameters that can be varied are Cp Cz and Crm The parameters that can be varied are Cy and Cmp gt Cr y u a The parameter that can be varied is Woyrrent The parameter that can be varied is 1 p The parameters that can be varied are Cp gt Cy gt Ci gt Cmr Ors and Uj Once the user has selected one of the sensitivity analysis options an input window will appear asking the user to define the lower and upper limit of the variable to be investigated The sensitivity plot is displayed after these limits are defined and the Plot option is selected The plot options are described in Section 2 1 4 of Part II of the manual The input parameters required to perform a sensitivity analysis for the longitudinal derivatives can be entered in the Transfer Function submodule Section 6 4 of Reference 11 The methodology used for the sensitivity analysis can be
37. of categories for various types of airplanes Aerodynamics Il 77 e General Aviation e Commercial e Military Transport Cargo e Military Bomber e USAF Fighter Land Based e Navy Fighter Land Carrier Based For homebuilts single engine props twin engine props agricultural regional turboprops with a Wro below 12 500 Ib low speed military trainers and small and low speed flying boats amphibious or float airplanes For business jets regional turboprops with a Wro above 12 500 lb jet transport commercial supersonic transports and large and high speed flying boats amphibious or float airplanes For military patrol bomb and transport airplanes supersonic patrol bomb and transport airplanes and military large flying boats amphibious or float airplanes For military patrol bomb and transport airplanes supersonic patrol bomb and transport airplanes For high speed military trainers fighter airplanes and attack airplanes operating from land bases For high speed military trainers fighter airplanes and attack airplanes operating from land and carrier bases Aerodynamics 2 Aerodynamics Module 2 1 General Description The purpose of the Aerodynamics module is to estimate the airplane lifting surface coefficients and the total airplane drag coefficient for different flight conditions Use of the Aerodynamics module options will be described in the following sections 2 2 Aerodynamics Main Window After
38. ooococnnoncncnconononnnoncnancnnnconoconoco nooo nono no nn nono nrnn cnn cnn E 5 2 1 2 Input Output Windows cc scecsesscesetsnes sheesseebesstescbpsbsnsebensoeecbeebseessetss ened 6 2 1 3 Input Output Window Command Bal ooooconocccoconoccnnnnonnncnnnnnnnn conc cono conoconocnnos 11 21 4 Plot Wind WS mies Es 12 2 1 5 Plot Window Command Bat cooocnnncnocncocononcnononnonnnonnnonnnonnnnnnonn conc nono c naco necnnos 14 2 22 E sc e one att elie din eis abe tea eed EE 15 2 2 1 Main Toolbar ninio sas osa cintia 15 2 22 The Pile Toolbar iii 23 2 2 3 Configuration Setup Toolbat oooncnncnnnnonnnonnnoncconcnncnncnnn crac cnn conc corona nono nccnnos 24 2 24 Certification ToolbaL cececiicosioidon lotssastinas docs i dior oan dosess bseestetesteavetvsnseets 36 225 System Setup Toolbaf sceno corista e En E e danna 40 23 Y NTE 46 23410 Ele lMpoltrain iii RO 48 232 File Export 22 cci cscces tants sais aras 48 3 Input Devices for the Software A ba EEE ESEPen Ee ESPERES E Sra TE EESE S En 51 3 1 Operation of the Mouse and Cursor esssseeseseeeseeesssreesesrrersstsresrrsresrsseerrssenrenresrenees 51 3 2 Operation of the Calculator ooonnncnnnnnnnnnnnnnonnnonconnnanonononanononcnnoconocn nono nc conc on corn ncnnccncnnss 51 3 3 gt Rey board cis reer iachantie es sass sua a EEr O EEEE E E n EEE EEr EEE ER 53 34 System Messages sosire eeen iens a 53 3 5 Flexible Put A 53 4 Generating Hardcopy O
39. rate to rudder transfer function must be defined in T F box 7 according to Gyro Tilt analysis theory see Section 13 1 1 2 of Reference 12 3 Yaw rate to rudder and roll rate to rudder are not part of the standard airplane transfer functions and must be defined using the User Defined option see Section 7 6 of the user s manual After all the required transfer functions have been defined the Plot Open Loop button is selected to calculate the root locus of the control loop A new input menu will appear on the screen asking the user to specify the steady state angle of attack plus the angle between the gyro sensitive axis and the Z axis of the body system a a In addition the user is asked to define the lower and upper limits for the system gain damping ratio and natural frequency the number of gain damping ratio and natural frequency points and the design gain If desired the design gain number of damping ratio points and number of natural frequency points may be left undefined After selection of the Plot option the root locus figure will be displayed A detailed explanation of the plot options can be found in Chapter 2 in Part II of the user s manual 7 8 Bode Method The function of the Bode analysis submodule is to develop a frequency response of an airplane in the form of a Bode plot On a Bode plot the frequency is plotted as the logarithm of frequency and the amplitude is plotted as twenty times the logarithm of the amp
40. selecting the Aerodynamics module the user can select from the five options displayed e Lift e Drag e Moment e Aerodynamic Center e Power Effects e Ground Effects Aerodynamics The Lift submodule can be used for estimations of the lifting characteristics of airplane lifting surfaces and high lift devices The options provided in this module are discussed in Section 2 3 The Drag submodule can be used for estimations of the drag characteristics of the airplane The options provided in this module are discussed in Section 2 4 The Moment submodule can be used for estimations of the moment characteristics of the airplane The options provided in this module are discussed in Section 2 5 The Aerodynamic Center submodule can be used to calculate aerodynamic center locations of airplane lifting surfaces The options provided in this module are discussed in Section 2 6 The Power Effects submodule can be used to calculate the effects of power on the horizontal tail dynamic pressure ratio and downwash angle and on the airplane lift and pitching moment coefficients This option is available for tail aft airplanes The options provided in this module are discussed in Section 2 7 The Ground Effects submodule can be used to calculate the ground effects on airplane lift The options provided in this module are discussed in Section 2 8 Il 79 e Dynamic Pressure Ratio 2 3 Lift The Dynamic Pressure Ratio submodule can be use
41. surfaces are disabled when an empennage surface is not selected Descriptions of the dialog box functions follow e Control Surface On the Wing page the user can select aileron drooped aileron aileron trim tab and elevon On other pages a flap type control elevator canardvator ruddervator and rudder variable incidence and trim tab can be chosen The horizontal tail can include a differential stabilizer for roll control e Surface Tip Shape The stability surface tip shape can be chosen for hinge moment estimations e Nose Shape The aileron elevator canardvator and rudder nose shapes can be chosen for hinge moment estimations e Type of Balance The type of balance of the aileron elevator canardvator and rudder can I 32 PART I be chosen e Type of Horn Balance If the aileron elevator canardvator or rudder includes a horn balance the type of horn balance can be chosen e Horn Balance Shape For fully or partially shielded aileron elevator canardvator and rudder horns the horn nose shape can be chosen Flap Slat Dialog Box When the Flap Slat button is selected the Flap Slat dialog box is displayed The dialog box displayed in Figure 2 19 allows the user to select the number and type of high lift devices per wing Descriptions of the dialog box follow e Number of High Lift Devices per Wing The user may specify the number of high lift devices per wing e Trailing Edge Device The user may specify the type of t
42. the missing weight component input data The error may be a result of undefined data in the Class Il Weight estimation menus or weight fractions not defined in the Class I Weight estimate submodule The Class II Weight input can be defined in the Structure Powerplant and Fixed Equipment Weight Estimation submodules The Class I weight fractions can be estimated in the Class Weight estimation module after selection of the Weight Fraction option Aerodynamics Il 75 The iteration table consists of an input group and a component weight table The hashed cells in the table indicate that weight estimation methods are not available for the specific weight component for the specific airplane category or engine type All other cells except the last column which represents the average weights represent a weight component for a specific weight estimation method and can be deselected Deselecting a component in the table will cause that specific component to not be used in the weight Class Il iteration process The user can deselect a method by undefining the specific box delete the input from the corresponding cell in the table When the iteration process is activated by selecting the Calculate button all blank cells will not be included in the iteration The estimated take off weight will be displayed below the iteration table To include an estimation method that has been deselected type any number in the corresponding box When the Calcu
43. the propeller driven option has been selected then the supersonic flow regime cannot be used Aerodynamics T 91 2 4 2 23 Inlet This module is used to calculate the inlet drag coefficient caused by subsonic inlet spillage 2 4 2 24 Nozzle This module is used to estimate the drag due to exhaust nozzles 2 4 2 25 Total Drag In this module the total drag coefficient can be calculated The calculation consists of the summation of all previous estimated component drag coefficients 2 4 2 26 Recalculate All This module allows the user to recalculate all of the drag coefficients by pressing calculate once The user can change a parameter in the input and view its effect on the drag coefficients A drag coefficient will only be recalculated if all of its associated input is complete 2 4 2 27 Trendline In this module a fifth order curve is fit to the Class II Drag Polar 2 4 2 28 Plot In the Class II Drag module different drag polars can be generated After invoking Plot the following five options will be displayed e Cr Cp M Const For generating drag polars for constant Mach number After selection of this option the drag component selection list will be displayed Each drag component selected in this menu will be used for generating the Class II drag polars e Cr Cp 9f Const For generating drag polars for constant flap deflection angles This option can only be used in the subsonic flow regime After selection of this op
44. transfer function will be shown in dialog boxes When certain output variables such as time constants damping ratios or frequencies cannot be calculated the program leaves a blank in the appropriate output box The transfer functions will be redisplayed when selecting the Calculate option These transfer functions can be printed either directly from the dialog box or by selecting the Print Parameters options in the Print dialog box 7 4 2 Lateral Directional Flying Qualities Before invoking the Flying Qualities option the user must select the aircraft certification category in the Certification and Type dialog box for which the flying qualities of the aircraft configuration are to be checked For FAR 23 certified aircraft the FAR 23 157 regulations are used for roll mode checking and MIL F 8785C regulations are used for spiral and Dutch roll mode checking Because the FAR 25 regulations do not have lateral directional flying quality requirements the MIL F 8785C for land based aircraft are used for FAR 25 certified aircraft In II 160 Dynamics case of a MIL F 8785C certified aircraft the user selects the aircraft base e Land Based For land based airplanes e Carrier Based For aircraft carrier based airplanes e Land amp Carrier Based For land and carrier based airplanes The user also needs to define the airplane class from the following options in the Classification dialog box e Class I Small light airplanes e Class II Medium w
45. two elevator tab related derivative options e Crs This submodule can be used to estimate the lift coefficient due to et elevator tab deflection derivative e Cms This submodule can be used to estimate the pitching moment et coefficient due to elevator tab deflection derivative 6 5 4 Canard Related Derivatives The methodology used to calculate the canard related derivatives can be found in Section 10 3 3 of Reference 7 Once the Canard button is selected a new menu appears with the following three canard related derivative options e Cp This submodule can be used to estimate the drag coefficient due C to canard incidence derivative e C L This submodule can be used to estimate the lift coefficient due to C canard incidence derivative e Crm This submodule can be used to estimate the pitching moment m Ic coefficient due to canard incidence derivative II 138 Dynamics 6 5 5 Canardvator Related Derivatives The methodology used to calculate the canardvator related derivatives can be found in Section 10 3 4 of Reference 7 When the Canardvator button is selected a new menu appears with the following three canardvator related derivative options e Cos This submodule can be used to estimate the drag coefficient due cv to canardvator deflection derivative e Cis This submodule can be used to estimate the lift coefficient due to CV canardvator deflection derivative e Crn 5 This submodule can be used
46. user selects the Climb option Note that the input elements which pertain to the specification to be satisfied contain default values These values can still be modified by the user as needed by selecting the Set Default option The drag polar parameters can be estimated in the Class Drag submodule See Chapter 2 The methodology used in sizing to climb requirements can be found in Section 3 4 of Reference 1 3 3 4 Maximum Cruise Speed Sizing To size the aircraft to maximum cruise speed requirements the user selects the Max Cruise Speed option The methodology used in sizing to maximum cruise speed requirements can be found in Section 3 6 of Reference 1 3 3 5 Maneuvering Sizing To size the aircraft to meet maneuvering requirements the user selects the Maneuvering option This submodule sizes the airplane for both pull up or push over load factor and a specific turn rate in the level turn maneuver The methodology used in sizing to maneuvering requirements can be found in Reference 1 Section 3 5 3 3 6 Landing Distance Sizing To size the aircraft to meet landing distance requirements the user selects the Landing Distance option The methodology used in sizing to landing requirements can be found in Reference 1 Section 3 3 Performance II 103 3 3 7 Matching Plot After entering the data in the desired performance sizing options the requirements can be shown in a performance matching plot by selecting the Matching Plot opt
47. value in the output group The A and B values should be left undefined If A and B are already defined they should temporarily be made undefined by using the Undefine button on the calculator described in subsection 4 2 When the user clicks the Calculate button on the input output window command bar the system checks all the variables in the input group Since A and B are undefined the program logic ignores the part related to the calculation of D Then it checks the database for the required parameters for calculation of E In this case it will find that D and C are given and E can be calculated from equation IV In this example the system ignores its internal logic to calculate D and gives authority to the user to supply the desired D value I 54 PART I 4 Generating Hardcopy Output Hardcopy output in the software can be in one of the following forms e Screendump This is performed by clicking the Print button on the main toolbar and selecting the Screendump option in the Print dialog box See Subsection 2 2 1 When the OK button is clicked a bitmap representation of the main window is sent to the default printer The print may take one or two minutes on slower printers e Active Window This is performed by selecting the Active Window option in the Print dialog box see Subsection 2 2 1 This option is only available when an application window input output window or plot window is open within the main window When the OK button is cl
48. 1 5 13 Weight Sizing The purpose of the Weight Sizing submodule is to rapidly estimate the following weight components and or sensitivity coefficients Aerodynamics Il 61 e Take off weight Wro e Empty weight Wg e Mission fuel weight Wr e Sensitivity of take off weight to aerodynamic propulsion and mission parameters These parameters are estimated on the basis of the following inputs e A mission specification optional e Statistical relations between empty weight and take off weight of existing airplanes Once this selection has been made a new menu appears with six options The six options displayed are 1 3 1 Mission Profile 1 3 2 Take off Weight 1 3 3 L D from Weights 1 3 4 Regression 1 3 5 Sensitivity 1 3 6 Useful Load Allows the user to calculate the overall mission fuel fraction by entering the flight segments from the start to the end of the mission Calculates mission fuel weight take off weight and empty weight A plot showing the iteration process is provided Calculates the climb cruise turn and loiter lift to drag ratios based on corresponding segment weights It iterates between Weight Sizing and Class I Drag since the wetted area is calculated from take off weight to obtain a converged solution for the take off weight With this option the user can enter statistical data by providing empty weights and take off weights in a table and calculate the regression coefficients A and B
49. 10 8 For estimation of the life cycle cost of an airplane program The life cycle cost is defined as the sum of R D T E cost acquisition cost operating cost and disposal cost The options provided in this module are discussed in Section 10 9 For estimation of the engine price propeller price and airplane price as well as a rapid method for estimating prices of future designs The options provided in this module are discussed in Section 10 10 After invoking the AMPR Weight button the user is presented with two options e Detailed e Alternate 10 4 R D T E Cost Application of this method requires a detailed knowledge of the airplane component weights In case detailed data of the airplane component weights is lacking this submodule estimates the AMPR weight for a given take off weight Selecting the R D T E Cost button will display the following five options e Engr amp Design e Dev Sup amp Test e Test Airplanes e Test Operations e Total Cost I 184 For estimation of the airframe engineering and design cost For estimation of the development support and testing cost For estimation of the flight test airplanes cost For estimation of the flight test operation cost For estimation of the total R D T E cost the test and simulation facilities cost the R D T E profit and the cost to finance the R T D E phases Cost 10 5 Prototype Cost The function of this option is to estimate the cost of
50. 118 4 6 3 Volume Coefficient neinni asseris e e i n E EE ei 119 4 6 4 Canardvator Tab cii e oean p teno e Ee N are 119 4 6 5 Chord Length aienea a a E A E S 119 46 607 119 ALOT EXpOS d E EE E TE 119 VaT ail GEometIy oeseri eea aida 119 47 1 Straight Tapered icon 120 A dl NA O 120 4 73 Volume Coefficient 0 eee ee eethoek iee oeii ane ko iest 120 407 4 Ruddervatot T ab cit ee vos 120 ATS Chord Lengths isre ernie sn ee eare p e E t n 120 ACRO gt Alone 120 ALT OS 120 Ventral Fin Geometria 121 Fuselage Geomethy sir sie Shes iis SO ea hes Aenea ie hes ek 121 Nacell Geometry EE EER ETT 121 lc A erani es dus shane we EEEE Ea E EEES 121 Float Geometry n E pias 121 Landing Gears aa 121 Cad Yate 122 Store GEOMCUY ON 122 Py lomsGeOmetyy dni 122 Propeller Geometria 122 Airplane Mi Wi ins rocas 122 Arles rd id 122 MC O es ira dd 122 Translat nn rai nit 123 Airfoil FOLGER aii 123 AeroP ack anti a aedaibred teape teens iths E EEEE 123 Table of Contents 4 24 Exporting to Shark SharkPX AP ooococoncnccnconnocnconnconoconoconocnnonononnncnn cnn nnnnnnn nc nora nccnnos 124 gt Propulsion Module iio a SAS Rin eas Re ik 125 SI General Descriptio rren ne n T e in orar 125 92 Type of Propulsion reie nioi r ee ies 125 5 3 Propulsion Main WiNdOW cccnnncnonnnonnconnconcconconnconocnnconncnnnonnnonn conan rn nono non sie i res 126 HA Power Extract E E A A ESR 126 5 Inlet Desire S 127 9 6 Nozzle Desi e a S EE
51. 8 4 1 1 Concentrated WelghtS oooonccnncnonnnonnconnconcconcnanonnnonnnnannnnnnnccnnoconocnnos 173 8 4 1 1 1 Add Concentrated Weight Component eee eee 174 8 4 1 1 2 Select Concentrated Weight Component 174 8 4 1 1 3 Copy Concentrated Weight Table eee cree 174 8 4 1 1 4 Concentrated Weights ooooonocnnonnonnnoconocconnnonocononanconoconocnnos 174 8 4 1 2 Distributed Weights ooncnncninnnonnconnconcconocononnnonnnonnnnancnnncnn nono conocnnos 175 8 4 1 2 1 Add Distributed Weight Component eee eee 175 8 4 1 2 2 Select Distributed Weight Component eee 175 8 4 1 2 3 Copy Distributed Weight Table ee eeeeeeeeeeee 175 8 4 1 2 4 Distributed Weights 00 eee eee eeeeeeeeeeeeeseeeeeeeceseenaee 176 8 4 1 3 User Aerodynamic Loads cooooccnoccconoconcnononnnconcnanonannnancnn nc noconocnnconnos 176 8 4 1 4 Load tao ic Dn 177 8 4 1 5 G ar EnSine MiSCs sen esirin ioeie rererere Rao onten ir srt 178 8 4 1 6 Total Internal sies oir oie E OENES 178 84 27 Load FactoiS eeann oie helen ate ta ois a EEA aha R 179 9 Structures Module irio dano otto EEEN RIRE Ees 181 91 General DescrUptiON sion E EEEE A A A O EEEE EEE 181 9 2 Structures Main Menu sx vce ccssesses eeesseveketbecs biceshetsncs pei eraen ea EE ri r E nir S oE 181 Table of Contents 9 2 1 Clas AAA as 181 10 Cost Analysis Module il 183 10 1 General Descrip teo cee doses sons ae a E rE EEn EEEE IREN ia 183 10 2 Cost M in Wind
52. Aerodynamic Center cece cece cseecseeeeeeeeeeeeeeseceeeeeseesseneenseesaes 98 2 6 9 Horizontal Tail Aerodynamic Center ooooccnonononononcnnnnononanonnnonn nono nonoc nono nocnnos 99 2 6 10 Vertical Tail Aerodynamic Center coooconnccconcnononononnonnnnnnnnnnnnnnonn nono nnnc crono necnnos 99 2 6 11 Canard Aerodynamic Center cooonocnnccnonoconoconccononononnnonnnnanonnonnnonn conc conc cn nccn recono 99 2 6 12 V Tail Aerodynamic Centel ssassn rk esisi cnn conc cnn nc neco recono 99 2 6 13 Ventral Fin Aerodynamic Center cooooocccocccononononononnnonnnonnnonnnnnnonnnnnncnncc nono necnnos 99 2 6 14 Aerodynamic Center of the Airplane eee ee eeeeceeeeeeeeereneeeseenaee 99 2 7 Power Electro dio 99 2 3 Ground Effects eeens oer E EEEE dd a ie 99 2 9 Dynamic Pressure Ratios s ccisisce seesseveseacees ea es EE EEr T EE EE EE iets 100 Performance Modules eseina iren a asa 101 3 1 General Descipton e ere n e r BRAG A A eRe ey 101 3 2 Performance M in Window cc csesssecissesssseoscosscessevsseessevesessevscssecesotssssosnesaseseosveese 102 3 3 Performance SILO ciendo EE ASEE addict eds 102 33 1 Stall Speed SIZIN onnen ap na dilata 102 3 32 Take off Distance SILO ins fiian 103 3 3 3 Climb S1Zin s c 3 svete ate hen eae e o seid AAT Ge aaeeess 103 3 3 4 Maximum Cruise Speed Sizing oooooccconnconnnonnnoncnancnnncnoconocnnconc cono conccnncnnnonns 103 3 30 gt Manet vertim SII sssr npese e aoi Ee E e 103 3 3 6 Landing Distance Sizing
53. All load stations are defined in the weight and balance axis system For horizontal surfaces the load station locations are defined in the Y direction to represent spanwise locations Vertical tail load stations are defined in the Z direction to represent spanwise locations Fuselage load stations are defined along the length of the fuselage in the X direction The number of load stations may be selected in the input parameter list with a maximum of 21 load stations The location of each load station may be changed in the load station table with the exception of the first The location of the first load station must be zero and the location of the last load station may not be larger than the span of the surface or length of the fuselage All other loads stations must be in sequential order 8 4 1 5 Gear Engine Misc The Gear Engine Misc module can be used to define any landing gear reactions engine thrust forces or miscellaneous loads not defined elsewhere These reactions can be specified outside of the structural component These are similar to the Concentrated Forces and Concentrated Moments defined in the User Loads and Aerodynamic Loads modules The structural component onto which each Gear Engine Misc load acts must be chosen The locations of these loads are defined by coordinates in the weight and balance axis system This module is flight condition dependent and will be reset for each flight condition 8 4 1 6 Total Internal The Tot
54. Brake In this module the speed brake drag coefficient can be calculated The methodology used to calculate the speed brake drag coefficient is described in Section 4 12 of Reference 7 Before entering the speed brake drag input output window the user will be required to define whether the speed brake is retracted or deployed for the particular flight condition 2 4 2 20 Miscellaneous Drag In this module the miscellaneous drag coefficient can be calculated Miscellaneous drag may consist of drag due to struts antennas or exhaust nozzle integration to name a few possibilities The methodology used to calculate the miscellaneous drag coefficient is described in Section 4 12 of Reference 7 2 4 2 21 Pylon Drag In this module the pylon drag coefficient can be estimated The methodology used to calculate the pylon drag coefficient is described in Section 4 5 of Reference 7 Before selecting the Pylon option the user must specify the number of pylons For estimation of the pylon drag coefficient the pylon drag coefficient consists only of a zero lift component It is assumed that the lift dependent drag component of a pylon is negligible 2 4 2 22 Windmilling Drag In this module the drag coefficient due to a windmilling condition can be estimated The methodology used to calculate the nacelle drag coefficient is described in Section 4 5 of Reference 7 Before selecting the Windmilling option the user must specify the type of engine If
55. Fuel Tank C Integral C Self Sealing Bladder C Non Self Sealing Type of Inlet Prop Type C Plenum C Composite C Straight Through C Metal Prop Pitch C Fixed C Variable 1 el 3 4 5 6 7 8 9 elele tele eeel e YD eee e e elelee e 1010 o ala uni elelee 110 eleele oT ee Ra e SNe NCH iG ee Ned elie o gt Engine Air Flow Location Number of Propellers 0 C Not Buried no airflow from fuselage base C Buried in Fuselage airflow from fuselage base X Cancel 7 Help Figures 2 17c Powerplant Selection Dialog Box for a Propeller Propfan Propulsion Engines Nacelles Propulsion Engine Type Yes No Number 0 C Yes C No C Jet C Piston Propeller C Propfan C Ducted Fan T p Engine Location Nacelle Location Wing Fuselage Other Wing Fuselage Other c 3 C E G E Fuel Tank C Integral C Self Sealing Bladder C Non Self Sealing Type of Inlet Prop Type C Plenum C Composite C Straight Through C Metal Prop Pitch Prop Reversing C Fixed C Reversibl C Variable C Non oon on a uni 1 OL BY 0 YL 11211 0116211 e 1111121101160 11 0 o wN omo an ES YL OL italie eiee ie Tied ey e teite See te tea Te eTe _ o a Engine Air Flow Location Number of Propellers 0 C Not Buried no airflow from fuselage base C Buried in Fuselage airflow from fuselage base C X Cancel Help Figures 2 17d Powerplant Selection Dialog Box for a Propeller Turboprop I 28 PARTI
56. July 2014 Advanced Aircraft Analysis User s Manual Version 3 6 The software described in this document is furnished under a license agreement The software may be used or copied only under the terms of the license agreement Advanced Aircraft Analysis User s Manual 2014 No part of this manual may be photocopied or reproduced in any form without prior written consent from Design Analysis and Research Corporation Microsoft Windows Windows 95 Windows 98 Windows NT Windows 2000 Windows XP Windows 7 are trademarks of Microsoft Corporation While the information in this publication is believed to be accurate Design Analysis and Research Corporation makes no warranty of any kind with regard to this material including but not limited to the implied warranties of merchantability and fitness for a particular purpose Design Analysis and Research Corporation shall not be liable for errors contained herein or for incidental or consequential damages in connection with the furnishing performance or use of this material Information in this publication is subject to change without notice Copyright 1989 2014 DARcorporation Table of Contents A A iach ft sol cau nal Mee oe deed cea at 1 1 Intro dC unir ia des 1 AS co hssasect css cteskspeceeedvebeesssavesoctesvpsssedasesdbesseesseebeesces ses 3 2 1 Windows and Command Bars 00 cece eesceseesecseceseceecaeecseseaeeeaeseeeeseeesseseeeseeaeeaeenaes 3 2 1 1 Application WindOWS
57. L This submodule can be used to estimate the lift coefficient due to flying wing incidence derivative e Cn This submodule can be used to estimate the pitching moment coefficient due to flying wing incidence derivative 6 5 11 Elevon Related Derivatives The methodology used to calculate the elevon related derivatives is based on theory in Section 10 3 5 of Reference 7 When the Elevon button is selected a new menu appears with the following three elevon related derivative options e CDs i This submodule can be used to estimate the drag coefficient due e to elevon deflection derivative e Crs This submodule can be used to estimate the lift coefficient due to e elevon deflection derivative e Cn oy This submodule can be used to estimate the pitching moment e coefficient due to elevon deflection derivative 6 5 12 Elevon Tab Related Derivatives The methodology used to calculate the elevon tab related derivatives derivatives is based on theory in Section 10 3 5 of Reference 7 When the Elevon Tab button is selected a new menu appears with the following two elevon tab related derivative options e CLs This submodule can be used to estimate the lift coefficient due to elt elevon tab deflection derivative Dynamics T 141 Cing This submodule can be used to estimate the pitching moment e coefficient due to elevon tab deflection derivative 6 6 Lateral Directional Control Derivatives The Lateral Directional
58. Program Options dialog box Figure 2 29 When the software is exited normally the contents of the Working Directory are cleared If the software exists abnormally due to a program or system crash the current project will remain in the Working Directory The next time the program is started it will attempt to recover the contents of the Working Directory as a recovery project It is strongly recommended that the user save the recovered project as a new project until 1t is confirmed that the recovery was successful and complete Warning The Working Directory should be reserved for use only by the software The user should not attempt to save projects or copy files from other locations to the Working directory I 58 PART I 6 Help System The Help system is an online system of help topics that give the user information about the use of the program and the theory and definitions behind the calculations The online help system can be accessed in six ways e Select the User s Manual icon in the Airplane Design and Analysis Program group e Select the Help button on the main toolbar see Subsection 2 2 1 e Select an item in the Help menu on the main window menu bar see Section 2 3 e Select the Info button on the calculator or any other dialog that includes a Help button e Select the button on an input output element e Select the Theory button on an input output window The online help system consists of three types of help t
59. Rate of Angle of Sideslip Derivatives The methodology used to calculate the rate of sideslip related derivatives can be found in Section 10 2 5 of Reference 7 When the Sideslip Rate button is selected a new menu appears with the following three rate of sideslip related derivative options C a This submodule can be used to estimate the side force coefficient p due to rate of sideslip derivative e C This submodule can be used to estimate the rolling moment coefficient due to rate of sideslip derivative Dynamics H 135 e Cr This submodule can be used to estimate the yawing moment coefficient due to rate of sideslip derivative 6 4 3 Roll Rate Derivatives The methodology used to calculate the roll rate related derivatives can be found in Section 10 2 6 of Reference 7 Once the Roll Rate button is selected a new menu appears with the following three roll rate related derivative options ec Yp This submodule can be used to estimate the side force coefficient due to roll rate derivative Ci This submodule can be used to estimate the rolling moment coefficient due to roll rate derivative Cry This submodule can be used to estimate the yawing moment coefficient due to roll rate derivative 6 4 4 Yaw Rate Derivatives The methodology used to calculate the yaw rate related derivatives can be found in Section 10 2 8 of Reference 7 When the Yaw Rate button is selected a new menu appears with the following thr
60. The methodology used in the dive and descent analysis can be found in Section 5 6 of Reference 8 When all the input parameters have been defined and the Plot option has been selected the desired curves will be displayed The two lines represent the two equations of motion for the dive condition The purpose of this is to determine the speed for a given flight path angle in dive The velocities at the intersections of the two curves represent solutions for the desired flight path angle 3 4 8 Maneuver To run the Maneuver submodule the thrust or power versus speed performance curves must first be defined The clean configuration drag polar must also be known Once the Maneuver option has been selected another menu appears showing the maneuver options e Pull Up Push Over For a pull up or push over maneuver e Level Turn For a level turn maneuver Once the Pull Up Push Over or Level Turn option has been selected another list appears showing the maneuver types e Instantaneous For an instantaneous maneuver e Sustained For a sustained maneuver When the maneuver option and the maneuver type have been selected another window appears showing the Class II Maneuver input and output parameters The methodology used in the maneuver analysis can be found in Section 5 7 of Reference 8 When the Plot option is selected a plot of the maximum load factor versus speed is generated using the input from either the instantaneous pull up push over o
61. ULION scott iia 82 Table of Contents 2 4 2 5 Table of Contents 2 3 1 3 Airplane Lift Coefficient and Downwash Zero angle of attack 82 2 3 2 Y ES ET Error Bookmark not defined Dec a 83 2 4 1 Estimation of Class I Drag Polars ooonnnnnnonnnicnnonnnoncnnncnnncnnnnoncnannnnccnnocnocnnos 84 2 4 2 Estimation of the Class II Drag Polar ooooncnnnninninocononconnnonnnancnancnnncnnccnocnnos 85 24 221 Wing Drags nics sat ations naa 87 2 4 2 2 Horizontal Tail Drage eee e n E Ea a E aa 87 2 4 2 3 Vertical Tan Drags cerrado or e teveenssvescensteetpsapeeress 87 242A Canard Drain o E E TE E E i ie EE E EE E E 88 242 30 N Fail Drain 88 2426 Ventral Pin Drag oca iia 88 24 2 7 Fuselage Drag canica 88 2ZA2 8B Nacelle Dra Errorea errereen ieseana adh catethcesterdsshas SEESE s oTo E ESTS 88 242 9 TA 89 2 42 10 Floats uml 89 2 4 2 11 Trailing Edge Flap Drag noens a eN RSE 89 2 4 2 12 Leading Edge Flap Drag ocooonccnocnnocnnccnoncconoconoconcconocononononnnonncnnonanono 89 2 4 2 13 Landing Gear Dr gen issie och cotetbsesseresssasistveststevebestebseesoesy 89 DAD IA Canopy Draw ateo bis 90 2 4 2 15 Windshield Drato aa inina n a nono nc nn nono nono N A a 90 24 2 16 Stores Drag iois eire eener e eE EEE E E E eso 90 ZAZIT TAM DEB oere na aaa eE E EEEE E e iaoe 90 24 218 Spouler DA ro iii 91 DADO Speed Brake cestas 91 2 4 2 20 Miscellaneous Drag ciinii isien enore erei ees ne eases 91 2 42 21 Pylon Dag inci 91 2 4 2 22 W
62. ab hingemoment derivative input output window the user must define the lifting surface tip shape from the following options in the Control Surfaces Configuration dialog box See Figure 2 18 e Elliptic Tip e Square Tip For a horizontal stabilizer vertical tail wing canard with rounded or elliptic tips For a horizontal stabilizer vertical tail wing canard with square tips Once the desired lifting surface tip shape is selected the user may define the airfoil nose shape from the following options e Round Nose e Elliptic Nose e Sharp Nose e Round Low Drag Dynamics For control surface with a rounded nose For control surface with an elliptic nose For control surface with a sharp or pointed nose For control surface with an aerodynamically shaped low drag nose I 147 Once the nose shape is selected the user must define the type of balance for the control surface from the following options e Nose Balance Control surface is balanced solely with the nose e Nose Horn Balance Control surface is balanced with a horn or a combination of horn and nose If the control surface is defined with a horn i e Nose Horn Balance the user must define the horn type from the following options e Unshielded Horn For a horn that is not protected from the free stream air e Fully Shielded Horn For a horn fully protected from the free stream air e Part Shielded Horn For a horn partially protected from the free stream air Final
63. ail vertical tail V tails and tailbooms e Empty Weight With this option empty weight components and the associated center of gravity coordinates can be tabulated The center of gravity of the empty airplane will be computed using only the components for which all data is supplied e Total C G This allows the user to calculate the Center of Gravity Location of the airplane for the specified loading scenario e Passengers This allows the user to specify passenger groups and calculates the C G of each group e Forward Aft C G This module calculates the most forward and the most aft C G of the airplane based on minimum and maximum weights specified by the user for each component It also plots the C G envelope 1 5 8 Inertias This option allows the user to calculate the airplane moment of inertias by entering components and their corresponding weights moments of inertia and center of gravity locations in a preformatted table Also the moments of inertia for the wing horizontal tail vertical tail V tail canard fuselage tailbooms stores nacelles and floats are calculated based on geometric and structural parameters 1 5 9 Set Category If the user wishes to use Class Il Weight estimation methods other than those chosen by the program the Set Category option can be used Selecting the Set Category button will display a dialog box with options to choose any of the categories described below The following is a recommendation
64. ailable 6 6 10 Differential Stabilizer Related Derivatives The methodology used to calculate the differential stabilizer related derivatives can be found in Section 10 3 7 of Reference 7 Once the Diff Stabilizer button is selected a new menu appears with the following three differential stabilizer related derivative options a Cin II 146 In preliminary design the side force coefficient due to differential stabilizer derivative is assumed to be negligible and will therefore Dynamics not be calculated The program sets this derivative equal to zero and no additional menus appear on the screen This submodule can be used to estimate the rolling moment coefficient due to differential stabilizer derivative This submodule can be used to estimate the yawing moment coefficient due to differential stabilizer derivative For the estimation of this derivative no reliable preliminary design methods are available 6 7 Hingemoment Derivatives The Hingemoment submodule is used to determine the hingemoment coefficient derivatives of the elevator rudder aileron canardvator ruddervator elevator tab rudder tab aileron tab canardvator tab and ruddervator tab Note that the elevon and elevon tab hingemoments are calculated in the aileron and aileron tab hingemoment module The methodology for the hingemoment calculations is found in Section 10 4 of Reference 7 Before accessing any control surface and trim t
65. airplane moment of inertia about the body axis To select a different airplane category 1 5 1 Structure Component Weight Estimation The airplane structure weight consists of the following components wing empennage fuselage landing gear and nacelles In this module the structure component weights can be estimated with the available methods described in Reference 5 The structure weight components menu consists of the following options e Wing e Horizontal Tail e Vertical Tail e V Tail e Canard e Fuselage e Tailboom e Nacelle e Landing Gear e Total Structure To estimate wing weight for strut braced or cantilever wings To estimate the horizontal tail weight To estimate the single or twin vertical tail weight To estimate the v tail weight To estimate the canard weight To estimate the fuselage weight To estimate the tailboom weight To estimate the nacelle weight To estimate the landing gear weight O Nose Gear To estimate the nose gear weight O Main Gear To estimate the main gear weight O Tail Gear To estimate the tail gear weight O Outrigger Gear To estimate the outrigger gear weight o Total Gear To estimate the total gear weight To calculate the total structure weight Aerodynamics 1 5 2 Powerplant Component Weight Estimation The airplane powerplant weight consists of the following components engines air induction system propellers fuel system and propulsion system In this m
66. al Internal loads module calculates all forces and moments acting on the structural component at each of the load stations defined in the Load Stations module All external loads selected to act on the structural component must be defined in the Conc Weights Dist Weights Aerodynamic User Loads or Gear Engine Misc loads modules Total Internal loads must have already been calculated for all surfaces for which surface reactions have been selected The internal loads are calculated in the Libove Coordinate System Horizontal surfaces are split into right and left halves The Total Internal loads table is flight condition dependent and will be reset for each new flight condition H 178 Loads 8 4 2 Load Factors The accelerations acting on the aircraft structure are found by solving the airplane equations of motion The kinematic and force equations in the airplane stability axis system are solved for the airplane linear accelerations angular rates and angular accelerations in the body fixed axis system By solving the angular rate and angular acceleration equations first the force equations can then be solved for the linear accelerations This module must be completed before Total Internal loads can be calculated for any structural component Loads II 179 TT 180 Structures 9 Structures Module 9 1 General Description The purpose of this module is to estimate the size and weight of the structural components This 1s done using mat
67. an be calculated for a specific flight condition The wing drag coefficient consists of a zero lift component and a lift dependent component The methodology used to calculate the wing drag coefficient is described in Section 4 2 of Reference 7 The Wing Drag estimation method depends on the selected flow regime 2 4 2 2 Horizontal Tail Drag In this module the horizontal tail drag coefficient can be calculated for a specific flight condition The horizontal tail drag coefficient consists of a zero lift component and a lift dependent component The methodology used to calculate the horizontal tail drag coefficient is described in Section 4 4 of Reference 7 The Horizontal Tail input output windows depend on the selected flow regime 2 4 2 3 Vertical Tail Drag In this module the vertical tail drag coefficient can be calculated for a specific flight condition The vertical tail drag coefficient consists of a zero side force component and a side force dependent component The methodology used to calculate the vertical tail drag coefficient is described in Section 4 4 of Reference 7 The Vertical Tail input output windows depend on the selected flow regime Aerodynamics Il 87 2 4 2 4 Canard Drag In this module the canard drag coefficient can be calculated for a specific flight condition The canard drag coefficient consists of a zero lift component and a lift dependent component The methodology used to calculate the canard drag coefficien
68. and Turboprop only The user can choose a reversible or non reversing propeller Controls Configuration Dialog Box When the Controls button is selected the Control Surface Configuration dialog box is displayed The dialog box contains different pages for wing horizontal tail vertical tail v tail and canard The user can display different pages by clicking on the corresponding tab on the lower left corner of the dialog box Figure 2 18 shows the Control Surface Configuration dialog box with the wing and horizontal tail page displayed respectively The canard vertical tail and v tail pages are similar to that of the horizontal tail PART I I 31 Control Surfaces Configuration k Control Surfaces Configuration Aileron Surface Tip Shape Elevator Surface Tip Shape C Yes C No C C Yes C No f E E Variable Incidence Type of Balance C Yes C No Type of Balance Trim Tab z Trim Tab 3 c E C C c Elevon Type of Horn Balance Differential Stabilizer Type of Horn Balance C Yes C No z C Yes C No 4 c G Nose Shape Nose Shape ie E of Horn Balance Shape 9 Horn Balance Shape C c g c E 3 Horizontal T ail Vertical Tail Horizontal Tail Vertical Tail X Cancel 7 Help X Cancel 7 Help Figure 2 18 Control Surface Configuration Dialog Box The dialog box displayed in Figure 2 18 allows the user to define the appropriate control surfaces for the airplane i e aileron elevator canardvator ruddervator and rudder Respective control
69. andimillin g Dro 91 ZAD 23 Inlet cara da ide 92 2 4 2 24 NOZZlE iia 92 2ZA2 25 Total Dra iia ion cisnes 92 2 4 2 26 Recalculate Ally ccoo tasas 92 24221 NN 92 Z 42 28 Ploti aio oil 92 2 4 3 Estimation of the Wind Tunnel Drag oooocnnocnnoninoncncnnonnnanonanonn nana nonocnocnocnnos 94 Moment rrna e aoa p a E EAN EEAS AREE a A EE IRERE E 94 2 5 1 Pitching Moment Coefficient at Zero angle of attack ooncnncnnnnnnnnnmmmmm 95 2 5 2 Zero Lift Pitching Moment Coefficient ee eee eee ceeeeeeeeereeseenseeeaee 95 TI 23 3 Trailing Edge Flapss 4 icicacecectiite sense a Al Wa oca ote conecto 95 2 54 Leading Edge FlapSiuiiiain chen anaes eae SR 95 2 5 5 Total Pitching Moment Coefficient at Zero Angle of Attack ee 96 2 6 Aerodynamic Contract o it etiueles a 96 2 6 1 Aerodynamic Center Shift due to the Fuselage oooncnnnnnnnnncnncnnonnnoncconccnnos 97 2 6 2 Aerodynamic Center Shift due to the NacelleS ooononnonnnnnncnncnnccnocnconnconocnnos 97 2 6 3 Aerodynamic Center Shift due to Stores 0 0 eee eeeceeeeeeeeeeeeseenseeeaes 98 2 6 4 Aerodynamic Center Shift due to Tailbooms ooooonnncnccnnccnnoconnnoncnonaconocanocnnos 98 2 6 5 Aerodynamic Center Shift due to Floats ononccnnnnnncnnccnncncnnnoncnonaconoconocnnos 98 2 6 6 Aerodynamic Center Shift due to PylOns oo eee essen eeeeeeeeenseenseeeaes 98 2 6 7 Aerodynamic Center Shift due to Power EffectS ooooocnnnnnncnnccnocnnonnconocnnos 98 2 6 8 Wing
70. ass I submodule the volume of fuel that the wing can hold is calculated It is assumed that the volume between the front and rear spars is used as the useful fuel volume By multiplying by the density of the fuel the fuel weight can also be found The program will compare this weight to the maximum fuel weight in the fuel tank at any point in the mission to determine if this amount of fuel can be stored in the wing The Class II submodule calculates the allowable fuel in the wing by taking into account the following e Rear Spar Location e Rear Spar Location e Inboard and Outboard Tank Stations e Factor to account for structure e Factor to account for fuel expansion The Cranked Wing submodule calculates the allowable fuel in the different panels of the wing It takes into accout the following e Rear Spar Location e Rear Spar Location e Inboard and Outboard Tank Stations e Factor to account for structure e Factor to account for fuel expansion e Thickness ratios of the panels at their root and tip 4 3 4 High Lift Device In this submodule the geometry of the high lift devices is specified Il 114 Geometry 4 3 5 Aileron Tab In this submodule the geometry of the aileron is added 4 3 6 Spoiler In this submodule the geometry of the spoiler is added 4 3 7 Chord Length In this submodule the chord length of the wing at any span wise location can be determined by using root chord length taper ratio and the given
71. ates a specific plot the software calculates the plot area to encompass the entire graph The values defining the first plot area will be saved and used the next time the plot is generated If the axes are changed to redefine the plot area those parameters will be saved and used the next time the graph is generated The parameters are saved for every plot that can be generated in the software The parameters are saved in the user database The user can always have the program recalculate the plot area to show the entire graph by selecting the Default button in the plot window command bar Change Axis Minimurn 0 0000 Maximum 12 0000 Divisions 4 Subdivisions 5 Decimals 14 Same for All Flight Conditions Y OK X Cancel 7 Help Figure 2 8 Change Axis Dialog The functionality of the plot window command bar buttons is described in the next subsection PART I I 13 2 1 5 Plot Window Command Bar The plot window command bar is displayed at the top of the plot window The plot window command bar is shown in Figure 2 9 Each button in the command bar represents an action that can be performed in the plot window A command bar button is not displayed if its action is not available for the particular plot window The Close button in the command bar closes the plot window and is always displayed The remaining buttons that can be displayed in the plot command bar are shown and described in Table 2 3 Figure 2 9 Plot Window Command
72. ation Number of Fans C Not Buried no airflow from fuselage base C Buried in Fuselage airflow from fuselage base Cd X Cancel 7 Help Figures 2 17f Powerplant Selection Dialog Box for a Ducted Fan Propfan PART I I 29 Propulsion Engines Nacelles Propulsion Engine Type Yes C No Number 0 C Yes C No C Jet C Piston C Propeller C Propfan Ducted Fan y Engine Location Nacelle Location Wing Fuselage Other Wing Fuselage Other C C C C C C Fuel Tank C Integral C Self Sealing Bladder C Non Self Sealing Type of Inlet Prop Type C Plenum C Composite C Straight Through C Metal 1 2 3 4 5 6 7 8 9 omen Oo M0 bh QQ NN Prop Pitch Prop Reversing C Fixed E C Variable C elelee a e eleele eee e e 1 to 1 teitei taia lei teite we RT eite Tiel tel te elelee neee o Engine Air Flow Location Number of Fans C Not Buried no airflow from fuselage base C Buried in Fuselage airflow from fuselage base or Y OK X Cancel Help Figures 2 17g Powerplant Selection Dialog Box for a Ducted Fan Turboprop Engines The user can specify if there are engines Number of Engines For all engine types the user can define the number of engines Engine Location The user can specify where the engines are located Nacelles The user can specify if there are nacelles Number of Nacelles The user can define the number of nacelles Nacelle Location The user can specify where the nacel
73. avity of the empty airplane will be computed using only the components for which all data is supplied This option will be described in Section 1 4 2 1 e Total C G With this option the Center of Gravity Location of the airplane can be calculated for the specified loading scenario e C G Excursion This option allows the user to generate graphs displaying the calculated center of gravity excursion in the X Y and Z directions This option will be described in Section 1 4 2 2 1 4 2 1 Empty Weight This option allows the user to calculate the airplane empty weight and associated center of gravity location This module consists of two different submodules The first one From Fractions uses the weight components previously calculated in the Weight Fractions module The second One Detailed is divided into the following four options e Structure With this option the user can tabulate the structure weight components and center of gravity coordinates The overall weight and center of gravity of the structure will be calculated using only the components for which all data is supplied If a component weight and center of gravity location has been entered in the From II 68 Aerodynamics Fractions Empty Weight Table their values will automatically be entered in the Structural Weight Table e Powerplant With this option the user can tabulate powerplant weight components and the associated center of gravity coordinates The overall w
74. ayed see Figure 2 1 The application window contains menu button selections that allow the user to select a calculation to be performed The software uses a flow chart method for the user interface as shown in Figure 2 1 This allows the user to see the path selected in reaching a certain location The software consists of various calculation modules that can be accessed through the application windows Table 2 1 presents the application buttons in the main window and the calculation modules accessed by that application module Table 2 1 Application Modules of the Program Application Button Calculation Modules Weight e Class I take off weight and fuel calculation e Class Land Class II weight amp balance analysis and center of gravity calculation for current loading Aerodynamics Class I wing and high lift devices design Class I lifting surface and airplane lift calculation Class I and Class II drag polar calculation Lift drag and moment distributions over a lifting surface Airplane aerodynamic center calculation Power effects on airplane lift and pitching moment Ground effects of airplane lift and pitching moment Dynamic Pressure Ratio Performance Class I performance sizing Class II performance analysis PART I I 5 Table 2 1 Continued Geometry e Class I wing fuselage and empennage layout e Aero Pack Interface e Lateral tip over analysis e Scale Propulsion e Class I installed thrust power cal
75. be used to determine the change in pitching moment and wing lift coefficient at zero angle of attack due to a trailing edge flap deflection Before the Trailing Edge submodule can be selected the user must define the desired trailing edge flap type from the Airplane Configuration dialog box 2 5 4 Leading Edge Flaps This submodule can be used to determine the change in pitching moment due to a leading edge flap deflection Before the Leading Edge submodule can be selected the user must define the desired leading edge flap type from the Airplane Configuration dialog box See Section 2 2 3 Figure 2 16 Aerodynamics Il 95 2 5 5 Total Pitching Moment Coefficient at Zero Angle of Attack This submodule can be used to determine the airplane pitching moment coefficient at zero angle of attack Before the Total submodule can be selected the user must define one of five types of airplane configurations in the Airplane Configuration dialog box See Section 2 2 3 Figure 2 16 The leading and trailing edge flap type and contributions must also be defined as presented in Sections 2 5 2 and 2 5 3 The total airplane pitching moment coefficient at zero angle of attack is calculated in components The effect of wing fuselage horizontal tail v tail and canard are included in the total pitching moment coefficient 2 6 Aerodynamic Center The purpose of the Aerodynamic Center submodule is to determine the Aerodynamic Center of the airplane The meth
76. button is selected a new menu appears with the following three ruddervator related derivative options e Cp 5 This submodule can be used to estimate the drag coefficient due rv to ruddervator deflection derivative e Cr 5 This submodule can be used to estimate the lift coefficient due to rv ruddervator deflection derivative e Crn 5 This submodule can be used to estimate the pitching moment rv coefficient due to ruddervator deflection derivative 6 5 9 Ruddervator Tab Related Derivatives The methodology used to calculate the ruddervator tab related derivatives is based on theory in Sections 10 3 2 and 10 3 8 of Reference 7 When the Ruddervator Tab button is selected a new menu appears with the following two ruddervator tab related derivative options e CLs This submodule can be used to estimate the lift coefficient due to rvt ruddervator tab deflection derivative e Cn 5 This submodule can be used to estimate the pitching moment rvt coefficient due to ruddervator tab deflection derivative 6 5 10 Flying Wing Related Derivatives The methodology used to calculate the flying wing related derivatives is based on theory in II 140 Dynamics Section 10 2 2 1 of Reference 7 Once the Flying Wing button is selected a new menu appears with the following three flying wing related derivative options e Cp This submodule can be used to estimate the drag coefficient due w to flying wing incidence derivative e C
77. calculate the horizontal tail volume coefficient and related parameters 1s described in Chapter 8 of Reference 2 After calculating the output parameters the user may display the horizontal tail geometry by selecting the Plot option 4 4 4 Elevator Tab In this submodule the geometry of the elevator is added 4 4 5 Chord Length In this submodule the chord length of the horizontal tail at any span wise location can be determined by using root chord length taper ratio and the given span wise location 4 4 6 Airfoil In this submodule the horizontal tail airfoil parameters at the horizontal tail mean geometric chord are calculated 4 4 7 Exposed In this submodule the exposed horizontal tail parameters are calculated 4 5 Vertical Tail Geometry When the Vertical Tail module is selected the user is presented with the following options TY 116 Geometry 4 5 1 Straight Tapered For calculation of straight tapered vertical tail geometry 4 5 2 Cranked For calculation of the equivalent straight tapered vertical tail geometry from given cranked vertical tail geometry 4 5 3 Volume Coefficient For calculation of the vertical tail Volume Coefficient 4 5 4 Rudder Tab For calculation of the rudder geometry 4 5 5 Chord Length For calculation of vertical tail chord based on span wise location 4 5 6 Airfoil For calculation of airfoil parameters at the vertical tail mean geometric chord 4 5 7 Exposed For calculation of expo
78. cccccececececececececececececececececeseseseseseseseeesees 142 6 6 2 Aileron Tab Related Derivatives ce eceesceeseeceenceceeeceeceeceeeeeeeecsaeeeeneeees 143 6 6 3 Spoiler Related Derivatives ooononnncnnnnncnnonnconnconcconccnnorononnncnnnranonnncnnncnnnnns 143 6 6 4 Vertical Tail Related Derivatives cee eeeccesseeeneeceeeeeeneeceereeeeeceteeeeneeens 144 6 6 5 Rudder Related Derivatives 20 00 00 eesceeeceeseeceseceeseeceeeeeeneeceeeeeeneeceeeeeenaeens 144 6 6 6 Rudder Tab Related Derivatives ooooonccnnoncconcccnonnconcnononaconoconnncnonccrnnnnnnnccnno 144 6 6 7 V Tail Related Derivatives ceesceceseceenceceseceeceeceseceeaeeceaeeeeeeecsaeeeeneeee 145 6 6 8 Ruddervator Related Derivatives 0 0 0 eeeescecsseceececeseceeneecereeeeecetreseneeees 145 6 6 9 Ruddervator Tab Related Derivatives oooonnccononococcccooncconcnonnnanoncconnnnnnnccnno 146 6 6 10 Differential Stabilizer Related Derivatives ee ceeeesseceeeeeeneeceteeeeeeees 146 6 7 Hing moment Derivatives ii rot a A 147 6 3 Recalculate All ici ici iii 148 6 9 Class I Stability and Control Empennage Sizing Analysis oooonnccnccnnccnncccnanconoconocnnos 148 6 9 1 Static Longitudinal Stability cece cee esse cneeceeeeeeeeeeseeeeeeeeereeeeens 149 6 9 2 Static Directional Stability oooonncnnnninnnocnconcconccononanonnnonnnrnn cnn nono c ran crnncnnnono 150 6 9 3 Minimum Controllable Speed with One Engine Inoperative Stability 151 6 10 Class IT
79. ccessively subtracted from the airplane weight according to the unloading scenario until the empty weight is reached Each component 1 up to 13 can be used only once in each of the Aerodynamics II 69 loading unloading columns of the table Once the loading scenario is provided and the center of gravity travel is calculated the C G excursion diagrams in the X Y and Z directions can be displayed with the Plot selection 1 4 3 Inertia Estimate This option allows the user to estimate the moments of inertia of the airplane with the radii of gyration method The user can provide the data or the preloaded database can be used The user has the option to select data from comparable airplanes By selecting a category a choice can be made between several available airplanes appearing in a table Because the Class I inertia estimation is based on the selected aircraft the user should choose the most comparable aircraft The user may add an airplane more than once to weigh the importance of a particular configuration To deselect an airplane the user can remove the airplane from the current selection list The Inertia Estimation module contains the options Add Airplane Select Airplane and Moment of Inertia which will be explained here e Add Airplane Allows the user to add inertia data for new airplanes to the already available airplanes appearing in the table The user has to provide the data for the Radius of Gyration Table This data is aut
80. cking on the Setup tab 2 2 1 Main Toolbar The main toolbar Figure 2 10 consists of seven fixed bitmap buttons at the bottom right of the main window The main toolbar supplies general functions needed throughout the program The functionality of the buttons in the main toolbar is described in Table 2 4 Figure 2 10 Main Toolbar Table 2 4 Toolbar Buttons Flight Condition Set and define each flight condition to be included in the analysis An airplane project can have up to 95 flight conditions defined Recalculate Calculates the output parameters of different modules for selected flight conditions Recalculate PART I I 15 Table 2 4 Contd Toolbar Buttons Notes Record general notes about the current project Notes are saved with the project Copy WMF Copy a representation of the active window into the clipboard in Windows Metafile Format The contents of the clipboard can then be pasted into a word processing or drawing program that supports Windows Metafiles The contents can also be saved to files if the Copy WMF to File option in the Program Options dialog box Figure 2 31 in the Setup toolbar is checked Print Make a hard copy of the data currently displayed on the screen on the selected printer Printing options are described in Chapter 4 Atmosphere Display an input output window for calculation of properties of the standard atmosphere at a given altitude and temperature offset The module also cal
81. cture button is selected the Structure dialog box is displayed The user can choose from the five different structure types as shown in the pull down list in Figure 2 21 for each surface The user can also specify if the surface is attached to the fuselage PART I I 35 Structure Amphibious Hull C Yes No Pressurized C Yes No Wing Type Cantilever Monoplane Strut Braced Monoplane Attached to Fuselage C Yes C No Structure z Canard Attached to Fuselage C Ye C N NO Structure as T ail Attached to Fuselage C Yes C No Structure AA Single spar Semimonocoque Two spar Semimonocoque Two spar Drag Truss Single spar Box Single spar Circular Vertical Tail Attached to Fuselage C Yes No Structure z Structure eee X Cancel Figure 2 21 Lifting Surface Cross Section Structure Dialog Box 2 2 4 Certification Toolbar The Certification toolbar Figure 2 22 consists of two bitmap buttons at the bottom of the main window The Certification toolbar buttons can be used to specify airplane certification type and class The functionality of the buttons on the Certification toolbar is described in Table 2 7 Figure 2 22 Certification Toolbar PART I The functions of the buttons shown in Figure 2 22 are described in Table 2 7 Table 2 7 Certification Toolbar Buttons military regulations Certification Define the airplane type category and certification under civ
82. culate the empty weight take off weight mission fuel weight used fuel weight mission fuel fraction fuel burned in each flight segment and the fuel available and airplane weight at the beginning of each flight segment will be calculated and displayed The calculation leads to one of the following cases 1 No solutions for take off weight 2 One solution for take off weight 3 Two solutions for take off weight The mathematical background for the above stated scenarios can be found in Appendix A of Reference 1 A plot function is available to show the iteration process used to arrive at these solutions For case 1 the input data must be changed to obtain a solution For case 2 no modifications are necessary If case 3 occurs the program automatically chooses the lowest weight solution 1 3 3 L D from Weights This command calculates the climb cruise turn and loiter lift to drag ratios based on II 64 Aerodynamics corresponding segment weights It iterates between Weight Sizing and Class I Drag since the wetted area is calculated from take off weight to obtain a converged solution for the take off weight 1 3 4 Regression This command is used to calculate the regression coefficients A and B by supplying empty weights and take off weights of actual airplanes in a table This option allows the user to generate a table of empty weights and take off weights The number of take off and empty weight combinations maximum 50
83. culates Mach number and Reynolds number per unit length Help Display the help system associated with the software Chapter 6 describes the help system Exit Exit the program The Flight Condition button displays the Flight Condition Definition dialog box which is shown in Figure 2 11 PART I PART I E i Flight Condition Definition ua Flight Phase Name Flight Condition 1 El oR New Edit Delete Gi Move Copy 7 Stores Engines Operating Trim Surface Speed Brake Altitude ft On Off On Off Stabilizer Retracted 16 c 10 e C Elevator C Deployed re te G pe 4c G AT deg F an cy fo c 3 0 c C Canard Spoiler Retracted ac E ac B C Canardyator U kts 5e e se E C Deployed C Y Tail 6 C e 65 c A Gear y Not Defined iea a E Ruddervator Cup 8 C c gc C C Elevon C Down c c c c Weurrent lb E a Drag Polar C G Location waits E 10 a Class C Forward Xog ft Class Il C Aft Engine Rating had C Wind Tunnel C Other Zeg ft Include in Excel Export Power Effects Ground Effects C On C On F Include in APP Export n g C Off C Off W Include in Recalculate A g Flight Condition Flying Qualities Category Notes X Cancel Help b S Figure 2 1la Flight Condition Dialog Box a Flight Condition Definition a Flight Phase Name Flight Condition 1 z LA New Edit Delete Gi Move Copy Flight Phase
84. culation e Inlet Nozzle sizing Stab amp Control e Longitudinal and lateral directional stability and control derivatives including thrust power e Control surface and trim tab hinge moment derivatives e Class I stability amp control empennage sizing e Class II longitudinal and lateral directional trim including stick force and pedal force calculations Dynamics e Open loop dynamics analysis e Automatic control system analysis Loads e Velocity Load Factor V n diagram generation e Structural component internal load estimation Structures e Material property tables e Class I component structural sizing Cost e Airplane program cost estimation Clicking on the appropriate buttons in the application window activates each module When the menu buttons leading to a calculation module have been selected the input output window for that calculation module is opened 2 1 2 Input Output Windows The input output window opens after selecting the type of calculation to be performed The input output window contains numeric data necessary to perform a calculation For some I 6 PART I calculations information about the airplane configuration and airplane certification type are required so that the correct calculation method can be used Before the input output window is displayed the program will display a dialog box allowing the user to specify configuration choices For example the program will ask the user t
85. culation of spanwise moment distribution on the canard The user can display it by selecting the Plot button For calculation of spanwise moment distribution on the v tail The user can display it by selecting the Plot button For calculation of Airplane pitching moment coefficient as well as moment derivatives for the airplane Aerodynamics 2 5 1 Pitching Moment Coefficient at Zero angle of attack This submodule can be used to determine the zero lift airplane pitching moment coefficients and pitching moment coefficients at zero airplane angle of attack This module includes flap effects To account for each airplane component the pitching moment coefficient components are calculated separately The user is presented with the following options 2 5 1 Zero Lift For estimation of the zero lift pitching moment coefficient 2 5 2 Trailing Edge For estimation of the effect of trailing edge flaps on lift and pitching moment coefficients 2 5 3 Leading Edge For estimation of the effect of leading edge flaps on lift and pitching moment coefficients 2 5 4 Total For calculation of the total pitching moment coefficient at zero angle of attack 2 5 2 Zero Lift Pitching Moment Coefficient This submodule can be used to determine the zero lift airplane pitching moment coefficient The coefficient is separated into a wing and fuselage component and the combined wing and fuselage effect is calculated 2 5 3 Trailing Edge Flaps This submodule can
86. d to perform computations related to the static longitudinal stability of an airplane The methodology used to calculate the static longitudinal stability can be found in Section 11 1 of Reference 2 When the Longitudinal submodule is selected another menu appears with three stability types These stability types are e Inherent Represents inherent stability which is required of all airplanes that do not rely on a feedback augmentation system for their stability The user has two options in this submodule O Static Margin For calculation of static margin for a given canard and horizontal tail area o Surface Area For calculation of canard and horizontal tail area for a given static margin e De Facto Represents de facto stability which is required of all airplanes that are stable only with a feedback augmentation system in place If the static margin exceeds the value of 0 05 no feedback augmentation system is required and the feedback gain in this condition will be set to zero e Volume Method Represents the sizing method based on volume coefficients The user has two options in this submodule O Quarter Chord For sizing the horizontal tail canard or v tail based on lifting surface position in reference to the wing Dynamics H 149 O A C C G For sizing the horizontal tail canard or v tail based on aerodynamic center location Before any of the above options can be selected the user must define one of three types of airplane config
87. d to trim e F Aileron stick or wheel force e F Rudder pedal force The remaining five variables not defined by the user are calculated and included in the output 6 10 4 T O Rotation The T O Rotation submodule can be used to determine the horizontal tail v tail canard area and lift coefficient required for take off rotation based on landing gear and center of gravity Dynamics I 153 locations as well as pitch attitude acceleration requirements 6 10 5 Trimmed Lift T from D This module is used for the determination of trimmed lift coefficients 6 10 6 Trimmed Lift T Const This module can be used to calculate the trimmed lift coefficients on the canard wing horizontal tail and v tail of an airplane at a given flight condition while accounting for flap effects This module allows the user to quickly determine the trimmed condition of the airplane Before the Trimmed Lift submodules can be selected the user must define one of five types of airplane configurations in the Airplane Configuration dialog box and trailing edge flap type in the Wing Configuration dialog box See Section 2 2 3 Figure 2 16 6 11 Stick Free Static Margin In this module the stick free static margin is calculated 6 12 Stick Force In this module the stick force and the stick force speed gradient are calculated 6 13 Aileron Force In this module the aileron wheel force is calculated 6 14 Rudder Force In this module the
88. d with the left mouse button a vertical insertion bar appears in the edit box and the keyboard can be used to type numeric input When the cursor is positioned outside the edit box it appears as a small calculator When the left mouse button is clicked while the cursor appears as a calculator the program calculator is opened see Section 3 2 When the Info button is clicked an information window is displayed for that variable The information window contains a definition of the variable with graphics and suggested values when available see Section 6 2 Figure 2 4 shows the Notes window When the Work Pad button is clicked this window is displayed and allows the user to type notes about that variable These notes are specific to that variable and will be saved with the project Notes may also be designated one of six colors to identify certain stages of the design process This is done by simply clicking on the desired color I 8 PART I in the Set Current Note Color box of the Work Pad Window If a color is not selected from this portion of the Work Pad Window the default color will be used with that particular note If notes have been entered for a variable the Work Pad button will change colors to the default notes color The default notes color can be set or changed in the Program Options window See Figure 2 30 The Work Pad Window also has options to allow the user to lock the value of the variable so that it does not get r
89. dialog box or by selecting the Print Parameters options in the Print dialog box 7 3 2 Longitudinal Flying Qualities Flying Qualities performs a check on the calculated longitudinal modes against the relevant regulations The flying qualities of the airplane are compared with the flying qualities as specified in the regulations chosen by the user in the Certification and Type dialog box Where no specifications are available for FAR part 23 or FAR part 25 MIL F 8785C specifications are used If the user chooses MIL Specs as the certification one of the following options must be chosen e Land Based For land based airplanes e Carrier Based For aircraft carrier based airplanes e Land amp Carrier For land and carrier based airplanes After selecting the airplane certification type the flight phase category must be defined in the Flight Condition Definition dialog box The user has the following options e Flight Phase A Non terminal flight phases that require rapid maneuvering precision tracking or precise flight path control Dynamics Il 157 e Flight Phase B Non terminal flight phases that are normally accomplished using gradual maneuvers and without precision tracking although accurate flight path control may be required e Flight Phase C Terminal flight phases that are normally accomplished using gradual maneuvers and usually require accurate flight path control The user also needs to define the airplane class from the foll
90. e Performance module see Section 3 or the user may estimate them The Plot option is used to display the wing geometry including the size and location of the high lift devices The program determines whether the take off condition or landing condition sizes the high lift device e Lift To allow the user to verify whether the selected high lift devices can produce the desired trimmed airplane lift coefficient for a specific flight condition 2 4 Drag After selecting Drag the following options are displayed e Class See Section 2 4 1 to use the Class I method e Class Il See Section 2 4 2 to use the Class II method e Drag Distribution The Drag Distribution can be used to calculate spanwise drag on the wing horizontal tail vertical tail canard and v tail To determine the spanwise drag distribution select the lifting surface button An input output window appears showing the surface planform input parameters The user can display the drag distribution by selecting the Plot button e Critical Mach The Critical Mach submodule can be used to calculate the Critical Aerodynamics II 83 Mach number for the airplane 2 4 1 Estimation of Class I Drag Polars The Class I Drag module has the following eight options e T O Gear Down For flaps in take off configuration with gear down e T O Gear Up For flaps in take off configuration with gear up e Clean For configuration with no flap deflection and gear up e Land Gear Up For fla
91. e The file can then be saved as a PART I I 57 project in this version This version does not support saving projects as files for use with previous versions Files saved with Unix versions 1 0 through 1 7 of AAA can be opened by choosing the ana analys option in the Files of Type box in the File Open dialog box described in Subsection 2 2 2 5 3 Projects and Their Databases A project consists of a project file and multiple databases for flight conditions and airplane components While the software is running the current project is stored in the folder lt dar folder gt Working which is called the Working Directory When the software is started with a new project the default name is projectl The new project is kept in the Working Directory until it is saved for the first time When the project is saved all of the files in the Working Directory are copied into one AAA project file specified by the user in the Save As dialog box see Subsection 2 2 2 Projects can be moved by using the Save As button in the File Management toolbar or the File Save As menu command When an existing project is opened the project is first copied to the Working Directory where it is used by the program When the user saves the project it is copied from the Working Directory back into its original folder The program has been set to automatically save the recovery file in the Working Directory every 5 minutes The user can alter the interval in the
92. e C Yy In preliminary design the side force coefficient due to ruddervator derivative is assumed to be negligible and will therefore not be calculated The program sets this derivative equal to zero and no additional menus appear on the screen Dynamics H 145 This submodule can be used to estimate the rolling moment coefficient due to ruddervator derivative This submodule can be used to estimate the yawing moment coefficient due to ruddervator derivative For the estimation of this derivative no reliable preliminary design methods are available 6 6 9 Ruddervator Tab Related Derivatives The methodology used to calculate the ruddervator tab related derivatives is based on theory in Sections 10 3 2 and 10 3 8 of Reference 7 Once the Ruddervator button is selected a new menu appears with the following three ruddervator tab related derivative options z rot eC rvt n ryt In preliminary design the side force coefficient due to ruddervator tab derivative is assumed to be negligible and will therefore not be calculated The program sets this derivative equal to zero and no additional menus appear on the screen This submodule can be used to estimate the rolling moment coefficient due to ruddervator tab derivative This submodule can be used to estimate the yawing moment coefficient due to ruddervator tab derivative For the estimation of this derivative no reliable preliminary design methods are av
93. e Nacelle input output window the user is required to input the number of nacelles If the propeller driven option has been selected then the supersonic flow regime cannot be used T 88 Aerodynamics 2 4 2 9 Tailboom In this module the tailboom drag coefficient can be estimated The methodology used to calculate the tailboom drag coefficient is described in Sections 4 3 and 4 13 of Reference 7 The tailboom input output window depends on the selected flow regime Before entering the tailboom input output window the user is required to define the number of tailbooms 2 4 2 10 Floats In this module the float drag coefficient can be estimated DARcorporation developed the methodology used to calculate the float drag The float drag estimation is only valid in the low subsonic speed regime Before entering the float input output window the user is required to define the number of floats 2 4 2 11 Trailing Edge Flap Drag In this module the drag coefficient of wing trailing edge lift generating devices can be estimated The assumption is made that flap deployment will occur only in the subsonic speed regime The flap drag estimation is therefore only valid in the subsonic speed regime The methodology used to calculate the flap drag coefficient is described in Section 4 6 of Reference 7 2 4 2 12 Leading Edge Flap Drag In this module the drag coefficient of wing leading edge lift generating devices can be estimated The assumption
94. e The Flight Condition List lets the user specify the number of flight conditions that appear in the Flight Condition drop down menu in the Flight Condition dialog box e Default Note Color Hint This allows the user to define which color is the default when notes are entered for a variable It also allows the user to associate a hint with each color to help identify certain stages of the design process e Variable Changed Default Color This option allows the user to define which color the variables backgrounds will change to if their values changes when the calculate button is used e Export Flight Condition This option allows the user to export flight conditions to a spreadsheet either in rows or in columns e Default Angle Units This allows the user to define whether degrees or radians will be the default angle unit throughout the project e Default 1 Angle Units This allows the user to define whether degree or radian will be the default 1 angle unit throughout the project e Default Length Units When British units are in use this allows the user to define whether inches or feet will be the default length unit throughout the project When S I units are in use it allows the user to define whether millimeters or meters will be the default length unit throughout the project e Default Area Units When British units are in use this allows the user to define whether square inches or square feet will be the default area unit througho
95. e canard airfoil parameters at the canard mean geometric chord are calculated 4 6 7 Exposed In this submodule the exposed canard parameters are calculated 4 7 V Tail Geometry When the V Tail module is selected the user is presented with the following options 4 7 1 Straight Tapered For calculation of straight tapered v tail geometry 4 7 2 Cranked Canard For calculation of the equivalent straight tapered v tail geometry from given cranked v tail tail geometry 4 7 3 Volume Coefficient For calculation of the v tail Volume Coefficient 4 7 4 Ruddervator Tab For calculation of the ruddervator geometry 4 7 5 Chord Length For calculation of v tail chord based on span wise location 4 7 6 Airfoil For calculation of airfoil parameters at the v tail mean geometric chord Geometry TT 119 4 7 7 Exposed For calculation of exposed v tail parameters 4 7 1 Straight Tapered The method for determining the geometry of a straight tapered v tail is identical to the method found in Section 4 4 1 of this manual 4 7 2 Cranked V Tail The method for determining the geometry of a cranked v tail is identical to the method found in Section 4 4 2 of this manual 4 7 3 Volume Coefficient This v tail submodule calculates the volume coefficient of a v tail planform The methodology used to calculate the v tail volume coefficient and related parameters is based on theory in Chapter 8 of Reference 2 After calculating the output parameters t
96. e form of a performance matching plot These plots depend on the type of airplane the applicable specification and the applicable regulation s With the help of such a plot the combination of the highest possible wing loading and the smallest possible thrust or highest power loading which meets all performance requirements can be determined The methodology used for performance sizing can be found in Reference 1 The purpose of the Analysis submodule is to provide the user with Class II analysis methods for predicting the performance characteristics of an airplane The methodology used to analyze the performance characteristics can be found in Chapter 5 of Reference 8 Use of the performance module options will be described in the following sections Performance H 101 32 Performance Main Window After invoking the Performance module two options are displayed e Sizing The options provided in this module are discussed in Section 3 3 e Analysis The options provided in this module are discussed in Section 3 4 e Export to APP The options provided in this module are discussed in Section 3 5 3 3 Performance Sizing After choosing the Sizing submodule another menu appears with seven options The seven options represent submodules as follows 3 3 1 Stall Speed For sizing to stall speed requirements 3 3 2 Take off Distance For sizing to take off distance requirements 3 3 3 Climb For sizing to climb requirements 3 3 4 Max Cru
97. e fuselage geometry methods The user has to input the tailboom cross section data and define the sections of the tailboom After calculating the output parameters the user may display the tailboom geometry by selecting the Plot option 4 12 Float Geometry The methodology used to calculate the float geometry and float related parameters is based on the fuselage geometry methods The user has to input the float cross section data and define the sections of the float After calculating the output parameters the user may display the float geometry by selecting the Plot option 4 13 Landing Gear In this module the lateral angle between the airplane center of gravity and the landing gear This Geometry TI 121 angle is useful in determining the risk of lateral tip over on the ground It also calculates the angle between the critical gear ground contact angle and the empty weight center of gravity and the current center of gravity This angle is useful in determining the risk of longitudinal tip over on the ground 4 14 Canopy This option is used to define the canopy geometry 4 15 Store Geometry The methodology used to calculate the store geometry and store related parameters is based on the fuselage geometry methods The user has to input the store cross section data and define the sections of the store After calculating the output parameters the user may display the store geometry by selecting the Plot option 4 16 Pylon Geo
98. e in the specified flight condition e Engines Operating The user can indicate which engines are operating in that flight condition e Engine Rating The user may select the following options using this drop down list 1 Take off 2 Max Continuous 3 Max Cruise e The user may select any of the given options Include in Recalculate Include in Export to include the flight condition to the Recalculate Dialog window and to include the flight condition to Export to Excel option in the Recalculate Dialog window e Trim Surface The user can indicate which control surface is used for trim in the current flight condition e Canard Trim For Canard configurations the user might select a variable incidence canard or a canardvator for trim This selection can be done simultaneously with any PART I I 19 tail aft control surface Drag Polar The user can indicate which drag polar is to be used for Trimmed Lift T from D Longitudinal Trim Trim Diagram and Performance gt Analysis gt Maximum Cruise Speed modules Power Effects For propeller driven airplanes the user can indicate if power effects are ON or OFF for the flight condition Speed Brake After selecting speed brake in the Configuration dialog box the user can specify if the speed brake is retracted or deployed for the corresponding flight phase Spoilers After selecting spoiler in the Configuration dialog box the user can specify if the spoiler is re
99. e toolbar Figure 2 14 consists of five bitmap buttons at the bottom of the main window The File toolbar buttons can be used to manage projects and files of the software The functionality of the buttons in the File toolbar is described in Table 2 5 Figure 2 14 File Toolbar PART I I 23 Table 2 5 File Management Toolbar Buttons New Create a new project Open Open an existing project analys files from AAA Versions 1 0 through 1 7 and gpr files from AAA 2 0 through AAA 2 2 can also be opened Save Quickly save the current project under its current name and directory Files have an aaa extension Save As Save the current project under a different name and or folder The project is saved in a directory folder with the same name as the project Save As y Delete Delete Remove any Project Each of the buttons in the File Management toolbar opens the Windows dialog box corresponding to that function Descriptions of the dialog boxes can be found in the documentation for your version of Windows 2 2 3 Configuration Setup Toolbar The Configuration Setup toolbar Figure 2 15 consists of six bitmap buttons at the bottom of the main window The Configuration Setup toolbar buttons can be used to define the airplane configuration for the current project The functionality of the buttons on the Configuration Setup toolbar is described in Table 2 6 SES PNT A Configuration Powerplant Controls Flap S
100. ecalculated export the value to an ASCII text file or select whether or not the variable is flight condition dependent The Default Unit box in the Work Pad Window allows the user to change the units for the variable associated with the window without changing the default units for the entire project The number of decimals can be increased or decreased using the Number of Decimals box The Go To button appears next to parameters which have been calculated by AAA in another module Selecting the Go To button will display the module in which the corresponding parameter was calculated This allows the user to see what variables were used in producing the parameter and confirm its validity Clicking on the Go To button a second time will return the user to the previous module DEK Wing Span Undefine SUBritish Number of Decimals 2 gt F Lock Value Does not get Recalculated Default Unit Set Current Note Color a a a a m al F Allow Yalue to be Exported to File co V Same Value for All Flight Conditions X Cancel 7 Help E Print 7 Info Figure 2 4 Work Pad Window PART I I 9 An input output window can also contain a table for numeric input and output data see Figure 2 2 Rows can be added to or subtracted from certain tables to account for multiple inputs of the same form For example the fuselage can be divided into two or more sections for moment calculatio
101. ed as the powerplant e Cmr This submodule can be used to estimate the thrust pitching u moment coefficient due to speed derivative This option is not available if None is specified as the powerplant NOTE There are two methods of entering values of the aerodynamic center shift due to the fuselage nacelles stores floats and tailbooms The user may 1 Enter the values of the aerodynamic center shift due to the fuselage nacelles stores and tailbooms in terms of the wing mean geometric chord in the C submodule if these values are known 2 Use the Aerodynamic Center module to calculate the values for the aerodynamic center shift due to the fuselage nacelles stores floats and tailbooms in terms of the wing mean geometric chord If the user chooses to first use the Aerodynamic Center module the values for the aerodynamic center shift will automatically be transferred into the Cin submodule input For a complete description of the Aerodynamic Center module see Part II Section 2 6 of this user s manual 6 3 3 Angle of Attack Derivatives The methodology used to calculate the angle of attack related derivatives can be found in Section 10 2 2 of Reference 7 Once the Angle of Attack button is selected a new menu appears with the following angle of attack related derivative options e Cp This submodule can be used to estimate the drag coefficient due to angle of attack derivative H 132 Dynamics e Cr This
102. ee yaw rate related derivative options ec yy This submodule can be used to estimate the side force coefficient due to yaw rate derivative e C This submodule can be used to estimate the rolling moment coefficient due to yaw rate derivative e C This submodule can be used to estimate the yawing moment coefficient due to yaw rate derivative 6 5 Longitudinal Control Derivatives The Long Control submodule has six longitudinal control surface and trim tab derivative submodules These twelve submodules are II 136 Dynamics 6 5 1 Stabilizer For estimation of the stabilizer control derivatives 6 5 2 Elevator For estimation of the elevator control derivatives 6 5 3 Elevator Tab For estimation of the elevator tab control derivatives 6 5 4 Canard For estimation of the canard control derivatives 6 5 5 Canardvator For estimation of the canardvator control derivatives 6 5 6 Canardvator Tab For estimation of the canardvator tab control derivatives 6 5 7 V Tail For estimation of the v tail control derivatives 6 5 8 Ruddervator For estimation of the ruddervator control derivatives 6 5 9 Ruddervator Tab For estimation of the ruddervator tab control derivatives 6 5 10 Flying Wing For estimation of the flying wing control derivatives 6 5 11 Elevon For estimation of the elevon control derivatives 6 5 12 Elevon Tab For estimation of the elevon tab control derivatives 6 5 1 Stabilizer Related Derivatives The methodology used to calcula
103. eight low to medium maneuverability airplanes e Class HI Large heavy low to medium maneuverability airplanes e Class IV High maneuverability airplanes only for MIL F 8785C After selecting the airplane category the flight phase category must be defined in the flight condition dialog box The user has the following options e Flight Phase A Plot the requirements for non terminal flight phases that require rapid maneuvering precision tracking or precise flight path control e Flight Phase B Plot the requirements for non terminal flight phases that are normally accomplished using gradual maneuvers and without precision tracking although accurate flight path control may be required e Flight Phase C Plot the requirements for terminal flight phases that are normally accomplished using gradual maneuvers and usually require accurate flight path control Once the aircraft certification category has been completely defined the user can choose from the following two options 7 4 2 1 Roll Performance 7 4 2 2 Spiral Dutch Roll Dynamics H 161 7 4 2 1 Roll Performance The Flying Qualities of the roll performance can be performed after selection of the Roll Performance option After the input is defined the output group will be filled 7 4 2 2 Spiral and Dutch Roll The Flying Qualities of the spiral and Dutch roll performance can be performed after selection of the Spiral Dutch Roll option After the inputs are defined the output gro
104. eight and center of gravity of the powerplant will be calculated using only the components for which all data is supplied If a component weight and center of gravity location has been entered in the From Fractions Empty Weight Table their values will automatically be entered in the Powerplant Weight Table e Fixed Equipment With this option the user can tabulate fixed equipment weight components and the associated center of gravity coordinates The overall weight and center of gravity of the fixed equipment will be calculated using only the components for which all data is supplied If a component weight and center of gravity location has been entered in the From Fractioms Empty Weight Table their values will automatically be entered in the Fixed Equipment Weight Table e Empty Weight The results from the Structures Powerplant and Fixed Equipment submodules described above are used to calculate the airplane current weight and center of gravity coordinates 1 4 2 2 C G Excursion This option allows the user to specify a component weight loading unloading scenario Loading starts from the empty weight of the airplane and successively adds weight components to the airplane until the maximum take off weight is reached At each step the new location of the airplane center of gravity is calculated and stored This procedure is repeated for the unloading case with the starting point at the maximum take off weight Mission loads and fuel are su
105. ent e Inertias The parameters that can be varied in this submodule are Lyg and Izp e Steady State The parameter that can be varied in this submodule is U Once the user has selected one of the sensitivity analysis options an input output window will appear asking the user to define the lower and upper limit of the variable to be investigated The sensitivity plot is displayed after these limits are defined and the Plot option is selected The plot options are described in Chapter 2 of Part II of the manual The input parameters required to perform a sensitivity analysis for the lateral directional derivatives can be entered in the Transfer Function submodule The methodology used for the sensitivity analysis can be found in Section 6 4 of Reference 11 7 5 Roll Rate Coupling Analysis The purpose of the roll rate coupling submodule is to analyze the roll pitch yaw coupling effect This allows the designer to determine whether or not the aircraft will be stable during steady roll and to adjust the level of longitudinal and or lateral directional stability if needed Once the Roll Coupling option is selected an input output window appears showing the roll pitch yaw coupling parameters This option is used to calculate the parameters associated with the critical roll rate After the Calculate option has been selected the user may see a roll rate coupling plot by selecting the Plot option The range of the variable to be investigated ma
106. er can select which features need to be recalculated or the user can choose to automatically run through a series of flight conditions marked in the Flight Condition window Trim diagrams can automatically be exported to WMF files and saved to the harddisk for each flight condition Marked variables can also be exported automatically to Excel spreadsheets This powerful tool allows the user to quickly and accurately create a picture of the aircraft aerodynamics and stability and control issues through a wide range of flight conditions at the click of a button PART I I 21 Recalculate ad Calculate O Stop i Elapsed Time Theoy a Close F Class Il Weight F Pitching Moment F Trimmed Lift T Const F Component C G F Stability amp Control Derivatives F Trimmed Lift T from D F Empty Weight C G F Class Il Drag F Transfer Functions F Fuel Weight C G l Trimmed Drag Trend Line F Flying Qualities F C G F Untrimmed Drag Trend Line F Static Margin F Forward Aft C G F Drag from Trend Line F Yen Diagram F Landing Gear Geometry F Critical Mach Number M Lift F Steady State Coefficients F Plot n Diagram F Maximum Lift F Hinge Moment Derivatives F Plot Trim Diagram File Location Export Plots F Export To Excel F Calculate All Marked Flight Conditions F Vary Mach Number for Given Altitude per FC Figure 2 12 Recalculate Dialog When the Print button on the main toolbar is selected the current print settings defined in
107. ere thrust is kept constant 6 10 1 Trim Diagram Analysis The Trim Diagram submodule is used to determine the trim characteristics of an airplane The methodology used to generate the trim diagram can be found in Chapter 4 of Reference 11 and Section 8 3 of Reference 7 Dynamics Il 151 The user has to select between Class I Class Il or Wind Tunnel Drag Polar modules see Section 2 4 from the Flight Condition dialog window It is not possible to generate a trim diagram without defining all input data The input window is different for each aircraft configuration and moving surface combination When all the input parameters are defined the trim diagram can be displayed by selecting the Plot button The plot may be altered using the conventional plot control buttons Section II 2 1 4 6 10 2 Longitudinal Trim The Longitudinal Trim submodule is used to solve the longitudinal equations for trim The methods in this module do not allow for canard or three surface airplanes The user must define a horizontal tail in the Airplane Configuration dialog box The input data required for the longitudinal trim submodule include the drag polar of the airplane The user has to select between Class or Class Il or Wind Tunnel Drag Polar modules see Section 2 4 from the Flight Condition dialog window The longitudinal trim calculations are executed by solving the following equations simultaneously e Drag equation e Lift equation e Pitching
108. erial properties and the results of the calculation of the Total Internal loads for the component see Chapter 8 Use of the Structures module will be described in the following sections 9 2 Structures Main Menu After selecting the Structures module the user can select from the two options displayed e User Materials In this submodule material properties that are not listed in the Available Materials table may be added and have their characteristics defined These materials will be added to the User Defined category of the Available Materials table e Class In this submodule the weight and size of a structural component can be estimated This submodule is explained in Section 9 2 1 9 2 1 Class I Structures This module is used to calculate the size and weight of a structural component to withstand the loads calculated in the Loads module based on the characteristics of the materials used Before these calculations can be done the component must be defined in the Airplane Configuration dialog box the Total Internal loads must be calculated and the structure of the lifting surfaces must be defined Structures T 181 The structure for the fuselage does not need to be defined and the structure for a lifting surface is defined in Lifting Surface Cross Section Structure dialog box which has the following options e Single Spar Semimonocoque e Two Spar Semimonocoque e Two Spar Drag Truss e Single Spar Circular e S
109. following sections 6 9 6 10 6 11 6 12 6 13 6 14 6 15 Class Class Il Stick Free Static Margin Stick Force Aileron Force Rudder Force Wing Location For estimation of the Class I empennage analysis For estimation of the airplane longitudinal and lateral directional trim characteristics For estimation of stick free Static Margin For estimation of stick force and stick force speed gradient For estimation of the aileron stick or wheel force For estimation of the rudder pedal force and rudder pedal force vs sideslip angle gradient For quick estimation of the wing location and its effect on stability 6 3 Longitudinal Stability Derivatives When the Long Stability has been selected a menu appears with five longitudinal stability derivative submodules These submodules are 6 3 1 6 3 2 6 3 3 6 3 4 6 3 5 I 130 Steady State Speed Angle of Attack A O A Rate Pitch Rate For estimation of the steady state lift drag moment and thrust coefficients For estimation of the speed derivatives For estimation of the angle of attack derivatives For estimation of the rate of angle of attack derivatives For estimation of the pitch rate derivatives Dynamics For certain submodules the user must define the lifting surfaces in the Airplane Configuration dialog box See Section 2 2 3 Figure 2 16 6 3 1 Steady State Coefficients This module can be used to calculate steady state lift d
110. g cost 10 8 4 Program Operating Cost After invoking the Program option the total program operating cost can be estimated Cost II 187 10 9 Life Cycle Cost The assumption is made that the R T D E program acquisition and program operating cost are estimated before entering this menu therefore the only input parameter needed to estimate the life cycle cost is the program disposal cost Program disposal cost is entered as a fraction of life cycle cost It is also possible in this window to redefine the other Life Cycle Cost components 10 10 Price Data After the selection of the Price Data option a menu is displayed with the following options e Airplane Before invoking the Airplane option a type of airplane must be defined in the certification dialog box The airplane price can be calculated after selecting the airplane category The airplane cost for a certain calendar year can be estimated in US e Engine Before selecting this option the user must specify the engine type in the engine dialog box from the following options o Jet Engine o Piston Engine o Propfan o Turboprop After the engine type selection is made the engine price can be calculated e Propeller Before selecting this option an engine other than a jet must be defined and the propeller type must be chosen in the engine dialog box from these options o Composite to calculate the composite propeller price o Metal to calculate the metal propeller p
111. ght estimate module the weight estimation methods are identified as follows e Cessna method e USAF method e General Dynamics GD method e Torenbeek method e Vought method The different weight component estimation options will be discussed in Section 1 5 1 through 1 5 4 Because airplane component weight modeling in the AAA program is a function of the airplane take off weight Wro the Class II weight estimation is an iterative process The weight iteration process will be discussed is Section 1 5 5 After selection of Class Il Weight the program will choose a weight estimation category based on the user supplied airplane certification type and powerplant specifications The Aircraft Weight Components menu will be displayed next with the following options 1 5 1 Structure To estimate the airplane structure component weights 1 5 2 Powerplant To estimate the airplane powerplant component weights 1 5 3 Fixed Equipment To estimate the airplane fixed equipment component weights Aerodynamics I 71 1 5 4 User Weight 1 5 5 Total Weight 1 5 6 Weight Iteration 1 5 7 Center of Gravity 1 5 8 Inertias 1 5 9 Set Category To estimate the user defined weights To calculate the take off weight To estimate the airplane take off weight by iteration of structure powerplant and fixed equipment component weights To estimate the airplane weight and center of gravity location for the specified loading scenario To estimate the
112. he Flying button a menu appears with the following options o Crew For estimating crew operating cost O Fuel amp Oil For estimating fuel and oil cost O Insurance For estimating insurance cost II 186 Cost e Maintenance e Depreciation e Fees e Total o Total For estimating the total flying cost Select this option for estimation of the direct operating cost of maintenance Before selecting the Maintenance option the user must specify the type of engine in the engine dialog box from the following options O Jet Engine O Piston Engine Oo Propfan O Turboprop Once the engine type is selected an input output window is displayed Select this option for estimation of the direct operating cost of depreciation Select this option for estimation of the direct operating costs of landing fees navigation fees and registry taxes Select this option for estimation of the total direct operating cost 10 8 3 Indirect Operating Cost In this submodule the indirect operating cost can be estimated After invoking the Indirect operating cost two options are displayed e Detailed e Alternate The indirect operating cost is a summation of the individual components Because the airplane designer has very little influence over this cost category the indirect operating cost is expressed as a fraction of the direct operating cost For this method the user need only define the fraction parameter and the direct operatin
113. he Mouse and Cursor The cursor usually appears as an arrow on the screen The arrow will change to an hourglass shape while the software is auto saving or during operations that take more time to perform By moving the mouse on the desktop the cursor can be moved to the desired position on the screen The left mouse button is used to select all buttons in the program When the cursor is placed over an input output element it will change to a calculator Also a brief description or hint will be displayed for that element The calculator can be displayed by clicking the left mouse button while the cursor is displayed as a calculator The calculator is described in Section 3 2 When the cursor is placed over an edit box in an input output element it will change to a bar If the left mouse button is clicked and insertion bar will appear in the edit box and the keyboard can be used to input a number 3 2 Operation of the Calculator The calculator pad is designed to minimize the use of the keyboard and to provide a user interface using a mouse The calculator is activated displayed on the screen by clicking the left mouse button when the cursor appears as a calculator The user may then use the calculator to input the value of the variable or the user may instead choose to use the number keypad on the keyboard to enter the data The calculator pad performs similar to an electronic calculator The buttons are selected with the left mouse button
114. he user may display the v tail geometry by selecting the Plot option 4 7 4 Ruddervator Tab In this submodule the geometry of the ruddervator is added 4 7 5 Chord Length In this submodule the chord length of the v tail at any span wise location can be determined by using root chord length taper ratio and the given span wise location 4 7 6 Airfoil In this submodule the v tail airfoil parameters at the v tail mean geometric chord are calculated 4 7 7 Exposed In this submodule the exposed v tail parameters are calculated I 120 Geometry 4 8 Ventral Fin Geometry In this submodule the ventral fin geometry is calculated 4 9 Fuselage Geometry The methodology used to calculate the fuselage geometry and fuselage related parameters is described in Chapter 4 of Reference 2 The user has to input the fuselage cross section data After calculating the output parameters the user may display the fuselage geometry by selecting the Plot option 4 10 Nacelle Geometry The methodology used to calculate the nacelle geometry and nacelle related parameters is based on the fuselage geometry methods The user has to input the nacelle cross section data and define the sections of the nacelle After calculating the output parameters the user may display the nacelle geometry by selecting the Plot option 4 11 Tailboom Geometry The methodology used to calculate the tailboom geometry and tailboom related parameters is based on th
115. hi en R E A E 68 14 22 CG EXCUISION iia 69 1 4 3 Inertia EStMale comodidad dois 70 1 44 Radii of Gyrationn ee n A E E E 71 S Class M Weight pesarese ene earra a e ON SINES REIRAS 71 1 5 1 Structure Component Weight Estimati0N ooncnnnnnnonononcnnnnonnnononanonncnncconocnnos 72 1 5 2 Powerplant Component Weight Estimation ooooonnnnnccnonnnonnnnncnanconocanocnnocnnos 73 1 5 3 Fixed Equipment Component Weight Estimation ooocnncnicnnnnnnnnnnnnconnconocnnos 73 1 5 4 User Weight Estimation oconncnonnnonnconnnoncnononnnonnncnnnnannnnnonnnconocnnonn nc nono necnnconnos 74 1 55 Total Weight Estimation 2 2 c scsssccscccsc essssseestenestesstetsseseeesseussdugsbebesveestetssesoens 74 1 5 6 Weight Iteration Process 20 0 cee cee cee cseeeseeeeeeeeceeeeeeeeeseceseeseenseenseesaeenaes 75 15 7 Center Of Gravito eenei ratoe senna anand EEEE ali 76 O O 77 1 5 9 Set Catesorysc costs chai a AR oie ees Sh ee 77 Aerodynamics Module ss ssscscc0ssisaeesoetceteassetavecvensscessdedenescebessasctetesascevdstasesusheacdescsphessetasesvess 79 2 1 General Description otitis aed ad wth E 79 2 2 Aerodynamics Main WiNdOW ccoccconnconoconocononnconnonononnnonnnrnncnnonnncnnncn nena nono neon ncnn corn ninns 79 A O RN 80 2 3 1 Wing Horizontal Tail Vertical Tail Canard V Tall oooonnnnnnnnninnnnn 81 2 3 1 1 Maximum Lift Coefficient oooconcnnoninonnnnnconnnonononcnancnnncnno canon nocnnccnnos 81 DIL IO DISTID
116. his command With this command a flight segment can be inserted between two previously entered segments When the Insert Segment command is selected a menu with twelve flight segments as described in New Segment will appear Upon selection of a new segment to be inserted the user must select the position of the new segment within the mission profile The new segment will be inserted before the segment selected as the new position Inserting at the end of the table is possible by selecting New Segment instead of Insert Segment To cancel insert segment select any button in another menu A segment in the mission profile table can be moved to another location by selecting 1t and selecting the flight segment after which II 63 it is to be inserted If the user wishes to exit without moving a segment select any menu to exit this command Input data in a particular flight segment can be changed by selecting the flight segment from the previously created table 1 3 2 Take off Weight This command is used to calculate the airplane take off weight from the mission segment fuel fractions and regression coefficients Once Take otf Weight is selected a display of all mission segments previously generated in Mission Profile will appear together with an input and output group The regression coefficients A and B can either be entered in the input section or calculated after invoking Regression After entering all input data and selecting Cal
117. icked a representation of the active window is sent to the default printer This option is faster than the Screendump option e Print Parameters This is performed by selecting the Print Parameters option in the Print dialog box See Subsection 2 2 1 Once print parameters is selected the user is prompted with the three options on the format of the output file See Subsection 2 2 1 Table 5 1 lists the resultant output for the three options Table 4 1 Resulting Output for the Three Formatting Options in Print Parameters Print Parameters Option Resulting Output Symbol Value Unit Advanced Aircraft Analysis Project Straight Tapered Wing Geometry ARy 10 00 500 00 ft 0 40 25 0 deg PART I I 55 Table 4 1 Resulting Output for the Three Formatting Options in Print Parameters Contd Description Value Unit Advanced Aircraft Analysis Project Straight Tapered Wing Geometry Wing Aspect Ratio 10 00 Wing Area 100 00 ft Wing Taper Ratio 0 40 Wing Quarter chord Line Sweep Angle 0 000 deg Description Symbol Value Unit Advanced Aircraft Analysis Project Straight Tapered Wing Geometry Wing Aspect Ratio 10 00 Wing Area Sw 100 00 ft Wing Taper Ratio Aw 0 40 Wing Quarter chord Line Sweep Angle Non 0 000 deg I 56 PART I 5 Software Databases and Projects The system features a database structure to store and retrieve airplane design and analysis project parameters Two types of da
118. il and Certification Classification Define the class of the airplane for US military flying qualities a regulations to evaluate flying qualities for both civilian and military airplanes Classificatio Certification and Type Dialog Box When the Certification button is selected the Certification and Type dialog box is displayed Figure 2 23 shows the Certification and Type dialog box for civil airplanes and military airplanes Descriptions of the dialog box functions follow PART I I 37 Certification and Type Type Military Civil Type C Civil UAY C Sailplane C Civil C Civil UAY C Transport e O Mil UAY C Ultralight Mil UAY Certification C Homebuilt Certification C FAR 23 C FAR 25 C FAR 23 C FAR 25 c C Light Airplane Bomber C JAR23 Mil Specs C JAR23 C Mil Specs C VLA CAS Specs Aaricultural C VLA C AS Specs C Light Sport C AmphibiousfFlying Boat C Light Sport Fighter C Business ieee C Trainer C Transport C Supersonic Cruise C Patrol C NACAJNASA fy X Cancel Help E X Cancel Help Figure 2 23 Certification and Type Dialog Box The Certification and Type dialog box offers the user the following options for defining the aircraft categories Military Civil The user can select from military or civil airplane types Type Depending on the Military Civil choice the user can choose the airplane type Certification The user can select the following certification standards
119. ilboom drag coefficient 2 4 2 10 Float For estimation of the float drag coefficient 2 4 2 11 Trail Edge Flap For estimation of the trailing edge flap drag coefficient 2 4 2 12 Lead Edge Flap For estimation of the leading edge flap drag coefficient 2 4 2 13 Landing Gear For estimation of the gear drag coefficient 2 4 2 14 Canopy For estimation of the canopy drag coefficient 2 4 2 15 Windshield For estimation of the windshield drag coefficient 2 4 2 16 Store For estimation of the stores drag coefficient 2 4 2 17 Trim For estimation of the trim drag coefficient 2 4 2 18 Spoiler For estimation of the spoiler drag coefficient 2 4 2 19 Speed Brake For estimation of the speed brake drag coefficient II 86 Aerodynamics 2 4 2 20 Miscellaneous For estimation of the miscellaneous drag coefficient 2 4 2 21 Pylon For estimation of the pylon drag coefficient 2 4 2 22 Windmilling For estimation of the drag coefficient due to windmilling 2 4 2 23 Inlet For estimation of the inlet drag coefficient 2 4 2 24 Nozzle For estimation of the exhaust nozzle drag coefficient 2 4 2 25 Total Drag For a summation of the different component drag coefficients 2 4 2 26 Recalculate All Used to recalculate all of the drag coefficients by pressing calculate once 2 4 2 27 Trendline For calculating a Trendline of the drag polar 2 4 2 28 Plot For plotting the trimmed and untrimmed drag polars 2 4 2 1 Wing Drag In this module the wing drag coefficient c
120. in reference to the wing Dynamics O A C C G For sizing the horizontal tail canard or v tail based on aerodynamic center location 6 9 3 Minimum Controllable Speed with One Engine Inoperative Stability The One Engine Out submodule is used to check the size of the vertical tail for the case of climb with one engine inoperative and applies only to airplanes with more than one engine The rudder deflection required to maintain controllability in the one engine out condition at minimum controllable speed is also calculated The methodology used to calculate one engine inoperative stability can be found in Section 11 3 of Reference 2 6 10 Class II Trim Analysis The methodology used to analyze the longitudinal and lateral directional trim characteristics can be found in Chapter 4 of Reference 11 When the Class Il button is selected the following options appear on the screen 6 10 1 Trim Diagram For determination of longitudinal trim characteristics 6 10 2 Long Trim For determination of elevator stick or control wheel forces 6 10 3 Lat Dir Trim For determination of aileron stick or control wheel forces and rudder pedal forces 6 10 4 T O Rotation For determination of the horizontal tail area required for rotation 6 10 5 Trimmed Lift T from D For determination of the trimmed lift coefficients where thrust is calculated from the drag 6 10 6 Trimmed Lift T Const For determination of the trimmed lift coefficients wh
121. in the software program group in Windows 2 PART I 2 Structure of the Software The software uses windows toolbars and dialog boxes to communicate with the user This chapter describes the structural elements of the software their purpose and their functionality The following elements of the software are described in this chapter 2 1 Windows and command bars 2 2 Toolbars 2 3 Menu bar 2 1 Windows and Command Bars The software is started by selecting the program icon in the Airplane Design and Analysis program group in Windows When the program is started the main window Figure 2 1 is displayed This window is open as long as the program is running The main window contains a Windows menu bar at the top the main menu of application modules the software toolbars and the status bar The status bar is located at the bottom of the main window and contains the company name and project name as specified by the user and the current date and time When an element of the status bar is double clicked with the mouse button a dialog box appears to change the content or format of that element see Section 2 2 PARTI I 3 Main Menus EN Advanced Aircraft Analysis 3 2 Project1 7 aa Flight Condition 1 Ele Edit Window Help AT Stap Main Window Application Window ATA Weight SS Aerodynami E Performance a Geomety a Propulsion j amp Control AK Dynamics aJi Loads SZ Structu
122. in the following sections 8 2 Loads Main Window After invoking the Loads module the user may enter the following modules e V n Diagram In this submodule velocity vs load factor V n diagrams can be constructed for the following types of airplanes FAR 23 certified FAR 25 certified and MIL A 8861 ASG certified airplanes Reference 17 The options presented in this module are discussed in Section 8 3 e Structural In this submodule the total internal loads for each structural component can be calculated in any combination that the user desires The options presented in this module are discussed in Section 8 4 83 V n Diagram Before invoking V n Diagram the user must choose one of three certification options from the Certification and Type dialog box Loads H 171 The options are as follows e FAR 23 V n diagrams can be constructed for FAR 23 certified airplanes After selecting FAR 23 in the certification dialog box a selection between the following FAR 23 airplane categories must be made o NORMAL o UTILITY o ACROBATIC e FAR 25 V n diagrams can be constructed for FAR 25 certified airplanes e MIL A 8861 ASG V n diagrams can be constructed for military certified airplanes according to the MIL 8861 ASG regulations Reference 17 The V n diagram output parameters are generated after all input data are provided and the Calculate option is selected Once the V n diagram input and output parameters are defined in the
123. inary design engineers to rapidly evolve a preliminary aircraft configuration from early weight sizing through open loop and closed loop dynamic stability and sensitivity analysis while working within regulatory and cost constraints The software is written for the Microsoft Windows graphical interfaces This allows the user to take full advantage of the interface It is recommended that it be run as a full screen program with a display resolution of at least 1024 x 768 or higher It will run with any display resolution but scrolling may become necessary for display resolutions less than 1024 x 768 The software consists of ten application modules a detailed help system environment setup and project handling tools Chapter 2 describes the structure of the program Chapters 3 and 4 describe and explain the various tools used in the program Chapter 5 discusses the project handling methods and capabilities of the software Chapter 6 describes the online help system in the software All information contained in this manual is also contained in the software help system The help PART I I 1 system can be accessed at any time from within the program by selecting the Help button at the bottom of the main window by selecting an option in the Help menu at the top of the main window by selecting the Theory button in the input output window or by pressing F1 on the keyboard The help system can also be accessed by selecting the User s Manual icon
124. ingle Spar Box When a structural component is selected the following options will become available e Comp Materials e Section Properties e Weight Estimate I 182 To select the materials used for different parts of the component The Component Material Properties must be defined before Weight Estimate can be performed When the Material Name column is selected for a section of a component the Available Materials table dialog is displayed To calculate the effective width and height of the structure at each load station by entering spar and surface locations The effective width and height are calculated differently for the different types of structures For the fuselage the structure type must be defined at each load station To size the component structure to the loads calculated in the Loads module Chapter 8 The Class method calculates an effective width and effective height of the structure at each load station and sizes the structural materials to withstand the loads Structures 10 Cost Analysis Module 10 1 General Description The purpose of this module is to estimate various costs of airplane design programs The estimation methods are presented in such a manner that they can be applied to civil and military airplanes of all types The various cost definitions and cost estimation methods used for this module are as discussed in Chapter 1 through 7 of Reference 9 Use of the Cost module options will be de
125. ining these quantities by following these steps for the equivalent parasite area II 84 Aerodynamics e Input the desired value for equivalent parasite area e Make a or b undefined e Select Calculate The same procedure applies to the airplane wetted area The wetted area is a function of the regression coefficients c and d Reference 1 Section 3 4 1 and are user defined in the input parameter section To provide a user selectable value for the wetted area follow these steps for the wetted area e Input the desired value for wetted area e Make c or d undefined e Select Calculate After calculating the drag polar a dialog window displays the drag polar equation This equation 1s printed with the Print Parameters option After calculating the drag polar equation the user may display the drag polar by selecting the Plot option After selecting Plot the following graphs will be displayed for the selected condition o CL VS Cp e Cr VS Cr Cp e Cy vs C IC For propeller powered configurations e Cr vs CP Cp For jet powered configurations A description of the standard plot options can be found in Chapter 2 of Part Il of this user s manual NOTE The C range must be specified in input parameters Changing the plot axis will not allow for calculation outside of the specified Cz range 2 4 2 Estimation of the Class II Drag Polar The purpose of the Class II Drag submodule is to supply a C
126. ion When the Matching Plot option is selected the system provides a list for the user to specify which performance sizing options are to be plotted The options can be plotted individually or in a group by selecting the desired combination of options A check mark indicates if an option has been selected Once the selection has been made the user selects the Continue button and an input output window appears with matching graph input parameters The user now specifies the desired range of take off wing loading values by defining the minimum and maximum W S ro limits When the user selects Plot the pre selected performance sizing options will be graphed With the help of such a plot the user may determine the highest possible wing loading and the smallest possible thrust or highest power loading which meet the performance requirements The standard plot options are described in Chapter 2 of Part Il of this manual 3 4 Performance Analysis After choosing the Analysis submodule another menu appears with ten options The ten options represent submodules as follows 3 4 1 Thrust Speed or 3 4 6 Cruise 3 4 2 Power Speed 3 4 7 Dive amp Descent 3 4 3 Stall Speed 3 4 8 Maneuver 3 4 4 Take off Distance 3 4 9 Glide 3 4 5 Climb 3 4 10 Landing Distance The use of these Class I mission performance analysis submodules will be described in the following sections Many of these submodules require the use of Class I drag polar data There are also n
127. ise Speed For sizing to maximum cruise speed requirements 3 3 5 Maneuver Turn For sizing to maneuvering requirements 3 3 6 Landing Distance For sizing to landing distance requirements 3 3 7 Matching Plot For plotting the performance sizing equations The user may choose to size an airplane by using any combination or all of the above sizing options Several of the performance sizing options require input only from the user while others also provide the user with output Once all of the inputs for the desired options have been entered and where applicable the output has been calculated the user may select the Matching Plot option to plot the performance sizing equations The procedures to use each performance sizing and matching plot submodule are presented in the following sections 3 3 1 Stall Speed Sizing To size the aircraft to meet stall speed requirements the user selects the Stall Speed option The methodology used in sizing to stall speed requirements can be found in Reference 1 Section 3 1 The input and output data for the Stall Speed submodule are for all airplane types and specifications H 102 Performance 3 3 2 Take off Distance Sizing To size the aircraft to take off distance requirements the user selects the Take off Distance option The methodology used in sizing to take off distance requirements can be found in Reference 1 Section 3 2 3 3 3 Climb Sizing To size the aircraft to meet climb requirements the
128. lass II method for predicting drag polars of airplanes during the preliminary design phase After invoking Class II Drag the program requires the definition of the altitude and speed in the Flight Condition Dialog to determine the Aerodynamics Il 85 flow regime for that particular flight condition The different flow regimes are divided according to Mach number ranges into e Subsonic Flow regime for flight Mach numbers not exceeding 0 6 e Transonic Flow regime for flight Mach numbers between 0 6 and not exceeding 1 2 e Supersonic Flow regime for flight Mach numbers between 1 2 and not exceeding 3 0 NOTE Drag estimation methods for the hypersonic flow regime flight Mach numbers greater than 3 0 are not available in this module Once the flow regime has been selected the user will enter the Class II Drag Polar Prediction menu which has the following options 2 4 2 1 Wing For estimation of the wing drag coefficient 2 4 2 2 Horizontal Tail For estimation of the horizontal tail drag coefficient 2 4 2 3 Vertical Tail For estimation of the vertical tail drag coefficient 2 4 2 4 Canard For estimation of the canard drag coefficient 2 4 2 5 V Tail For estimation of the v tail drag coefficient 2 4 2 6 Ventral Fin For estimation of ventral fin drag coefficient 2 4 2 7 Fuselage For estimation of the fuselage drag coefficient 2 4 2 8 Nacelle For estimation of the nacelle drag coefficient 2 4 2 9 Tailboom For estimation of the ta
129. lat Gear Structure Figure 2 15 Configuration Setup Toolbar I 24 PARTI Table 2 6 Configuration Setup Toolbar Buttons Configuration Define the basic configuration of the airplane which includes empennage nacelle s store s pylon s tailboom s speed brake and spoiler In the case of stores speed brake and spoiler the flight condition dialog box is used to define whether the devices are deployed in the current flight phase Engine Define various aspects of the propulsion system of the airplane Controls Define longitudinal and directional control surfaces for the airplane Controls Flap Slat Define the number and type of high lift devices per wing Gear Define the type of landing gear position retraction and attachment point In the case of retractable gears the Flight Condition dialog is used to define whether the gear is extended or retracted in the flight phase Structure Define the cross section structure type of wing and empennage amphibious hull pressurization and attachment types for the wing and the Structure empennage Each of the buttons in the Configuration Setup toolbar opens a dialog box when selected The remainder of this subsection will present the dialog boxes opened by the Configuration Setup toolbar buttons Airplane Configuration Dialog Box When the Configuration button is selected the Airplane Configuration dialog box is displayed Figure 2 16 shows the Airplane Configuration dialog box
130. late button is selected that method will be included in the iteration The weight iteration component weight table contains a Wuser column This column can be used to fix the weight of a specific component The column can also be used to influence the average weight of a component To fix the weight of a component enter the known weight in the Wuser column and delete the contents of the other estimation method cells for the component To influence the average leave one or more weight estimation cells defined and enter a number in the Wuser column 1 5 7 Center of Gravity This option allows the user to calculate the airplane weight and center of gravity location for a specific loading scenario by entering weight components and their corresponding center of gravity locations in a preformatted table After selecting Center of Gravity a menu with three options appears e Copy Class C G With this option the values for the Class I center of gravity locations from the Class I Detailed center of gravity tables are copied to the Class II center of gravity table Using this option the center of gravity values in the Class II table will be overwritten II 76 Aerodynamics e Component C G This module calculates the center of gravity location in terms of the wing mean geometric chord based on basic geometry and center of gravity input parameters The module also calculates the C G locations of the fuselage fuel wing canard horizontal t
131. les are located Propulsion The user can define a jet propeller or a ducted fan Engine Type The user can specify the engine to be a piston propfan or a turboprop for a propeller or a ducted fan Aspiration For a piston driven propeller or ducted fan the user can specify whether the engine is normally aspirated or supercharged or turbocharged Fuel Tank For all engine types the type of fuel tank can be specified Type of Inlet For all engine types the type of engine inlet can be specified Number of Propellers For all engine types except jet the user can define the number of propellers The number of propellers can be different from the number of engines PART I e If Jet is selected for the powerplant type Figure 2 17a the type of starting system can be specified e If Propeller or Ducted Fan is selected for the Propulsion type and Piston is selected Engine Type the following options are available o Propeller Pitch The user can choose a variable or fixed pitch propeller o Propeller Type The user can choose a metal or composite propeller o Aspiration The user can define aspiration e If Propeller or Ducted Fan is selected for the Propulsion type and Propfan or Turboprop is selected as the Engine Type the following options are available o Propeller Pitch The user can choose a variable or fixed pitch propeller o Propeller Type The user can choose a metal or composite propeller o Propeller Reversing For Ducted Fan
132. light Condition List 50 Entries Variable Changed Default Color HBOOoodu lt Export Flight Condition In Rows C In Columns Default Note Color Hint on Oa ca ga ga X Cancel Help Figure 2 32 Program Options Dialog Box Toolbar Button Options The user can choose from the three types of toolbar buttons which will be displayed when the current project is opened The button options are shown in Figure 2 32 Parameter Notes The user can choose whether to switch on the parameter Notes button see Figure 2 32 Parameter Info The user can choose whether to switch on the parameter Info button see Figure 2 32 Tips at Startup The user can choose whether to have the tips show up at the start of the program See Figure 2 32 Export All Values The user can choose whether to export all values to ASCH Save WMF to File The user can choose whether to save the WMF graphics created by using the Copy WMF button in the main toolbar Figure 2 1 to a file Show Variable Name Displays the database variable name in the work pad dialog window PARTI e Auto Save Recovery Project The default interval for auto saving a recovery project is set to 5 minutes Users can alter the interval according to individual preference Detail on the recovery project is presented in Section 5 3 e The Recently Used File List lets the user specify the number of recently used files that appear on the File menu
133. litude ratio or decibels see Section 11 4 of Reference 12 After selecting the Bode Method option the single feedback loop control system window appears This new window consists of eight separate transfer function T F boxes in addition to the Plot Open Loop button After all the required transfer functions have been defined the Plot Open Loop button is selected to calculate the open loop dynamic characteristics of the control system After the user defines a frequency range the Plot button is used to generate the Bode plot A detailed explanation of the plot options can be found in II 168 Dynamics Chapter 2 in Part II of the user s manual 7 9 Human Pilot The Human Pilot calculation can be used to estimate a human pilot transfer function for use in the open loop control system analysis The Human Pilot module can be used to model pilots of differing abilities reaction times and physical fitness This module can even be used to show the dangers of a drunken pilot in the loop The methodology used to analyze a human pilot transfer function can be found in Chapter 10 of Reference 12 7 10 Time Delay The Time Delay calculation can be used to account for time delays and bus transmission delays due to flight control computer systems 7 11 Erase T F The user should be aware that the program database is structured so that the transfer functions which are loaded into the T F boxes will be passed from one control analysis optio
134. lso used to calculate dimensional loads from the coefficients The Forces module and Moments module can be used to directly define the dimensional loads created by a control surface deflection e Concentrated The Concentrated module can be used to define any loads acting on the fuselage or lifting surfaces as pure concentrated forces or moments These forces and moments may be defined in terms of changes in the coefficients of the surface in the Concentrated Coefficients Aerodynamic only Forces may be defined in the Concentrated Forces module Moments may be defined in the Concentrated Moments module e Distributed The Distributed module can be used to define any loads acting on the fuselage or lifting surfaces as pure distributed forces or moments These forces and moments may be defined in terms of changes in the coefficients of the surface in the Distributed Coefficients Aerodynamic only Forces may be defined in the Distributed Forces module Moments may be defined in the Distributed Moments module 8 4 1 4 Load Stations Before the Total Internal loads module may be accessed the load stations of the structural component must be defined The internal loads will be calculated at each of the load stations in the Total Internal module The load stations module simply divides the span of the surface or length of the fuselage into equal segments The location of each load station is displayed in the Loads Y 177 load station table
135. ly 1f the horn is fully or partially shielded 1 e Fully Shielded Horn or Part Shielded Horn the user is given the following options to define the horn nose e Blunt Nose For a horn with a blunt nose e Elliptic Nose For a horn with an elliptic nose 6 8 Recalculate All This module allows the user to recalculate all of the stability and control derivatives by pressing calculate once The user can change a parameter in the input and view its effect on the stability and control derivatives A stability or control derivative will only be recalculated if all of its associated input is defined 6 9 Class I Stability and Control Empennage Sizing Analysis The methodology used to analyze the Stability and Control characteristics can be found in Chapter 11 of Reference 2 When the Class submodule has been selected a new menu appears on the screen with three Class I Stability and Control Empennage Sizing options 6 8 1 Longitudinal For longitudinal stability and empennage sizing 6 8 2 Directional For directional stability and empennage sizing II 148 Dynamics 6 8 3 One Engine Out For minimum control speed with one engine inoperative stability empennage sizing Other important stability and control characteristics such as trimmability take off rotation crosswind controllability and a variety of dynamic stability considerations are not covered by the Class I method 6 9 1 Static Longitudinal Stability The Longitudinal submodule is use
136. metry This option is used to define the pylon geometry 4 17 Propeller Geometry This option is used to define the propeller geometry 4 18 Airplane 3 View This option is used to display the front side and top view of the airplane 4 19 Angles This option all the pertinent angles used in the analysis 4 20 Scale This option is used to uniformly scale an airplane using a user input scale factor This feature is useful when analyzing a scale model such as a wind tunnel model I 122 Geometry 4 21 Translate This option is used to translate the airplane in the X and Z directions 4 22 Airfoil Folder This option is used specify the location of the airfoil files 4 23 AeroPack This module can be used to export 3 Dimensional geometric data to the AeroPack drafting program Reference 18 After selecting the AeroPack option the following options are displayed e Wing e Horizontal Tail e Vertical Tail e Canard e V Tail e Ventral Fin e Fuselage e Tailboom e Nacelle e Trail Edge Flap e Lead Edge Flap e Landing Gear e Canopy e Aileron e Elevator e Rudder e Canardvator e Ruddervator e Spoiler e Store e Pylon Geometry TT 123 e Speed Brake e Dorsal Fin e Propeller e Float e Airfoil Folder e Export Data The components that have been completely defined in the Geometry module can be selected for exporting 4 24 Exporting to Shark SharkFX AP After selecting the Export Data option a dial
137. metry The options provided in this module are discussed in Section 4 17 This option can be used to output airplane three view The options provided in this module are discussed in Section 4 18 This option lists all the pertinent angles The options provided in this module are discussed in Section 4 19 This option can be used to scale an entire airplane uniformly Additionally the airplane can be translated in the X and Z direction The options provided in this module are discussed in Section 4 20 This option can be used to translate an entire airplane in the X and Z direction The options provided in this module are discussed in Section 4 21 This option can be used to specify the airfoil file location The options provided in this module are discussed in Section 4 22 This option can be used to output airplane geometry that can be shared with Shark AP or SharkFX AP The options provided in this module are discussed in Section 4 23 Geometry 4 3 Wing Geometry The methodology used to calculate the wing geometry related parameters is described in Chapter 6 of Reference 2 The Wing Geometry submodule has eleven options 4 3 1 Straight Tapered For calculation of straight tapered wing geometry 4 3 2 Cranked Wing For calculation of the equivalent straight tapered wing geometry from a given cranked wing geometry based on the tip chord or Mean Geometric Chord 4 3 3 Fuel Volume For calculation of the volume of fuel the wing can hold ba
138. module are only valid in the subsonic speed regime In the transonic and supersonic speed regimes the wave drag generated by the windshield can be significant In these speed ranges it will be necessary to employ area ruling to reduce the wave drag to a minimum The methodology used to calculate the windshield drag coefficient is described in Section 4 8 of Reference 7 2 4 2 16 Stores Drag In this module the store s drag coefficient can be calculated The methodology used to calculate the stores drag coefficient is described in Section 4 9 of Reference 7 Before entering the Stores Drag input output window the user will be required to enter the number of stores and to define which of the stores are on the airplane for the particular flight condition The Stores Drag input output windows depend on the selected flow regime 2 4 2 17 Trim Drag In this module the trim drag coefficient can be calculated for a specific flight condition The methodology used to calculate the trim drag coefficient is described in Section 4 10 of Reference 7 Il 90 Aerodynamics 2 4 2 18 Spoiler Drag In this module the spoiler drag coefficient can be calculated The methodology used to calculate the spoiler drag coefficient is described in Section 4 12 of Reference 7 Before entering the spoiler drag input output window the user will be required to define whether the spoiler is retracted or deployed for the particular flight condition 2 4 2 19 Speed
139. n 1 HH cia Preso A eat 2 0000 Lift Coefficient Dy e Ent i Qu AER ES D S 170 00 ft 1 5000 4 1 0000 4 0 5000 4 o o000 Nes pa A 0 0000 0 0500 0 1000 0 1500 0 2000 0 2500 0 3000 Drag Coefficient Cp L 1 I 1 1 1 J 0 0 25 50 75 10 0 125 15 0 UCA L 1 i 1 1 1 J 00 25 0 50 0 75 0 100 0 1258 160 0 Ccp TIA E E E ie Ea j E a New Open Save SaveAs Delete Flight Cond Recalculate Notes Copy WMF Print Atmosphere Help Exit File Configuration Certification Setup DARcorporation Advanced Aircraft Analysis 3 2 Project 05 01 09 12 13 PM Figure 2 7 Plot Window I 12 PART I Most plot windows contain a legend at the top right corner of the window Plot windows also contain one or more vertical and one or more horizontal axes When the cursor is moved over an axis it appears as horizontal and vertical axes When the cursor changes the left mouse button can be double clicked and a dialog will be displayed allowing the user to change the axes see Figure 2 8 The minimum and maximum values the major and minor divisions and the number of displayed decimal places can be changed If the axes are expanded beyond the original range of the calculation the plotted parameters will not be recalculated for the expanded range To recalculate the parameters the user should close the plot window and increase the range of calculation in the input output window The first time the program cre
140. n a w in elelee e e e ay gelee Saa elelee ee eee _ o a Engine Air Flow Location Starting System C Not Buried no airflow from fuselage base C Pneumatic C Buried in Fuselage airflow from fuselage base C Electric C X Cancel Help Figures 2 17a Powerplant Selection Dialog Box for a Jet Propulsion Engines Nacelles Propulsion Engine Type Yes No Number 0 C Yes No Jet p 3 Propeller C Propfan Engine Location Nacelle Location C Ducted Fan Turboprop Wing Fuselage Other Wing Fuselage Other c E aj Y al c Fuel Tank Aspiration C Integral C Normally Aspirated C Self Sealing Bladder C Supercharged Non Self Sealing C Turbocharged Type of Inlet Prop Type C Plenum C Composite C Straight Through C Metal Prop Pitch C Fixed C Variable oon 0 mia wl ny eleele te el tetee elelee e e eee e eee e eli e e e wow o ana uni 211 1162111160110 11801189 elelee eee e e elelee lelene e o Engine Air Flow Location Number of Propellers 0 C Not Buried no airflow from fuselage base C Buried in Fuselage airflow from fuselage base Cd X Cancel 7 Help Figures 2 17b Powerplant Selection Dialog Box for a Propeller Piston PART I I 27 Propulsion Engines gt pNacelles Propulsion Engine Type Yes C No Number 0 C Yes C No C Jet C Piston Propeller Engine Location Nacelle Location j 2 Ducted Fan C Turboprop Wing Fuselage Other Wing Fuselage Other C C C C C
141. n gain can be specified In a double loop control system the user is also asked to define the inner loop gain If the user only wishes to see the root locus plot for a range of system gains the design gain number of damping ratio lines and number of natural frequency arcs may be left undefined If the design gain is defined it is displayed in the root locus plot After selection of the Plot option the root locus plot is displayed A detailed explanation of the plot options can be found in Chapter 2 of Part II of the user s manual 7 7 2 S Plane Gyro Tilt Analysis An important consideration in the analysis of single loop control systems is the study of gyro tilt angle effect This study is essential in the design of yaw dampers After selecting the Gyro Tilt submodule the gyro tilt angle analysis window appears This new window consists of six separate transfer function T F boxes in addition to the Plot Open Loop button The required transfer functions must be defined by methods described in Section 7 6 It should be noted that there are additional constraints when defining transfer functions in the Dynamics II 167 Gyro Tilt submodule which are described below 1 For gyro tilt analysis the inner loop is not a feedback loop To analyze the gyro tilt effect the yaw rate and roll rate transfer functions must be defined in the inner loop 2 In addition the yaw rate to rudder transfer function must be defined in T F box 5 and the roll
142. n terms of rate for example roll rate to aileron transfer function the angle transfer function can be multiplied by S which the user can input in another T F box on the same path to convert to a rate transfer function Similarly a rate transfer function can be converted to an angle transfer function by multiplying the rate transfer function by 1 S Dynamics 7 7 S Plane Root Locus Analysis Using the Root Locus s submodule the user can generate a root locus plot in the S plane for continuous systems This submodule consists of the following five options 7 7 1 Single Loop For single loop feedback control systems 7 7 1 Double Loop K fwp For double loop control systems with the inner loop gain in the forward path 7 7 1 Double Loop K fbp For double loop control systems with the inner loop gain in the feedback path 7 7 2 Gyro Tilt To analyze the gyro tilt angle effect 7 7 1 S Plane Single and Double Loop Analysis After selecting a single or double control loop analysis and defining all required transfer functions see Section 7 6 the Plot Open Loop button is selected to calculate the root locus of the control loop An input output window will be displayed in which the lower and upper limit for the system gain the lower and upper limit for damping ratio lines the lower and upper limit for natural frequency arcs the number of gain points the number of damping ratio lines the number of natural frequency arcs and the desig
143. n this submodule the installed performance of the propulsion system can be estimated for a single set of conditions In case there is no engine and or propeller data file available the user Dynamics H 127 can estimate the installed characteristics for a single condition This requires uninstalled engine propeller performance data It is suggested that the manufacturer s data be used 5 8 Propeller In this submodule propeller performance maps can be imported and used to determine the efficiency of the propeller under various flight conditions The efficiency can also be plotted as a function of speed and altitude H 128 Dynamics 6 Stability amp Control Module 6 1 General Description Stability and Control consists of two submodules Derivatives and Analysis The purpose of the Derivatives submodule is to compute the non dimensional aerodynamic stability and control derivatives and hingemoment derivatives for a rigid airplane in a given flight condition i e for a given weight altitude and speed The purpose of the Analysis submodule is to help the user rapidly determine whether or not a proposed aircraft configuration will have satisfactory stability and control characteristics The Stability amp Control module also allows the user to calculate the required vertical tail canard and or horizontal tail surface area needed to satisfy user defined stability and control characteristics The mathematical equations used to comp
144. n to another That is if the user has defined transfer functions in T F boxes 1 and 3 in the single loop root locus option these same transfer functions will remain in T F boxes 1 and 3 should the user quit the single loop option and execute the double loop gyro tilt analysis or Bode method option This saves the user from having to define these same transfer functions again should the need arise to use another root locus option If the user does not wish to pass the transfer functions defined in one control system to another the Erase T F button may be used This option initializes all the user defined transfer functions to a unity transfer function In addition to using the Erase T F button the transfer function can also be deleted by right clicking on each T F box and selecting Erase T F Dynamics Il 169 H 170 Loads 8 Loads Module 8 1 General Description The purpose of this module is to estimate loads placed on airplane components and to determine important information for structural design and sizing The Loads module consists of two submodules V n Diagram and Structural Loads The purpose of the V n Diagram submodule is to determine load factors and their corresponding speeds These methods are based on Section 4 2 of Reference 5 The purpose of the Structural Loads submodule is to calculate the internal forces and moments in the structural components of an airplane Use of the Loads module options will be described
145. nerated I 42 PARTI Print Print Options Print Parameters Options C Screendump Symbol Value Unit Active Window C Description Value Unit C Print Parameters C Description Symbol Value Unit Print Page Numbers MV Print Date l Print File Name MV Print Time x Cancel 7 Help amp Setup Figure 2 30 Print Setup Dialog Box When the user selects Calculator on the System Setup toolbar the Calculator Setup dialog box is displayed The calculator is described in Section 3 2 The Calculator Setup dialog box is displayed in Figure 2 31 The dialog box allows the user to specify a standard or RPN calculator type Calculator Setup Calculator Type Figure 2 31 Calculator Setup Dialog Box When the user selects Options on the System Setup toolbar the Program Options dialog box is displayed The Program Options dialog box is displayed in Figure 2 32 Descriptions of the dialog box functions follow PART I 1 43 Program Options Toolbar Button Options C Medium Buttons without Captions C Small Buttons without Captions W Parameter Notes M Parameter Info l Export All Values I Save WMF to File F Show Variable Name Default Angle Units deg rad Default 1 Angle Units d eg rad Default Length Units ft M Tips at Startup Default Area Units 2 s2 ft in Auto Save Recover Project Every 5 a Minutes Recently Used File List 5 Entries Default Speed Units F
146. nnonononnnonn nono n nan non conc cnn cc nono necnnos 159 TAT Transfer Functions esea eE E Seca vbecbesevins souneaeas Ea 160 Vill Table of Contents 7 4 2 Lateral Directional Flying Qualiti8S ooncnnnnnonocnnonnnonccnncranonnnonnncancnanonnnono 160 TAQ Roll Performance iii l 162 7 4 2 2 Spiral and Dutch Roll oonooncnoncniconocanacanoconnnconoconoconocnnorononnnonnncnos 162 7 4 3 Lateral Directional Stability Derivative Sensitivity Analysis 162 7 5 Roll Rate Coupling Analysis 0 eee eeseeseceseceseceseceeecaeecaeeeaeseeeeseeeeeeeeeseenseenaes 163 7 6 Defining Control Transfer Functions ee cee csecseeeseeeeeeeeeeeeeeeeeeseceseeeseeaeenaes 164 7 7 S Plane Root Locus Analysis oooocnnncnncnnonnconoconoconccnnocnnonnnonnncnn cnn onnnonn no nro naco nera nconnos 167 7 7 1 S Plane Single and Double Loop AnNalysiS o oonncnnocnnoniconononcnnncnnncnncnancnnnono 167 7 71 2 S Plane Gyro Tilt AnalySIS ooonononnnnnnnconocnnonnconoconoconoconocn nono nonnncnn crac crnncnnnono 167 7 8 gt Bode Method it 168 29 Human Plot ir a ias 169 FAO Time Dela viciosa Gea eis oases oi ia as 169 TAT Erase MP orita pa tias ss 169 So Loads Mod le ted idad doit 171 8 1 General Description eee ee ea aE S ea eRe IERE ES eiS 171 8 2 L ads Main Window O 171 8 amp 3 Wen Diaria nia 171 8 4 Structural Lords ici EE SEEE E AA E eA 172 8 4 1 Fuselage Wing Horizontal Tail Canard Vertical Tail V Tall oooonnnninnn 173
147. ns Figure 2 2 shows an input output window with a table for fuselage section input The table can be resized rows added or subtracted using the spin edit element which appears as the last element in the input menu see Figure 2 2 The spin edit element is similar in appearance to an input output element The number of rows of a table can be changed by clicking on the arrows in the spin edit element The work pad window can be used to maintain notes of a particular project This window can also be used to change the units of a specific parameter and maintain notes about the parameter see Figure 2 4 The input menu of the input output window may also contain a combo box element see Figure 2 5 The combo box element is similar in appearance to an input output element but does not contain an edit box The combo box element contains a list of choices that affect the calculation results The list of choices is displayed by clicking on the arrow at the right side of the element and holding down the left mouse button A choice can then be made by moving the cursor to the appropriate choice and releasing the mouse button Figure 2 5 Combo Box Element Most input output windows contain an output group of elements showing the results of the calculation performed in the window The output group can contain output elements or a table of values The output results of some input output windows can be displayed as a graph or plot The plot of the out
148. nt selection list After selecting one or more airplanes this button will show the weight fraction data of the selected airplanes By pressing Calculate the average weight fractions will be calculated If a weight component fraction for a selected airplane is not available the representative box will be blank and this component will not be used in determining the average By scrolling down the list the remaining selected airplane data can be shown in case they did not appear on the current screen Input and output menus will appear In this menu the Class I component weight estimation is completed The estimation is based on the calculated average weight fractions By selecting Calculate the estimated component weights will be displayed When the calculated empty weight does not match the user supplied empty weight calculated in Weight Sizing an adjustment is Il 67 made which is distributed over all items in proportion to their component weight values listed in the last column 1 4 2 Center of Gravity This option allows the user to calculate the airplane center of gravity by entering weight components and their locations in a preformatted table The Center of Gravity module consists of three separate submodules After invoking Center of Gravity a menu appears with the following options e Empty Weight With this option empty weight components and the associated center of gravity coordinates can be tabulated The center of gr
149. o define empennage surfaces before the input output window for longitudinal stability calculations is displayed Input output windows contain a command bar at the top of the window an input group and an output group See Figure 2 2 The command bar contains a menu of buttons one for each command available to the input output window The input output window command bar is described in Subsection 2 1 3 Main Window Input Output Window Command Bar Input Group Spin Edit Element Input Output Table Output Group y a E dol Notes Copy WME P nt Almosphere Help Fight Cond Res Save File Corfigure on ACetheson ASetup DARcorporation Advanced Alrcraft Analysis 3 2 Project 05 01 09 11 58 AM Figure 2 2 Input Output Window Input output windows contain one or more input output elements Figure 2 3 shows an input output element The input output element contains the following PART I 1 7 e Variable Symbol e Edit Box for keyboard input e Unit SI or British e Info button e GoTo lt button e Work Pad button Variable Symbol Edit Box Info Button Unit Go To Button Notes Button Figure 2 3 Input Output Elements When the cursor is positioned over an input output element a brief description of the parameter is displayed When the cursor is located over the edit box of the input output element it appears as a vertical bar When the edit box is selecte
150. o the aerodynamic center shift due to nacelles 2 6 6 Aerodynamic Center Shift due to Pylons This submodule calculates the location of the pylon Aerodynamic Center based on the pylon geometry and the location on the airplane 2 6 7 Aerodynamic Center Shift due to Power Effects This submodule calculates the shift in aerodynamic center due to power effects 2 6 8 Wing Aerodynamic Center This submodule calculates the location of the wing Aerodynamic Center I 98 Aerodynamics 2 6 9 Horizontal Tail Aerodynamic Center This submodule calculates the location of the horizontal tail Aerodynamic Center 2 6 10 Vertical Tail Aerodynamic Center This submodule calculates the location of the vertical tail Aerodynamic Center 2 6 11 Canard Aerodynamic Center This submodule calculates the location of the canard Aerodynamic Center 2 6 12 V Tail Aerodynamic Center This submodule calculates the location of the v tail Aerodynamic Center 2 6 13 Ventral Fin Aerodynamic Center This submodule calculates the location of the ventral fin Aerodynamic Center 2 6 14 Aerodynamic Center of the Airplane This submodule calculates the total airplane Aerodynamic Center based on lifting surface Aerodynamic Center locations and shift contributions of other components such as fuselage nacelles etc 2 7 Power Effects The purpose of the Power Effects submodule is to determine the effects an operating engine has on the aircraft The power effect
151. odology used for finding the Aerodynamic Center can be found in Chapter 8 of Reference 7 Once the Aero Center button is selected a new menu appears allowing the user to calculate the aerodynamic center shift due to the fuselage nacelles stores tailbooms floats wing horizontal tail vertical tail canard v tail and aerodynamic center for the airplane The methodology used to calculate the aerodynamic center shift due to the fuselage nacelles stores and tailbooms can be found in Section 8 2 5 3 of Reference 7 Procedures outlining the use of each aerodynamic center shift submodules are presented in the following subsections 2 6 1 Fuselage For calculating the aerodynamic shift due to the fuselage 2 6 2 Nacelle For calculating the aerodynamic shift due to nacelles 2 6 3 Store For calculating the aerodynamic shift due to stores 2 6 4 Tailboom For calculating the aerodynamic shift due to tailbooms 2 6 5 Float For calculating the aerodynamic shift due to floats 2 6 6 Pylon For calculating the aerodynamic shift due to pylons 2 6 7 Power For calculating the aerodynamic shift due to power 2 6 8 Wing For calculating the aerodynamic shift due to the wing II 96 Aerodynamics 2 6 9 Horizontal Tail For calculating the aerodynamic shift due to the horizontal tail 2 6 10 Vertical Tail For calculating the aerodynamic shift due to the vertical tail 2 6 11 Canard For calculating the aerodynamic shift due to the canard 2 6 12 V Tail
152. odule the powerplant component weights can be estimated with the methods of Reference 5 The powerplant weight components menu consists of the following options e Propeller e Engine e Fuel System e Air Induction e Propulsion Syst e Total Powerplant To estimate propeller weight This option is only available for piston turboprop and propfan engines To estimate engine weight The engine weight includes engine exhaust cooling supercharger afterburners thrust reversers and lubrication systems To estimate fuel system weight To estimate air induction weight The air induction component weight includes inlet ducts other than nacelles ramps spikes and associated controls To estimate propulsion system weight The propulsion system weight components include engine controls starting systems propeller controls and provisions for engine installation To estimate total powerplant system weight 1 5 3 Fixed Equipment Component Weight Estimation In this module the fixed equipment component weights can be found with the methods of Reference 5 The fixed equipment weight components menu consists of the following options e Flight Control e Hydr Pneum Syst e Instr Avion Elec e Electrical Syst Aerodynamics To estimate the weight of the flight controls systems To estimate the weight of the hydraulic and pneumatic systems To estimate the weight of the instruments avionics and electronics The weight es
153. of AAA displays the support and services and description on how to use the online help Tip of the Day Displays the tip of the day Displays the DARcorporation Frequently Asked Questions web page 2 3 1 File Import The following file types are supported for import into the software e Weight Table Take off vs Empty Weight regression coefficient tables from AAA versions 1 0 through 1 7 These files have the extension re e Geometry Geometry parameter files for use in AeroPack These ASCII files have the extension geo e ASCII Parameter files from the program ACE IT by Tecolote Research These files have the extension txt 2 3 2 File Export e Weight Table Take off vs Empty Weight regression coefficient table These files have the extension re e Geometry Geometry parameter file for use in AeroPack These ASCII files have the extension geo I 48 PART I e ASCII Parameter file from the program ACE IT by Tecolote Research These files have the extension txt e S amp C Derivatives ASCII formatted file containing all Stability and Control Derivatives and Data for all flight conditions of the current project These files have the extension txt PART I I 49 PART I 3 Input Devices for the Software The following devices are used as input for the software mouse calculator and keyboard The devices are described in this chapter 3 1 Operation of t
154. of loads due to a mass that is distributed over the structural component Section 8 4 1 2 e User Loads For calculation of user specified loads acting on a structural component Section 8 4 1 3 Used for the fuselage only e Aerodynamic For calculation of the aerodynamic loads acting on a structural component Section 8 4 1 3 Used for all components except fuselage e Load Stations For calculation of the load stations used for each structural component Section 8 4 1 4 e Gear Engine Misc For calculation of the loads that result from landing gear engine or miscellaneous reactions Section 8 4 1 5 e Total Internal For calculation of the total internal loads acting on the structural component Section 8 4 1 6 Different concentrated distributed user aerodynamic landing gear engine or miscellaneous loads can be defined for each of the structural components 8 4 1 1 Concentrated Weights The Concentrated Weights module is used to define the weight and location of masses acting at concentrated positions on a structural component When Conc Weights is selected the following options will appear 8 4 1 1 1 Add Component 8 4 1 1 2 Select Compon 8 4 1 1 3 Copy Table 8 4 1 1 4 Conc Weights Loads II 173 8 4 1 1 1 Add Concentrated Weight Component This option is used to define weight components other than those found in the weight and balance tables of the Weight module see Chapter 1 The number of new objects is chosen
155. og is displayed which allows the user to save the exported data as a geo file Specify the folder and file name and click OK to save the file The file can then be imported into Shark SharkFX AP See the Shark SharkFX AP manual for import instructions H 124 Geometry 5 Propulsion Module 5 1 General Description In this Module the installed power and thrust of airplanes can be calculated In addition to the installed power thrust estimation this module also provides options for e Inlet and nozzle sizing e Estimation of inlet pressure recovery e Estimation of inlet and nozzle drag Calculations in this module are based on the methodologies outlined in Chapter 6 of Reference 7 Use of the Propulsion module options will be described in the following sections The assumption will be made that the following uninstalled propulsion characteristics are known e For piston engines Shaft horsepower data for a specific flight condition e For turboprop and propfan engines Shaft horsepower and thrust data for a specific flight condition e For jet engines Thrust data and engine mass flow data for a specific flight condition e For propeller driven airplanes Power coefficient and propeller efficiency or thrust coefficient for a specific flight condition Furthermore the quantity of the engine power extraction needs to be known 5 2 Type of Propulsion Before the Propulsion module can be accessed some basic parameters m
156. omatically added to the Select Airplane menu under the Airplane Category labeled User Defined e Select Airplane Allows the user to select existing airplanes from the program database for use in the Class I moment of Inertia estimation All the available airplanes in all categories can be displayed The user may select any combination of these to use in the Inertia estimation by adding them to the current selection list Moment of Inertia After selecting one or more airplanes this button will display an input output window The average radii of gyration are already calculated and will be shown in the input section If data are not available these airplane radii of gyration are not used in averaging The user needs to supply the remaining data and select Calculate to generate the moments of inertia I 70 Aerodynamics 1 4 4 Radii of Gyration This option allows the user to estimate the radii of gyration of the airplane using moment of inertia values The user provides the inertia data gross weight span and length and the radii of gyration are calculated for each of the three axes 1 5 Class II Weight The purpose of this module is to present a Class II method for estimating airplane component weights The Class II estimation methods used in this module are based on those described in Reference 5 and Reference 6 These methods employ empirical equations that relate component weights to airplane design characteristics In the Class Il wei
157. on the System Setup toolbar the Date Time Setup dialog box is displayed The Date Time Setup dialog box is displayed in Figure 2 27 The dialog box allows the user to choose the format of the date and time displays on the status bar Date and Time Format Date Time 07 31 06 5 08 PM x HE X Cancel 7 Help Figure 2 27 Date Time Setup Dialog Box PART I I 41 When the user selects Project on the System Setup toolbar the Project Name Setup dialog box is displayed The Project Name Setup dialog box is displayed in Figure 2 28 The dialog box allows the user to specify a project name or other description which will be displayed on the status bar Project Name Project Name Advanced Aircraft Analysis 3 1 Project Figure 2 28 Project Name Setup Dialog Box When the user selects Company on the System Setup toolbar the Company Name Setup dialog box is displayed The Company Name Setup dialog box is displayed in Figure 2 29 The dialog box allows the user to specify a company name or other description which will be displayed on the status bar Company Name Company Name DARcorporation X Cancel 7 Help Figure 2 29 Company Name Setup Dialog Box When the user selects Printer on the System Setup toolbar the Print Setup dialog box is displayed The Print Setup dialog box is displayed in Figure 2 30 The dialog box allows the user to specify the default printer and printer options used when printouts are ge
158. opics The three types of help topics and the method of accessing them are described in the following sections 6 1 User s Manual topics 6 2 Variable Info topics 6 3 Theory topics 6 1 User s Manual Topics These topics describe the functionality of the various windows toolbars and dialog boxes described in Chapter 2 of this part of the User s Manual These topics can be accessed in the following ways e When there are no application input output or plot windows open in the main window select the Help button on the main toolbar see Subsection 2 2 1 e When there are no application input output or plot windows open in the main Window select an item in the Help menu on the main window menu bar see Section 2 3 e Select the Help button on any dialog PART I I 59 6 2 Variable Info Topics Every variable that is displayed in an input output element see Subsection 2 1 2 has a topic that defines that variable Whenever possible typical values for the variable and a graphical depiction of the variable are shown in the variable info topic The variable info topic can be displayed in two ways e Select the Info button on the input output element If the Info button is not visible on the input output element the Options button on the System Setup toolbar see Subsection 2 2 4 can be selected to change the input output elements display properties e Select the Info button on the calculator see Section 3 2 6 3 Theory To
159. ormance curves must first be defined The clean configuration drag polar must also be known See Chapter 2 Part III of the user s manual Once the Climb option has been selected another menu appears showing the climb options e Climb Perform Contains rate of climb climb gradient and specific excess power e Time to Climb Represents the time to climb from one altitude to another When either the Climb Perform or Time to Climb options have been selected the Class II Climb input output window is displayed The methodology used in the climb analysis can be found in Section 5 3 of Reference 8 II 106 Performance 3 4 6 Cruise To run the Cruise submodule the thrust or power versus speed performance curves must first be defined The clean configuration drag polar must also be known See Chapter 2 Part IM Once the Cruise option has been selected another menu appears showing the cruise and range options 3 4 6 1 Maximum Speed 3 4 6 2 Range 3 4 6 3 Endurance 3 4 6 4 Payload Range 3 4 6 1 Maximum Cruise Speed When the Max Cruise Speed option has been selected a new window appears showing the Class II maximum cruise speed input and output parameters The program iterates over thrust power and speed to solve the cruise equations of motion until the required thrust power is equal to the available thrust power The maximum cruise speed is determined by locating the intersection of the thrust power required versus speed and thrust p
160. orming certain actions The first type of message is informative The second type of system message is automatically given by the system These messages usually contain critiques of user input These contain warnings or guidelines The user still has the option to ignore such warnings or guidelines However in the case where such an act could lead to a mathematical problem i e divide by zero the system ignores the user s request 3 5 Flexible Input One of the features of the program is the capability to bypass certain computation sequences The software is designed to ignore the computation sequence and make use of user defined values instead This feature is not always available and should be used with caution Using this feature demands a clear understanding of the computation sequence in the application module PART I I 53 For example I A B and C are input values into the calculation module I D and E are output values of the calculation module HI D is calculated by D f A B IV E is calculated by E f C D In a normal case the user provides all input values A B and C and the application module calculates the output parameters D and E with equation HI and IV Consider the case where the user does not want to calculate D from equation III but wishes to provide his her own value and still wants to use the system logic to calculate E To do this it is sufficient to define the C value in the input group and the desired D
161. ow The System Setup toolbar buttons can be used to manage the program environment The functionality of the buttons on the System Setup toolbar is described in Table 2 8 El Units Date Time Project Company Printer Calculator Options Figure 2 25 System Setup Toolbar Table 2 8 System Setup Toolbar Buttons Units Select British or S I units for the input and output parameters Date Time Select the date and time format in the status bar Date Time Project Specify the name of the project to be displayed in the status bar Company Specify the company name to be displayed in the status bar Printer Access the system print manager to define various printer attributes Printer Calculator Select the calculator type Standard or RPN Calculator I 40 PARTI Table 2 8 Contd System Setup Toolbar Buttons Options Select the size of the toolbar buttons Choose whether parameter info and notes buttons are displayed on input output elements see Subsection 2 1 2 Choose Dptions whether to save WMF to file and specify the length of the recovery project auto saving interval When the user selects Units on the System Setup toolbar the Units Setup dialog box is displayed The Units Setup dialog box is displayed in Figure 2 26 The program allows the user to work in the British or SI Metric system of units Units Type Figure 2 26 Units Setup Dialog Box When the user selects Date Time
162. ow sei cccs ic 5 ciieteckcvtoes etused nce cs abe cbiedous ooscee ep daspouecs sasbencblebeve cvuowespessensegs 183 10 3 AMPR Weight errors abesieb atin ase EEEE EE Tr TE ES ae 184 104 RED FIRB Costos sovscons tenet eden a 184 10 5 Prototype Costienorioas Bhs iin estes GALS Le ee RRR ea 185 10 6 A O AN 185 10 7 Operating Cost for Military AlrplaneS oooocnnnonnnonnconocononononnnonnncnn crac onanonnnonn nc nocnnocnnos 185 10 8 Operating Cost for Civil Airplanes oooccnncnnonnonnnonoconccononnnonnnonnncnn oran oran nnnnonnnonnccnnocnnos 186 10 84 Block Dti ni 186 10 8 2 Direct Operating Cost cooooocccocconncononononanonanonncnnoco nono no cono nn ErenEv E n EES 186 10 8 3 Indirect Operating Cost s c ccssccsccccsceesdesssestes sschsctstsssceevaspesesstssecdeessvbsesedsvesss 187 10 84 Program Operating Costs e s2 sceveessasevadeay pevecusessesdead event eosdenyedeer shes isn 187 10 9 Life Cycle Cost a a eh de Ad e 188 10 10 Price Data vis sscsssereinn eenia ear EEr EEE E EEE EE E TEE ES 188 REFERENCE ii SA at 189 X Table of Contents Table of Contents XI PART I Software System Organization This part describes the general functions of the mouse and program options used in the software and gives an overview of the different types of files used 1 Introduction The software provides a powerful framework to support the iterative and non unique process of aircraft preliminary design It allows students and prelim
163. ower available versus speed curves If a solution is not obtained the plotting option enables the user to quickly see the relative positions of the two curves 3 4 6 2 Range The methodology used in the range analysis can be found in Section 5 4 of Reference 8 Once the Range submodule has been selected another menu appears showing the range options e Constant Speed For the range at constant speed When the Constant Speed option has been selected a new window appears showing the Class II Range constant speed input and output parameters The angle of attack is iterated to solve the cruise equations of motion until the velocity matches user supplied input value e Constant Alt For the range at a constant altitude When the Constant Alt option has been selected a new window appears showing the Class II Range constant altitude input and output parameters If the lift Performance II 107 coefficient in the input menu is left undefined the lift coefficient for maximum range at constant altitude shown in the output will be used to determine the range 3 4 6 3 Endurance To run the Endurance submodule the thrust or power versus speed performance curves must first be defined The clean configuration drag polar must also be known Once the Endurance option has been selected another menu appears showing the endurance and loiter options e Constant Speed For the endurance at a constant speed In this option the angle of attack is i
164. owing options in the Classification dialog box e Class I Small light airplanes e Class II Medium weight low to medium maneuverability airplanes e Class II Large heavy low to medium maneuverability airplanes e Class IV High maneuverability airplanes Once the airplane certification type airplane flight phase and airplane class have been selected the input output window for the longitudinal mode checking is displayed In this submodule the user can compare the phugoid and short period mode with flying qualities as specified in MIL F 8785C After the input is defined the output group will be filled The output group displays data for the selected flight phase category When the user selects an output parameter information concerning that parameter can be accessed by selecting Help on the calculator Once the longitudinal mode checking has been performed the user can get a plot of the short period frequency requirements as specified in MIL F 8785C by selecting the Plot button A description of the standard plot options can be found in Section 2 1 4 of Part II of this user s manual The short period frequency characteristic of the current aircraft configuration is also plotted 7 3 3 Longitudinal Stability Derivative Sensitivity Analysis The Sensitivity option provides the capability to perform a sensitivity analysis for all longitudinal derivatives weight and inertia parameters The resulting sensitivity graphs show how different modes
165. pics Every calculation module of the software has a theory topic which explains the methods behind the calculation The theory topic can be displayed in two ways e Select the Theory button on the command bar of the input output window e Select the Help button on the main toolbar see Subsection 2 2 1 when the calculation module s input output window is open and active I 60 PARTI PART II Application Modules This part describes the function and use of the application modules The application modules are based on References 1 through 12 1 Weight Module 1 1 General Description The purpose of this module is to estimate airplane component weights and to determine whether or not the center of gravity of the airplane is within the desirable range for different loading and unloading scenarios The methods are based on Chapter 2 Sections 2 1 to 2 5 and Section 2 7 of Reference 1 Chapter 10 of Reference 2 and Chapter 2 of Reference 5 This module also contains the moments of inertia calculation based on Chapter 3 of Reference 5 Use of the Weight module options will be described in the following sections 1 2 Weight Main Window After invoking the Weight module 3 options are displayed e Weight Sizing The options provided in this module are discussed in Section 1 3 e Class Weight The options provided in this module are discussed in Section 1 4 e Class Il Weight The options provided in this module are discussed in Section
166. ping ratios and the associated transfer functions After choosing Transfer Function the user is presented with the following options e Stabilizer For calculating the stabilizer open loop transfer function e Elevator For calculating the elevator open loop transfer function e Canard For calculating the canard open loop transfer function e Canardvator For calculating the canardvator open loop transfer function e V Tail For calculating the v tail open loop transfer function e Ruddervator For calculating the ruddervator open loop transfer function e Flying Wing For calculating the flying wing open loop transfer function e Elevon For calculating the elevon open loop transfer function Il 156 Dynamics e None For calculating the airplane modes without a control surface NOTE The corresponding control surface must be defined by the user in the Control Surfaces Configuration dialog box before the dynamic analysis can be conducted After all input parameters are defined and the Calculate option is selected the longitudinal transfer function output parameters will be displayed and the transfer function will be shown in dialog boxes When certain output variables such as time constants damping ratios or frequencies cannot be calculated the program leaves a blank in the appropriate output box The transfer functions will be redisplayed when selecting the Calculate option These transfer functions can be printed either directly from the
167. played the first table will be displayed after selecting this button Clear Out Allows the user to erase all output parameters in the output EEA Clear Out Loo section of the calculation window Import Imports an Excel file in the same format as the table in the active Import T able window PART I I 11 Table 2 2 Contd Input Output Window Command Bar Functions Export Export input and output data to a text file ASCID or to an Excel Spreadsheet Theory Opens a Help window containing the calculation methods corresponding to the input output window see Section 6 3 Close Window Closes the input output window The window minimize button can be used to iconize the window if desired 2 1 4 Plot Windows The plot window contains a graphical representation of a calculation in an input output window Figure 2 7 shows a plot window of a Class I drag polar of a jet powered airplane The plot window contains a command bar at the top The functionality of the plot window command bar is described in Subsection 2 1 5 EN Advanced Aircraft Analysis 3 2 Project1 aaa Flight Condition 1 Elle Edit Window Help ATA Weight ZS Aerodynamics MI Performance em Geometry eX Propulsion SP Stab amp Control NA Dynamics ar Loads SZ Structures BES Cost EN Airplane Drag Polar Clean Configuration Flight Conditio
168. ps in landing configuration with gear up e Land Gear Down For flaps in landing configuration with gear down e OEI For one engine inoperative configuration e Current Fit Cond For the current airplane flight condition e All Polars Calculates all the drag polars for Class I Drag The methodology used to calculate the drag polar can be found in Chapter 3 of Reference 1 To calculate the drag polar the user first selects the switch corresponding to the configuration of interest The other five flight condition options will have similar input and output parameters The Class I Drag module relates the total airplane lift coefficient to the total airplane drag coefficient by the drag polar equation 1 2 Cp Ni oil i TAR EL Where Cp and Bpp are coefficients calculated by the program using the methods in Section 3 4 1 of Reference 1 The program allows the user to perform this calculation in different ways The standard method of computing C Dp and Bpp is to specify all the input parameters and select the Calculate button Equivalent parasite area is a function of the regression coefficients a and b Reference 1 Section 3 4 1 and are user defined in the input parameter section If a and b are defined the program automatically calculates the value of the equivalent parasite area However if the resulting equivalent parasite area or airplane wetted area is unacceptable the user has the capability of redef
169. put is presented in a plot window when the Plot button on the input output window command bar is selected The plot window is described in the Subsection 2 1 4 I 10 PARTI 2 1 3 Input Output Window Command Bar The input output window command bar is displayed at the top of the input output window The input output window command bar is shown in Figure 2 6 Each button in the command bar represents an action that can be performed in the input output window A command bar button is not displayed if its action is not available for the particular input output window The Close button in the command bar closes the input output window and is always displayed The remaining buttons that can be displayed in the input output command bar are shown and described in Table 2 2 Calculate HEE Clear Out BES gt Theory Close Plot E Export i Import T able A Next Nacelle x co Figure 2 6 Input Output Window Command Bar Buttons Table 2 2 Input Output Window Command Bar Functions Calculate Using the specified input the calculations for the input output window are performed The results of the calculations are displayed in the output parameters Plot Opens the corresponding Plot window when applicable Next Item If there are multiple tables of input for example tables for different nacelles tailbooms and stores needed for the calculation the next table of parameters will be displayed If the last table is currently dis
170. r the instantaneous level turn option the input windows for both cases contain the same information Performance II 109 3 4 9 Glide To run the Glide submodule the thrust power versus speed performance curves must first be defined The clean configuration drag polar must also be known Once the Glide option has been selected another menu appears showing the Class II Glide options e Rate of Descent When the Rate of Descent option is selected another window appears showing input and output parameters The methodology used in the glide analysis can be found in Reference 8 e Time of Glide When the Time of Glide option is selected another window appears showing input and output parameters The methodology used in the glide analysis can be found in Reference 8 e Range of Glide When the Range of Glide option is selected another window appears showing input and output parameters The methodology used in the glide analysis can be found in Reference 8 When the Plot option is selected the selected parameter is plotted with respect to angle of attack The rate of descent time in the air and glide range versus angle of attack plots are generated using the input from each respective input output window A detailed explanation of the plot options can be found in Part II Chapter 2 of the user s manual 3 4 10 Landing Distance To run this submodule the landing configuration drag polar must first be known Once the Landing Distance option
171. rag moment and thrust derivatives When the Steady State button is selected a new menu appears with the following options e Cp o Ch e Corr 6 3 2 Speed Derivatives This submodule can be used to estimate the steady state drag coefficient This submodule can be used to estimate the steady state lift coefficient This submodule can be used to estimate the steady state pitching moment coefficient This submodule can be used to estimate the steady state thrust force coefficient This option is not available if None is specified as the powerplant This submodule can be used to estimate the steady state thrust pitching moment coefficient This option is not available if None is specified as the powerplant The methodology used to calculate the speed related derivatives can be found in Section 10 2 1 of Reference 7 Once the Speed button is selected a new menu appears with the following five speed related derivative options e Cp e CL Dynamics This submodule can be used to estimate the drag coefficient due to speed derivative This submodule can be used to estimate the lift coefficient due to speed derivative H 131 e Cn This submodule can be used to estimate the pitching moment coefficient due to speed derivative This submodule can be used to estimate the thrust force Ty u coefficient due to speed derivative This option is not available if None is specifi
172. railing edge device from the list given in Figure 2 19 e Leading Edge Device The user may specify the type of leading edge device from the list given in Figure 2 19 PART I I 33 High Lift Devices Number of High Lift Devices per Wing 5 Trailing Edge Device Leadling Edge Device Plain Flap C Slat Split Flap C Krueger Flap Single Slotted Flap Type Double Slotted Flap Type Il Double Slotted Flap Fowler Flap Triple Slotted Flap Drooped Aileron 2 Triple Slotted x Cancel Help Figure 2 19 High Lift Devices Dialog Box Landing Gear Dialog Box When the Gear button is selected the Landing Gear dialog box is displayed Figure 2 20 shows the Landing Gear dialog box Descriptions of the dialog box functions follow I 34 PART I Landing Gear Number of Gears 3 Position Attachment Retraction Nose C Wing C Retractable C Fuselage C Fixed C Tail C Nacelle C Outrigger C Tailboom C Other ra x Cancel Help Figure 2 20 Landing Gear Configuration Dialog Box e Number of Gears The user may specify the number of gears Each strut should be considered as one landing gear e Position The relative position nose main or tail of each gear can be specified e Retraction The retraction capability retractable or fixed of each gear can be specified e Attachment The location wing fuselage nacelle tailboom or other of where the gear is attached can be specified Structure Dialog Box When the Stru
173. res BSS Cost lt lt Stability and Control of x Derivatives I Analysis Wind Tunnel Menu Bar Stability Control and Hingemoment Derivative Long Stability I Lat Dir Stability Long Control Lat Dir Control Hingemoment Recalculate All Lateral Directional Stability Derivatives Sideslip Sideslip Rate Roll Rate Yaw Rate File Toolbar Main Toolbar A a z ay Status Bar J A RE me lt 7 E New Open Save Save As Delete Flight Cond Recalculate Notes Copy WMF Print Atmosphere Help Exit File AContiguration Certification Setup DARcorporation Advanced Aircraft Analysis 3 2 Project 05 01 09 11 18 AM Company Name Project Name Date and Time Figure 2 1 The Main Window Three types of windows can be contained within the main window There are application windows input output windows and plot windows Application windows input output windows and plot windows are child windows and are always displayed within the main window Descriptions of each of these window types and their components are presented in the following subsections 2 1 1 Application windows 2 1 2 Input Output windows 2 1 3 Input Output window command bar 2 1 4 Plot windows 2 1 5 Plot window command bar 1 4 PARTI 2 1 1 Application Windows When one of the application buttons at the top of the main window is selected the corresponding application window is displ
174. rice TT 188 Cost REFERENCES 1 Roskam J Airplane Design Part I Preliminary Sizing of Airplanes DARcorporation Lawrence Kansas 2005 2 Roskam J Airplane Design Part II Preliminary Configuration Design and Integration of the Propulsion System DARcorporation Lawrence Kansas 2004 3 Roskam J Airplane Design Part III Layout Design of Cockpit Fuselage Wing and Empennage Cutaways and Inboard Profiles DARcorporation Lawrence Kansas 2002 4 Roskam J Airplane Design Part IV Layout Design of Landing Gear and Systems DARcorporation Lawrence Kansas 2004 5 Roskam J Airplane Design Part V Component Weight Estimation DARcorporation Lawrence Kansas 2003 6 Raymer D Aircraft Design A ConceptualApproach 4 Edition American Institute of Aeronautics and Astronautics Inc Reston Virginia 2006 7 Roskam J Airplane Design Part VI Preliminary Calculation of Aerodynamic Thrust and Power Characteristics DARcorporation Lawrence Kansas 2004 8 Roskam J Airplane Design Part VII Determination of Stability Control and Performance Characteristics FAR and Military Requirements DARcorporation Lawrence Kansas 2006 9 Roskam J Airplane Design Part VIII Airplane Cost Estimation Design Development Manufacturing and Operating DARcorporation Lawrence Kansas 2006 10 Lan C E Roskam J Airplane Aerodynamics and Performance DARcorporation Lawrence Kansas 2003 11 Roskam J Ai
175. ro and no additional menus appear on the screen This submodule can be used to estimate the rolling moment coefficient due to spoiler derivative H 143 e Crs This submodule can be used to estimate the yawing moment sp coefficient due to spoiler derivative 6 6 4 Vertical Tail Related Derivatives Once the Vert Tail button is selected a new menu appears with the following three vertical tail related derivative options Cy This submodule can be used to estimate the sideforce coefficient Vv due to vertical tail incidence derivative C This submodule can be used to estimate the rolling moment V coefficient due to vertical tail incidence derivative Cn This submodule can be used to estimate the yawing moment V coefficient due to vertical tail incidence derivative 6 6 5 Rudder Related Derivatives The methodology used to calculate the rudder related derivatives is based on theory in Sections 10 3 2 and 10 3 8 of Reference 7 Once the Rudder button is selected and the vertical tail configuration has been chosen a new menu appears with the following three rudder related derivative options ec ys This submodule can be used to estimate the side force coefficient r due to rudder derivative e Cis This submodule can be used to estimate the rolling moment r coefficient due to rudder derivative ec This submodule can be used to estimate the yawing moment n y g coefficient due to rudder derivative 6
176. rol systems The options provided in this module are discussed in Sections 7 6 through 7 9 Once the Dynamics module has been selected in the main window a new menu appears with three submodules 7 3 Longitudinal For estimation of the longitudinal dynamic characteristics 7 4 Lateral Direct For estimation of the lateral directional dynamic characteristics 7 5 Roll Coupling For the roll pitch yaw coupling effect of the dynamic analysis Dynamics H 155 Once the Control module has been selected in the main window a new menu appears with four submodules 71 7 Root Locus s For analysis of a single and double loop feedback control system 7 8 Bode Method For development of an airplane frequency response Bode plot 7 9 Human Pilot For estimation of a human pilot transfer function 7 10 Erase T F For erasing one or all of the calculated transfer functions Section 7 6 presents the two methods used for defining transfer function and for loading data into the Control analysis module 7 3 Longitudinal Dynamics After selecting the Longitudinal option the following options will be displayed 7 3 1 Transfer Function For transfer function computation 7 3 2 Flying Qualities For flying qualities checking 7 3 3 Sensitivity For longitudinal derivative sensitivity analysis 7 3 1 Calculate Transfer Function Transfer Function is used to calculate the longitudinal characteristic equations associated modes including frequencies and dam
177. rplane Flight Dynamics and Automatic Flight Controls Part I DARcorporation Lawrence Kansas 2003 12 Roskam J Airplane Flight Dynamics and Automatic Flight Controls Part II DARcorporation Lawrence Kansas 2003 NOTE References 1 through 5 and 7 through 12 are published by DARcorporation 1440 Wakarusa Drive Suite 500 Lawrence Kansas 66049 USA Phone 785 832 0434 Fax 785 832 0524 www darcorp com REFERENCES 189 13 14 15 16 17 18 Lofts M Thomas H Analysis of Wind Tunnel Data on Horn Balance RAE report Aero 1994 1944 Thomas H Comments on Wind Tunnel Investigation of Rounded Horns and of Guards on a Horizontal Tail Surface RAE technical note Aero 1641 1945 Hildebrand F B A Least Square Procedure for the Solution of the Lifting Line Integral Equation NACA technical note 925 1944 Crandall S M Swanson R S Lifting Surface Theory Aspect Ratio Corrections to the Lift and Hinge Moment Parameters for Full Span Elevators on Horizontal Tail Surfaces NACA technical note 1175 1947 MIL A 8861 ASG Military Specifications Airplane Strength and Rigidity Flight Loads May 1960 DARcorporation AeroPack User s Manual DARcorporation Lawrence Kansas August 2006 190 REFERENCES
178. rudder pedal force and the rudder pedal force gradient are calculated 6 15 Wing Location In this module the C G range and Static Margin are plotted as a function of the wing location allowing the user to quickly make a decision on the wing location which would result in a desired Static Margin and a C G range H 154 Dynamics 7 Dynamics Module 71 General Description The purpose of the Dynamics module is to help the user analyze the open loop dynamic characteristics of the airplane in a given flight condition and to help the user analyze single and double loop feedback control systems of the airplane If the open loop dynamic characteristics of the airplane are known the Control analysis submodule can be used to perform root locus analyses The Control analysis can also be used to analyze a system open loop transfer function in the frequency domain The methodology used to analyze open loop dynamic characteristics can be found in Reference 11 The methodology used to analyze feedback control systems can be found in Reference 12 Use of the Dynamics module options will be described in the following sections 7 2 Dynamics Main Window After selecting the Dynamics module the user can select from the two options displayed e Dynamics For analyzing the airplane open loop dynamic characteristics for a given flight condition The options provided in this module are discussed in Sections 7 3 through 7 5 e Control For analyzing feedback cont
179. s 136 6 5 1 Stabilizer Related Derivatives ooooonoccconononococonononnncconanonnncconnnonnncconenononcnnnos 137 6 5 2 Elevator Related DerivativeS oononccnncnnonnnnnccconnncnoccconncnonccnnnnnnonccnnnnnnnnccno 137 6 5 3 Elevator Tab Related Derivatives ooooocnnoccconononocccononononcconononnncnnnnconancnncnnnns 138 6 5 4 Canard Related Derivatives cee eeseceseceeseeceseceeseeceeeeeeneeceeeeeeaeecetreeenaeens 138 6 5 5 Canardvator Related DerivativeS ooocccnnnncnncccnonncnoccnnnnnnoncconnnnnnnccnnnnncnnccno 139 6 5 6 Canardvator Tab Related Derivatives ooonnnocinnncconcnononcconononnnconcconnnncnncnonns 139 6 5 7 V Tail Related Derivatives cee eesccceeeceesceceneeeenceceeeeeeneeceeeeeneecetreeenaeens 139 6 5 8 Ruddervator Related Derivatives 0 0 0 eeeecceesseceeeeceseceeececereeeneecenreeeneeees 140 6 5 9 Ruddervator Tab Related Derivatives oooonnccnnonconccnoonnconcnonnncnonccononnnnnccnno 140 6 5 10 Flying Wing Related Derivatives ooooonocncocnconocononanonnncnnonanonanonn conc cnnncnnncnnnnns 140 6 5 11 Elevon Related Derivatives oococnncccnncocoocnconcnconnnconcnconnncnnn cono nocnnnnconnncnnnnnnns 141 6 5 12 Elevon Tab Related Derivatives cesceesecsseceecceceseceeneeceeeeeneecenreeenaeess 141 6 6 Lateral Directional Control Derivatives ooooocnncccnonccnoncconnncnnnccnnnnnnon ccoo nnnnon cnn nnnonnccno 142 Table of Contents VI 6 6 1 Aileron Related Derivatives ccc
180. s submodule is only available when the Include Power Effects field is selected in the Flight Condition Dialog Box The methodology used for determining Power Effects can be found in Chapter 8 of Reference 7 and also Reference 17 2 8 Ground Effects This module calculates the effect of the ground on airplane lift and pitching moment as well as calculating the effect of variation of the airplane s height above the ground This module is Aerodynamics Il 99 available only when the Include Ground Effects field is selected in the Flight Condition Dialog Box 2 9 Dynamic Pressure Ratio This module calculates the effects of free stream velocity on the horizontal tail vertical tail or v tail as it varies with angle of attack The contribution of power effects is also included in the analysis II 100 Aerodynamics 3 Performance Module 3 1 General Description The purpose of the Performance Sizing submodule is to allow for a rapid estimation of those airplane design parameters having a major impact on airplane performance Airplanes are usually required to meet performance objectives in different categories depending on the mission profile Meeting these performance objectives normally results in the determination of a range of values for e Wing loading W S e Thrust loading T W or power loading W P e Airplane maximum lift coefficients C Lmax C Dos and C LRT The variables listed above are plotted in th
181. scribed in the following sections 10 2 Cost Main Window After invoking the Cost button seven options are displayed e AMPR Weight e R D T E Cost e Prototype Cost e Acquisition Cost Cost For estimation of the Aeronautical Manufacturers Planning Report AMPR weight which is defined on page 25 of Reference 9 This weight parameter is needed for estimation of the various costs in an airplane program The options provided in this module are discussed in Section 10 3 For estimation of the research development test and evaluation cost The options provided in this module are discussed in Section 10 4 For estimation of the cost of development manufacturing and flight testing of the prototypes This submodule is to be used only for those airplane programs that are not intended for eventual production The options provided in this module are discussed in Section 10 5 For estimation of the manufacturing and acquisition costs The difference between these costs is the profit made by the manufacturer The options provided in this module are discussed in Section 10 6 II 183 e Operating Cost military For estimation of the military airplane operating costs The options e Operating Cost civil e Life Cycle Cost e Price Data 10 3 AMPR Weight provided in this module are discussed in Section 10 7 For estimation of the civil airplane operating costs The options provided in this module are discussed in Section
182. sed after highlighting text in an edit box The three options described in Table 2 10 I 46 PART I Table 2 10 Edit Menu Options Delete the selected text and place in the Windows clipboard Copy the selected text into the Windows clipboard Paste the current text from the Windows clipboard into the edit box The Window menu options can be used to arrange application input output and plot windows within the main window The three Window menu options are described in Table 2 11 Table 2 11 Window Menu Options Size and arrange the open windows so that all of the windows can be seen on the screen Cascade Arrange all of the open windows so that the title bar of each window can be seen Arrange All If any windows are minimized within the main window this option will arrange them at the bottom of the main window In addition the Window menu displays a list of all open windows within the main window Selecting a window name within the list will cause that window to be displayed and become active The Help menu options can be used to access the software help system and online manual The five Help menu options are described in Table 2 12 PART I I 47 Table 2 12 Help Menu Options Help Displays the help topic relating to the currently active window This corresponds to the Help button on the main toolbar or the Theory button on the input output window Online Help Manual Displays the help on the internal layout
183. sed on its geometry 4 3 4 High Lift Device For calculation of the high lift device geometry trailing edge flaps slats Krueger flaps etc geometry 4 3 5 Aileron Tab For calculation of aileron geometry 4 3 6 Spoiler For calculation of the spoiler parameters 4 3 7 Chord Length For calculation of wing chord based on span wise location 4 3 8 Airfoil For calculation of wing airfoil parameters 4 3 9 Exposed For calculation of exposed wing geometry 4 3 1 Straight Tapered The Straight Tapered wing geometry submodule has four different sets of input and output variables Any one of these data sets can be accessed by using the appropriate menu button After calculating the output parameters the user may display the wing by selecting the Plot option 4 3 2 Cranked Wing This submodule can be used to calculate the equivalent straight tapered wing geometry for a cranked wing The equivalent wing is defined as a straight tapered wing with the same net wing area as the cranked wing and with the same tip chord length or same mean geometric chord The user can define a cranked wing by entering input data for all the panels comprising the wing Once all input is defined the user can calculate the equivalent straight tapered wing parameters by selection of the Calculate button A plot of the cranked wing can be obtained after selection of Geometry Tr 113 the Plot option 4 3 3 Fuel Volume Class I Class II and Cranked Wing In the Cl
184. sed vertical tail parameters 4 5 1 Straight Tapered The method for determining the geometry of a straight tapered vertical tail is identical to the method found in Section 4 5 1 of this manual 4 5 2 Cranked Vertical Tail The method for determining the geometry of a cranked vertical tail is identical to the method found in Section 4 5 2 of this manual 4 5 3 Volume Coefficient This vertical tail submodule calculates the volume coefficient of a vertical tail planform The methodology used to calculate the horizontal tail volume coefficient and related parameters is described in Chapter 8 of Reference 2 After calculating the output parameters the user may display the vertical tail geometry by selecting the Plot option 4 5 4 Rudder Tab In this submodule the geometry of the rudder is added 4 5 5 Chord Length In this submodule the chord length of the vertical tail at any span wise location can be Geometry HI 117 determined by using root chord length taper ratio and the given span wise location 4 5 6 Airfoil In this submodule the vertical tail airfoil parameters at the vertical tail mean geometric chord are calculated 4 5 7 Exposed In this submodule the exposed vertical tail parameters are calculated 4 6 Canard Geometry When the Canard module is selected the user is presented with the following options 4 6 1 Straight Tapered For calculation of straight tapered canard geometry 4 6 2 Cranked Canard For calcula
185. span wise location 4 3 8 Airfoil In this submodule the the airfoil parameters at the wing mean geometric chord is calculated 4 3 9 Exposed In this submodule the exposed wing geometry is calculated 4 4 Horizontal Tail Geometry When the Horizontal Tail module is selected the user is presented with the following options 4 4 1 Straight Tapered For calculation of straight tapered horizontal tail geometry 4 4 2 Cranked For calculation of the equivalent straight tapered horizontal tail geometry from given cranked horizontal tail geometry 4 4 3 Volume Coefficient For calculation of the horizontal tail Volume Coefficient 4 4 4 Elevator Tab For calculation of the elevator geometry 4 4 5 Chord Length For calculation of horizontal tail chord based on span wise location 4 4 6 Airfoil For calculation of airfoil parameters at the horizontal tail mean geometric chord 4 4 7 Exposed For calculation of exposed horizontal tail parameters Geometry r 115 4 4 1 Straight Tapered The method for determining the geometry of a straight tapered horizontal tail is identical to the method found in Section 4 4 1 of this manual 4 4 2 Cranked Horizontal Tail The method for determining the geometry of a cranked horizontal tail is identical to the method found in Section 4 4 2 of this manual 4 4 3 Volume Coefficient This horizontal tail submodule calculates the volume coefficient of a horizontal tail planform The methodology used to
186. sponding weight sensitivities Aerodynamics Il 65 1 3 6 Useful Load This module calculates the range for different payload scenarios starting with a minimum of 1 pilot and ending with the maximum payload specified in the Take off Weight module 1 3 7 Remarks 1 The payload weight excludes expended payload such as bombs and ammunition Payload dropped during the mission has to be specified in the Mission Profile Payload which is not dropped must be included in payload and crew weights 2 Refueled fuel weight during the mission must be specified in the Mission Profile 3 A change in expended payload and refueled fuel can only be made in the Mission Profile submodule since the location of the payload expenditure and refueling segment in the flight profile is essential for the calculation of the weights 1 4 Class I Weight The Class I Weight module consists of four separate submodules After invoking Class Weight a menu appears with the following options 1 4 1 Weight Fractions With this option the user can determine the aircraft component weights by using the weight fraction methods 1 4 2 Center of Gravity This option allows the user to calculate the airplane empty and current weights and center of gravity locations 1 4 3 Inertia Estimate With this option the user can calculate the airplane moments of inertia with radii of gyration methods 1 4 4 Radii of Gyration With this option the user can calculate the airplane radii of
187. ss the OK button An undefined variable appears blank in the input output parameter PARTI Clear Selecting this button will erase the current number from the calculator display In case the current number replaced an old number the old value will appear The number will not be cleared in the database SI British Selecting this button will cause the displayed unit to switch between the British and SI units The displayed value will also be converted to the selected unit In the input output element the value will be shown in the units set during the software installation or with the Units button in the System Setup toolbar 3 3 Keyboard The keyboard can be used in the software when the user is prompted for text input within a dialog window entering a number once the calculator is displayed or entering data directly into an input output element If the keyboard is used to enter data directly into an input output element selecting the lt Enter gt or lt Tab gt key will enter the currently displayed number into the database and move the cursor to the next element in the input or output group Keyboard shortcuts exist for the Calculate and Close icons found in the input output windows The lt F9 gt key will perform the calculation and lt F4 gt will close the active input output window 3 4 System Messages There are two types of messages which warn the user of possible pitfalls in the design process or demonstrate methods of perf
188. stimation of the total program operating cost Cost II 185 10 8 Operating Cost for Civil Airplanes Before selecting the Operating Cost button the user must specify the airplane civil military designation For civil aircraft a menu is displayed for estimating the operating cost of commercial airplanes The menu has the following options 10 8 1 Block Data For estimation of the time speed and distances corresponding to the mission from the point of beginning taxi before take off to ending taxi after landing 10 8 2 Direct For estimation of direct operating cost 10 8 3 Indirect For estimation of indirect operating cost 10 8 4 Program For estimation of total program operating cost 10 8 1 Block Data All the parameters needed for operating cost estimation can be calculated in the Block Data input output window After invoking the Block Data switch two options are displayed e Time amp Speed For calculating speed range and time parameters used in the operating cost estimation method e Utilization In this submodule the annual utilization and total annual block distance can be calculated 10 8 2 Direct Operating Cost In the Direct Operating submodule the user can estimate the direct operating cost DOC in US nm incurred while operating commercial airplanes After selecting the Direct Operating button five options are displayed e Flying Select this option for estimation of the direct operating cost of flying After invoking t
189. submodule can be used to estimate the lift coefficient due to angle of attack derivative e C This submodule can be used to estimate the pitching moment coefficient due to angle of attack derivative e C This submodule can be used to estimate the thrust pitching moment coefficient due to angle of attack derivative This option is only available for propeller powered airplanes NOTE There are two ways to enter values for the aerodynamic center shift due to the fuselage nacelles stores floats and tailbooms The user may 1 Enter the values for the aerodynamic center shift in terms of the wing mean geometric chord due to the fuselage nacelles stores floats and tailbooms in the Cing submodule if these values are known 2 Use the Aerodynamic Center module to calculate the values of the aerodynamic center shift due to the fuselage nacelles stores floats and tailbooms in terms of the wing mean geometric chord If the user chooses to use the Aerodynamic Center module first the values for the aerodynamic center shift will automatically be transferred into the Cing submodule input For a complete description of the Aerodynamic Center module see Part II Section 2 6 of this user s manual 6 3 4 Rate of Angle of Attack Derivatives The methodology used to calculate the rate of angle of attack related derivatives can be found in Section 10 2 3 of Reference 7 Once the A O A Rate button is selected a new menu appears with
190. t coefficient of the wing with flap deflection This option is described in Section 2 3 2 For calculation of the lift coefficients for nacelles and pylons Aerodynamics e Airplane For calculation of the airplane steady state lift coefficient and angle of attack The procedure to use each lift submodule is presented in the following sections The methods used in calculating the lift coefficient of all airplane lifting surfaces are the same 2 3 1 Wing Horizontal Tail Vertical Tail Canard V Tail The Wing Horizontal Tail Vertical Tail Canard and V Tail submodules have three options Cig To determine the lifting surface lift curve slope eg To determine the horizontal tail canard v tail downwash upwash gradients displayed only when the airplane is defined as having either a horizontal tail canard or v tail e CL amp Qo To determine the lifting surface zero angle of attack lift coefficient and zero lift angle of attack e CL max To determine the airfoil and lifting surface maximum lift coefficients This option is described in Section 2 3 1 1 e Linear Cr To determine the lift coefficient for a given angle of attack in the linear portion of the lift curve e Non Linear Cr To determine the lift coefficient for a given angle of attack in the non linear portion of the lift curve e Lift Distribution To determine the lifting surface spanwise lift distribution This option is described in Section 2 3 1 2
191. t is described in Section 4 4 of Reference 7 The Canard input output windows depend on the selected flow regime 2 4 2 5 V Tail Drag In this module the v tail drag coefficient can be calculated for a specific flight condition The v tail drag coefficient consists of a zero lift component and a lift dependent component The methodology used to calculate the v tail drag coefficient is based on theory in Section 4 4 of Reference 7 The V Tail input output windows depend on the selected flow regime 2 4 2 6 Ventral Fin Drag In this module the v tail drag coefficient can be calculated for a specific flight condition 2 4 2 7 Fuselage Drag In this module the fuselage drag coefficient can be calculated for a specific flight condition The fuselage drag coefficient consists of a zero lift component and a lift dependent component The methodology used to calculate the fuselage drag coefficient is described in Section 4 3 of Reference 7 The Fuselage input output windows depend on the selected flow regime 2 4 2 8 Nacelle Drag In this module the nacelle drag coefficient can be estimated The methodology used to calculate the nacelle drag coefficient is described in Section 4 5 of Reference 7 For estimation of the nacelle drag coefficient the nacelle drag consists of an isolated nacelle component and an interference component The Nacelle drag input output window depends on the selected flow regime and the propulsion type Before entering th
192. t phase Edit The user can change the name of the current flight phase Delete The user can delete a flight phase from the current project All information associated with the selected flight phase will be deleted Move The user can move a flight phase within the current project Copy The user can copy a flight phase within the current project PART I e Altitude The user enters the altitude corresponding to the defined flight phase e Temperature Increment The user can enter a temperature offset from standard atmosphere e Velocity The user enters the velocity for the defined flight condition British or S L units are automatically supplied depending on the setting in the Units dialog box see Figure 2 25 e Flap Deflection After defining trailing edge flap in the configuration dialog box the user enters the flap deflection angle corresponding to the flight condition e Current Weight The user enters the current weight of the aircraft corresponding to the defined flight phase e C G X Location The user enters the current Center of Gravity location along the X axis for the defined flight condition e C G Z Location The user enters the current Center of Gravity location along the Z axis for the defined flight condition e Load Factor The user enters the load factor corresponding to the defined flight phase e Stores After defining stores in the Configuration dialog box the user can specify which stores are on the airplan
193. tabases are used The first type of database is fixed and used by the program to retrieve information that does not change with use of the program The first type of database is discussed in Section 5 1 and 5 2 The second type of database is used to store parameter values and characteristics that change with use of the program A project consists of a collection of this type of databases Projects and their individual databases are described in Section 5 3 5 1 The System Database The system database is used to store information about parameters that do not change The system database is located in the folder lt dar folder gt Database where lt dar folder gt is the folder where the software is installed The system database should never be deleted or modified in any way The system database stores the following information about every numeric parameter used in the software e Name of the variable e Engineering Symbol of the variable e Name of the info topic for the variable e Conversion factor between British and SI units e Number of decimal places to display for the variable e Minimum allowable value for the variable e Maximum allowable value for the variable e Short definition of the variable that is displayed as described in Figure 2 3 5 2 File Compatibility with Previous Versions Files created with AAA versions 1 0 through 2 4 and other airplane design software by DARcorporation can be opened with this version of the softwar
194. te the horizontal stabilizer related derivatives can be found in Section 10 3 1 of Reference 7 Once the Stabilizer button is selected a new menu appears with the following three stabilizer related derivative options CD This submodule can be used to estimate the drag coefficient due to horizontal stabilizer incidence derivative e C e This submodule can be used to estimate the lift coefficient due to horizontal stabilizer incidence derivative Cmi This submodule can be used to estimate the pitching moment coefficient due to horizontal stabilizer incidence derivative 6 5 2 Elevator Related Derivatives The methodology used to calculate the elevator related derivatives can be found in Section 10 3 2 of Reference 7 When the Elevator button is selected a new menu appears with the following three elevator related derivative options Dynamics II 137 e Cos This submodule can be used to estimate the drag coefficient due e to elevator deflection derivative e Cr 5 This submodule can be used to estimate the lift coefficient due to e elevator deflection derivative e Cms This submodule can be used to estimate the pitching moment e coefficient due to elevator deflection derivative 6 5 3 Elevator Tab Related Derivatives The methodology used to calculate the elevator tab related derivatives can be found in Section 10 3 2 of Reference 7 When the Elevator Tab button is selected a new menu appears with the following
195. terated to solve the equations of motion until the velocity matches the user supplied input data e Constant Alt For the endurance at constant altitude If the lift coefficient in the Constant Alt input group is left undefined the lift coefficient for maximum endurance at constant altitude shown in the output group will be used to determine the endurance The methodology used in the endurance analysis can be found in Section 5 5 of Reference 8 3 4 6 4 Payload Range Diagram Once the Payload Range submodule has been selected another menu appears showing the payload range diagram options e Constant Speed For the payload range diagram for range at constant speed e Constant Alt For the payload range diagram for range at constant altitude The methodology used in the payload range diagram analysis can be found in Section 5 4 of Reference 8 3 4 7 Dive amp Descent To run the Dive amp Descent submodule the thrust or power versus speed performance curves must first be defined The clean configuration drag polar must also be known The purpose of H 108 Performance this submodule it to determine the flight path angle for a given dive speed When the Dive 8 Descent option is selected another window appears showing the Class II dive and descent input and output parameters for a given speed The angle of attack is iterated to solve the dive and descent equations of motion until the velocity matches the user supplied input data
196. termine electrical power requirements the user should prepare an Electrical Power Load Profile as described on page 320 in Reference 7 H 126 Dynamics e Pneumatic For estimation of extracted pneumatic power This option will not be offered for a piston engine selection A list of airplane services that require bleed air as the source of power must be prepared e Total For estimation of the power required for mechanical electrical and pneumatic operation of the airplane 5 5 Inlet Design The purpose of the Inlet Design submodule is to estimate the streamtube cross section at the inlet the inlet pressure recovery and the inlet extra drag After invoking Inlet Design the following options will be displayed e Inlet Area To calculate the required inlet area NOTE The definitions of the inlet geometric parameters depending on engine and inlet choice are shown in Section 6 2 of Reference 7 e Pressure Loss To calculate the pressure loss in the inlet Before invoking Pressure Loss the inlet type must be specified in the engine dialog box as follows O Plenum For all engine types except supersonic jet engines O Straight Through For all engine types e Manifold Pressure To calculate the manifold absolute pressure 5 6 Nozzle Design In this submodule some nozzle design objectives are provided After invoking Nozzle Design the exhaust exit area window appears asking for take off shaft horsepower 5 7 Installed Data I
197. timation in this module is obsolete for modern EFIS type cockpit installations and for modern computer based flight management and navigation systems To estimate the weight of the electrical systems I 73 e Airc Press Icing e Oxygen Systems e Auxiliary Power e Furnishings e Baggage Cargo e Operational ltems e Armament e Weapon Provision e Other Items e Total Fixed Equip To estimate the weight of the air conditioning pressurization and anti or de icing systems To estimate the weight of the oxygen systems To estimate the weight of the auxiliary power unit To estimate the weight of the furnishings To estimate the weight of the baggage and cargo handling equipment To estimate the weight of the operational items To estimate the weight of the armament This option is not available for General Aviation or Commercial airplanes To estimate the weight of the guns launchers and weapon provisions This option is not available for General Aviation or Commercial airplanes To estimate the weight for other not previously tabularized weight items To estimate the total fixed equipment weight 1 5 4 User Weight Estimation The user defined weights for components that are not listed in the Structure Weight Fixed Equipment Weight or Powerplant Weight modules are estimated in this module 1 5 5 Total Weight Estimation The Total Weight submodule allows the user to estimate the airplane take off weight
198. tion the drag component selection list will be displayed Each drag component selected in this menu will be used for generating the Class II drag polars I 92 Aerodynamics e Cp M C Const For generating Cp M curves for constant lift coefficient After selection of this option the drag component selection list will be displayed Each drag component selected in this menu will be used for generating the Cp M curves e Drag Build Up For generating drag polars built up of different drag components After selection of this option the drag component build up list will be displayed The selected order of the drag components in the drag build up menu will determine how the Class II drag polar will be built up e Cp a For generating Cy a curves Each drag component selected in this menu will be used for generating the Cp a curves A component can be selected by clicking on the component name or the selection box to the left of the component After the component selection the user can select either trimmed or untrimmed drag polar option The lift coefficient that is plotted is dependent on which component or combination of components is selected All selected combinations result in a Class II drag polar plot in terms of the airplane steady state lift coefficient with a few exceptions see Table 2 1 For the first three options listed above a check mark will appear in the selection box For the Drag Build Up option a number is displa
199. tion of the equivalent straight tapered canard geometry from given cranked canard tail geometry 4 6 3 Volume Coefficient For calculation of the canard Volume Coefficient 4 6 4 Canardvator Tab For calculation of the canardvator geometry 4 6 5 Chord Length For calculation of canard chord based on span wise location 4 6 6 Airfoil For calculation of airfoil parameters at the canard mean geometric chord 4 6 7 Exposed For calculation of exposed canard parameters 4 6 1 Straight Tapered The method for determining the geometry of a straight tapered canard is identical to the method found in Section 4 4 1 of this manual 4 6 2 Cranked Canard The method for determining the geometry of a cranked canard is identical to the method found in Section 4 4 2 of this manual T 118 Geometry 4 6 3 Volume Coefficient This canard submodule calculates the volume coefficient of a canard planform The methodology used to calculate the canard volume coefficient and related parameters is described in Chapter 8 of Reference 2 After calculating the output parameters the user may display the canard geometry by selecting the Plot option 4 6 4 Canardvator Tab In this submodule the geometry of the canardvator is added 4 6 5 Chord Length In this submodule the chord length of the canard at any span wise location can be determined by using root chord length taper ratio and the given span wise location 4 6 6 Airfoil In this submodule th
200. to calculate the dynamic pressure at the horizontal tail in the linear lift range The options provided in this module are discussed in Section 2 9 The purpose of the Lift submodule is to Determine whether or not a selected wing geometry gives the required value of clean airplane maximum lift coefficient Determine the type and size of high lift devices needed to meet the maximum lift requirement for the take off and landing conditions The methodology used for verifying clean airplane maximum lift coefficient and for sizing high lift devices can be found in Chapter 7 of Reference 2 The methodology used to calculate the airplane drag polars is described in References 1 and 7 After selecting the Lift submodule in the Aerodynamics main window eight new options will be displayed The options are the following e Wing e Horizontal Tail e Vertical Tail e Canard e V Tail e Flap e Nacelle Pylon For calculation of the lift coefficient of the clean wing This option 1s described in Section 2 3 1 For calculation of the lift coefficient of a horizontal tail This option is described in Section 2 3 1 For calculation of the Side Force of a vertical tail This option is described in Section 2 3 1 For calculation of the lift coefficient of a canard This option is described in Section 2 3 1 For calculation of the lift coefficient of a v tail This option is described in Section 2 3 1 For calculation of the lif
201. to estimate the pitching moment cV coefficient due to canardvator deflection derivative 6 5 6 Canardvator Tab Related Derivatives The methodology used to calculate the canardvator tab related derivatives can be found in Section 10 3 4 of Reference 7 When the Canardvator Tab button is selected a new menu appears with the following two canardvator tab related derivative options e Cis This submodule can be used to estimate the lift coefficient due to ct canardvator tab deflection derivative e Cms This submodule can be used to estimate the pitching moment ct coefficient due to canardvator tab deflection derivative 6 5 7 V Tail Related Derivatives The methodology used to calculate the v tail related derivatives is based on theory in Section 10 3 1 of Reference 7 Once the v Tail button is selected a new menu appears with the following three v tail related derivative options e Cp This submodule can be used to estimate the drag coefficient due vee to v tail incidence derivative Dynamics H 139 e C Li This submodule can be used to estimate the lift coefficient due to vee v tail incidence derivative Cin This submodule can be used to estimate the pitching moment vee coefficient due to v tail incidence derivative 6 5 8 Ruddervator Related Derivatives The methodology used to calculate the ruddervator related derivatives is based on theory in Sections 10 3 2 and 10 3 8 of Reference 7 When the Ruddervator
202. ton allows the user to switch to the next nacelle table When the user wishes to input data for each of the defined nacelles the Next Nacelle button should be used This Aerodynamics Il 97 procedure is followed until the last nacelle appears on the screen If the last nacelle is displayed selecting the Next Nacelle button will display the first nacelle table The X and Y coordinates of each nacelle must be entered in the input group for every nacelle The user may add or delete a nacelle by changing the number of nacelles in the Airplane Configuration dialog box After the input parameters and all of the input table s have been defined the user selects the Calculate button to observe the computed output parameters The value of the aerodynamic center shift due to nacelles in terms of the wing mean geometric chord will be transferred into the Cin and Cing submodule input menus 2 6 3 Aerodynamic Center Shift due to Stores When the user selects the Stores button the menu selections and procedures are identical in every respect to the aerodynamic center shift due to nacelles 2 6 4 Aerodynamic Center Shift due to Tailbooms When the user selects the Tailbooms button the menu selections and procedures are identical in every respect to the aerodynamic center shift due to nacelles 2 6 5 Aerodynamic Center Shift due to Floats When the user selects the Floats button the menu selections and procedures are identical in every respect t
203. tracted or deployed for the corresponding flight phase Gear After defining the number and location of the gear in the Gear dialog box the user can select landing gear position retracted up or extended down for the corresponding flight phase C G Location The user can specify whether the flight condition corresponds to forward or aft C G or any other C G location Ground Effects The user can indicate whether the ground effects are ON or OFF for the flight condition Flight Qualities Category The flight phase and category used in flying quality evaluation can be specified for the flight condition Notes The user may enter notes that will be saved with the flight condition The Recalculate button see Figure 2 12 on the main toolbar allows the user to recalculate Tp o m e Th 8 Class II Weight Component C G Empty Weight C G Fuel Weight C G C G Forward Aft C G Landing Gear Geometry Lift PART I i Maximum Lift j Pitching Moment k Stability and Control Derivatives 1 Class If Drag m Trimmed Drag Trend Line n Untrimmed Drag Trend Line o Drag from Trend Line p Critical Mach Number q Steady State Coefficients r Hinge Moment Derivatives s Trimmed Lift T Const t Trimmed Lift T from D u Transfer Functions v Flying Qualities w Static Margin x V n Diagram y Plot V n Diagram z Plot Trim Diagram This can be done for each flight condition separately where the us
204. tributed weight The program checks whether the distributed weight input is defined completely and logically If there are any errors the program will display a message indicating what the error is The program will check both Dist Weights options The distributed weights table is flight condition dependent and will be reset for each new flight condition The internal loads generated by each of the distributed weights selected can be calculated in the Total Internal loads module using the data contained in the distributed weights table 8 4 1 3 User Aerodynamic Loads Any aerodynamic control surface flap or user defined loads acting on structural components can be defined in the Aerodynamic module for the wing horizontal tail canard or vertical tail For the fuselage user defined loads can be defined in the User Loads module The following options are available when selected e Steady State Aerodynamic only The steady state aerodynamic loads acting on each lifting surface are defined separately as lift drag and pitching Il 176 Loads moment They can be defined in the Lift module Drag module and the Moment module Chapter 2 e Control Flap Defl Aerodynamic only Control surface and flap deflection loads can be defined as changes in the surface coefficients or as dimensional forces and moments The Coetficients module is used to define the change in coefficients of the surface created by the control surface deflection It is a
205. ubmodule in Weight Chapter 1 then they should be entered here This data will be used to calculate the internal loads due to concentrated weights in the Total Internal loads module H 174 Loads 8 4 1 2 Distributed Weights The Distributed Weights module is used to define the location and weight of masses that are distributed over the structural component The following options will be available when Dist Weights is selected 8 4 1 2 1 Add Component 8 4 1 2 2 Select Compon 8 4 1 2 3 Copy Table 8 4 1 2 4 Dist Weights 8 4 1 2 1 Add Distributed Weight Component This option is used to define weight components other than those found in the weight and balance tables of the Weight module see Chapter 1 The number of new objects is chosen and then the name weight and center of gravity positions can be entered for each mass 8 4 1 2 2 Select Distributed Weight Component This option is used to select components to be used as a distributed weight to act on the structural component The dialog box contains all of the weight items used in the weight and balance tables of the Weight module see Chapter 1 Components that are defined using the Add Component option subsection 8 4 1 1 1 can be found in the User Defined category The user can select objects by using the mouse button and then dragging down until all desired objects are highlighted When the gt gt Add button is pressed the objects will be added to the Current Selection list
206. ue to aileron derivative H 142 Dynamics 6 6 2 Aileron Tab Related Derivatives The methodology used to calculate the aileron tab related derivatives can be found in Section 10 3 5 of Reference 7 Once the Aileron Tab button is selected a new menu appears with the following three aileron tab related derivative options Sar e Cis at aj In preliminary design the side force coefficient due to aileron tab derivative is assumed to be negligible and will therefore not be calculated The program sets this derivative equal to zero and no additional menus appear on the screen This submodule can be used to estimate the rolling moment coefficient due to aileron tab derivative This submodule can be used to estimate the yawing moment coefficient due to aileron tab derivative 6 6 3 Spoiler Related Derivatives Before accessing the spoiler input output window the user must define the type of spoiler in the Airplane Configuration dialog box See Section 2 2 3 Figure 2 16 The methodology used to calculate the spoiler related derivatives can be found in Section 10 3 6 of Reference 7 When the Spoiler button is selected a new menu appears with the following three spoiler related derivative options e ee o cl Dynamics In preliminary design the side force coefficient due to spoiler derivative is assumed to be negligible and will therefore not be calculated The program sets this derivative equal to ze
207. umerous plotting options incorporated into these submodules See Part II Chapter 2 of the user s manual for a description of the Drag Polar module A detailed explanation of the plot options can be found in Part II Chapter 2 of the user s manual H 104 Performance 3 4 1 Thrust Speed Performance Curve To define a Thrust Speed curve the user needs to know three points on the thrust versus speed curve The user should specify the three known points in the input section of the menu in addition to the minimum and maximum limits of speed applicable to the curve 3 4 2 Power Speed Performance Curve To define a Power Speed curve the user needs to know three points on the power versus speed curve The user should specify the three known points in the input section of the menu in addition to the minimum and maximum limits of speed applicable to the curve 3 4 3 Stall Speeds To run the Stall Speed submodule the thrust or power versus speed performance curve must first be defined Once the Stall Speed option has been selected another menu appears showing the stall speed options e Take off For the stall speed in the take off configuration e Clean For the stall speed in a clean configuration e Landing For the stall speed in the landing configuration e Current Fit Cond For the current flight condition stall speed The Stall Speed module option works two ways 1 The user inputs the thrust power at which the stall speed is to be
208. up will be filled The selected flight phase category data are displayed in the output group Once the lateral directional Dutch roll mode checking has been performed by selecting the Plot button the user can plot the Dutch roll frequency and damping ratio requirements as specified in MIL F 8785C The current aircraft configuration Dutch roll characteristic is plotted in the same figure A description of the standard plot options can be found in Chapter 2 of Part II of this user s manual 7 4 3 Lateral Directional Stability Derivative Sensitivity Analysis The Sensitivity option provides the capability to perform a sensitivity analysis for all lateral directional derivatives weight and inertia parameters The resulting sensitivity graphs show how different modes of an airplane may vary if one of the derivatives weight or inertia parameters is varied Once the input parameters have been entered in the Transfer Function submodule the user can enter the Sensitivity submodule The following sensitivity analysis options will be displayed for the lateral directional case e Sideslip The parameters that can be varied in this submodule are Cy p Cig Cng and Curg e Roll Rate p The parameters that can be varied in this submodule are Cy Ci and Cap Il 162 Dynamics e Yaw Rate r The parameters that can be varied in this submodule are Cy C and Cp e Weight The parameter that can be varied in this submodule is Wenrr
209. urations in the Airplane Configuration dialog box See Section 2 2 3 Figure 2 16 6 9 2 Static Directional Stability The Directional submodule is used to perform all the computations related to the static directional stability The methodology used to calculate the static directional stability can be found in Section 11 2 of Reference 2 Once the Directional submodule has been selected a menu appears with three stability types The user must define the number of vertical tail surfaces in the Airplane Configuration dialog box to use these submodules These stability types are e Inherent e De Facto e Volume Method T 150 Represents inherent stability which is required of all airplanes that do not rely on a feedback augmentation system for their stability The user has two options in this submodule ox OF B for calculation of C for a given vertical tail area np O Surface Area for calculation of vertical tail area for a given C p Represents de facto stability which is required of all airplanes that are stable only with a feedback augmentation system in place If Ca B exceeds the value of 0 001 deg no feedback augmentation system is required and the feedback gain in this condition will be set to zero Represents the sizing method based on volume coefficients The user has two options in this submodule Quarter Chord For sizing the horizontal tail canard or v tail based on lifting surface position
210. ust be defined The user is required to specify the type of engine from the engine dialog box The options are as follows e Jet Engine For jet engine driven airplanes Dynamics If 125 e Piston Engine For piston engine driven airplanes e Propfan Engine For propfan engine driven airplanes e Turboprop Engine For turboprop engine driven airplanes In the case of a Jet Engine a flight condition must be specified to define the type of jet from the following options e Subsonic Jet e Supersonic Jet 5 3 Propulsion Main Window After the propulsion type has been specified a main menu is displayed This menu provides the following options e Power Extraction The options provided in this module are discussed in Section 5 4 e Inlet Design The options provided in this module are discussed in Section 5 5 e Nozzle Design The options provided in this module are discussed in Section 5 6 e Installed Data The options provided in this module are discussed in Section 5 7 5 4 Power Extraction In this submodule the user can define and calculate the quantity of the total extracted power This quantity is needed for installed thrust or installed power estimation The next menu displays the following options e Mechanical For estimation of extracted mechanical power The user should prepare a list of those systems which require mechanical power 1 e fuel pumps hydraulic pumps etc e Electrical For estimation of extracted electrical power To de
211. ut the project When S L units are in use it allows the user to define whether square millimeters or square meters will be the default area unit throughout the project e Default Speed Units This option allows the user to define the default speed unit as feet per second or knots PART I I 45 2 3 Menu Bar The main window menu bar is located at the top of the main window Most options in the Menu Bar sub menus have a corresponding option in the toolbars The menu bar contains the File Edit Window and Help menus The File menu contains nine options which are described in Table 2 9 Table 2 9 File Menu Options Create a new project Corresponds to the New icon on the File toolbar Open Open an existing project Corresponds to the Open icon on the File toolbar Save Save the current project Corresponds to the Save icon on the File toolbar Save As Save the current project under a different name or location Corresponds to the Save As icon on the File toolbar Import Import a supported file type See subsection 2 3 1 Export a supported file type See subsection 2 3 2 Print Make a printout Corresponds to the Print icon on the main toolbar Print Setup Set the default printer and printer options Corresponds to the Printer icon on the System Setup toolbar Displays the various properties of the current project Exit the program Corresponds to the Exit icon on the main toolbar The Edit menu options can be u
212. ute stability and control derivatives can be found in Reference 7 6 2 Stability amp Control Main Window After selecting the Stab amp Control module the user can select from the two options displayed e Derivatives For estimation of all longitudinal and lateral directional stability longitudinal and lateral directional control and control surface hingemoment derivatives The options provided in this module are discussed in Sections 6 3 through 6 8 e Analysis For estimation of inherent airplane stability and control surface sizing The options provided in this module are discussed in Sections 6 9 through 6 13 Once the Derivatives button has been selected in the main window a new menu appears with six options which are discussed in the following sections 6 3 Long Stability For estimation of the longitudinal stability derivatives Dynamics II 129 6 4 6 5 6 6 6 7 6 8 Lat Dir Stability Long Control Lat Dir Control Hingemoment Recalculate All For estimation of the lateral directional stability derivatives For estimation of the longitudinal control derivatives For estimation of the lateral directional control derivatives For estimation of the control surface and trim tab hingemoment derivatives For quick re calculation of all derivatives due to changes in flight condition When the Analysis button has been selected in the main window a new menu appears with three options that are discussed in the
213. utput ssie iese iape ee eri 55 Software Databases and Projects eessesessseeessseeeeseeeresrsrterrstesresrersestenresteereseerrnsesreeeesreere 57 5 1 The System Database s 2 scc s ceici sees csesssesheeieeg eseis eae oe EE E ee EEE EE EEEE EESE 57 5 2 File Compatibility with Previous Versions ooooonoccnonconnconnnonononcnancnno canon nooo n crono nnnonnnnnos 37 5 3 Projects and Their Databases ooooonncninninocononconononononcnnncnnonnonnnonnncn no cono cnn con nono ncnnncnncnnos 58 G Help System E 59 6 1 User s Manual Topics sieneen eeneioe ekkoer eeso EEEE EEs EEE Kors Eee EEE aE 59 6 2 Variable Info Topics iii 60 60 3 Theory Topics cutis 60 Table of Contents I A os 61 Weisht Mod le A ioa enra oe evi anes ota NEL I i eet AN Rate ad 61 1 1 General Description seriis tes istascceeesscessceaspescebesshsstedsnesdbesseesseodanastesessuseteteaceoet 61 12 Weight Main Win dOWemscsid stated rs dla 61 LI WES id 61 1 3 1 Mission Prove ss sssrinin saeara aristides 63 1 3 2 Take off Weight A sists dehiscteue segs e aE Ean nEs EEE nase 64 1 33 L D from Weights 2825 Hessel iaed Hee EAR a 64 LIA REESS ON e e eea e ee ee a Spe eats oe 65 IS A eee Bes UR ARASH EEEE EE EE eh S ES E 65 AI eeens Ea EEEE EEE eT E EEEE E E SEEE SE ER s 66 13 7 Remarks iunn a a a R N 66 LA gt Class Weights cnn A EAR E EE 66 1 41 AA TOO 66 142 Centeriof GraVllyaacapona da lid ia 68 1 4 2 1 Empty We
214. y be entered in the input parameter list The plot displays the current configuration along with symbols marking the points of steady state roll rate yaw divergence and pitch divergence Dynamics II 163 7 6 Defining Control Transfer Functions After selecting the Control option the user is presented with the various windows for control systems described in Sections 7 7 and 7 8 Once a control system is displayed the user may select any combination or all of the eight transfer function boxes as needed Should the user choose not to use all the available T F boxes the transfer functions in the unused boxes will default to a unity transfer function The user may execute the Control analysis module with the control loop system containing e Standard airplane transfer functions e User defined transfer functions e A combination of both standard airplane and user defined transfer functions Once a T F box is selected a list is displayed allowing the user to select the desired transfer function The transfer function options presented in this list are e Elevator Speed For the Speed to Elevator Transfer Function e Elevator Angle of Attack For the Angle of Attack to Elevator Transfer Function e Elevator Pitch Angle For the Pitch Angle to Elevator Transfer Function e Stabilizer Speed For the Speed to Stabilizer Transfer Function e Stabilizer Angle of Attack For the Angle of Attack to Stabilizer Transfer Function e Stabilizer
215. yed in the selection box indicating the order of plotting The user can deselect an option by clicking on it again Aerodynamics I 93 Table 2 1 Plot Lift Coefficients Based on Component Selected Component Selected Plot Lift Coefficient Wing only Clean Wing Lift Coefficient Horizontal Tail only Horizontal Tail Lift Coefficient Canard only Canard Lift Coefficient Flap only Clean Wing Lift Coefficient Wing and Flap only Clean Wing Lift Coefficient 2 4 3 Estimation of the Wind Tunnel Drag This module is used for the correction of wind tunnel drag on the various components The full scale zero lift drag coefficients are scaled using appropriate Reynold s number and Mach number corrections Drag polars at various control surface deflections are entered and scaled up to full scale and can be used in Class II trim analysis 2 5 Moment After selecting the Moment submodule in the Aerodynamics main window five new options will be displayed The options are the following e Wing e Horizontal Tail e Vertical Tail e Canard e V Tail e Airplane For calculation of spanwise moment distribution on the wing The user can display it by selecting the Plot button For calculation of spanwise moment distribution on the horizontal tail The user can display it by selecting the Plot button For calculation of spanwise moment distribution on the vertical tail The user can display it by selecting the Plot button For cal

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