*
*   File : C150L.dat
*
*   Purpose : This is an input file for the Digital Datcom program.
*
*   Author : Harald JOHNSEN
*
*   This file is copied to the file 'FOR005.DAT', and then the comments
*   are striped out of it. The DATCOM program is then executed, and the 
*   results are contained in the file 'FOR006.DAT'. Other output formats
*   are also generated.
*
*
*   Model-Specific notes:
*      o The CG locations were based on the empty aircraft. When we get the
*        model up, change this as required.
*
*
*   General Notes :
*      1. In general, we aren't concerned with anything faster than the 
*         subsonic speed regime. DATCOM can handle transonic, supersonic, 
*         and hypersonic, so check the book for more information.
*
*      2. DATCOM pg 47 states :
* 
*           "In general, the eight flap types defined using SYMFLP
*           (variable FTYPE) are assumed to be located on the most
*           aft lifting surface, either horizontal tail or wing if 
*           a horizontal tail is not defined."
* 
*         We are going to assume that this means that we will have to
*         perform a wing-body-vertical_tail analysis in order to get
*         the flap effects, then add the horizontal tail in order to 
*         get it's effects and trim points. The first case is used
*         ONLY to get the flap effects.
*
*      3. The page references below are from The USAF Stability and 
*         Control Digital DATCOM, volume 1, Users Manual, April 1979
*
*

************************
*   List of Command Card
************************
*
*   Note : Command cards MUST start in column 1. They may appear in
*          any order, with the exception of NEXT CASE.


*   The contents of each applicatble namelist is dumped for the case in 
*   the input system of units. (Very useful to see input parameter
*   values).
*NAMELIST

*   Preserve the input data for the case to be used in the following 
*   cases. This would be useful if multiple or comparison cases
*   are built.
*SAVE

*   Sets the system units of measure. Options are FT, IN, M, CM
DIM FT

*   Terminates the reading of input cards and begins execution of
*   the case. Case data are destroyed following execution of a case,
*   unless the SAVE card is present.
*NEXT CASE

*   Trim calculations will be performed for each subsonic Mach number 
*   within the case. A vehicle may be trimmed by deflecting a control
*   device on the wing or horizontal tail or by deflecting an all-moveable
*   horizontal stabilizer.
*TRIM

*   Provides dynamic-derivative results in addition to the standard
*   static-derivative output
DAMP

*   Provides a case identification that is printed as part of the
*   output header.
* CASEID we set it below based on the configuration

*   Prints the contents of the named arrays in the foot-pound-second
*   system. See Appendix C for list of arrays and their contents.
*
*   DUMP CASE will print all the arrays that are used during case
*             execution prior to the conventional output. (Not particularily
*             useful, as all data is in nasty arrays)
*   DUMP INPT will print dump of all input data blocks used for the case.
*             (also not useful)
*   DUMP IOM will print all the output arrays for the case. (not useful)
*   DUMP ALL will print all program arrays, even if not used for the case.
*            (surprisingly, not useful)
* DUMP ALL

*   Defines the output units of measure for the static and dynamic
*   stability derivatives, either RAD or DEG. The following parameters
*   are affected: CLa, Cma, Cyb, Cnb, Clb, CLq, Cmq, Clp, Cyp, Cnp,
*   Cnr, Clr, CLad, CMad. JSBSim XML output is also switched between 
*   degrees and radians for alpha, beta, etc.
DERIV RAD

*   Provides auxiliary and partial outputs at each Mach number in the
*   case. These outputs are automatically provided for all cases at
*   transonic Mach numbers.
*
*      Flight Conditions
*         Mach number, Altitude, Velocity, Pressure, Temperature,
*         Reynolds number
*
*      Reference dimensions
*         Reference area, Long. reference length, Lat reference length,
*         Horiz. moment ref. center, Vert. moment ref. center.
*         Also, if NACA cards are used, airfoil dimensions are output
*
*      Basic body properties
*         Wetted area, XCG, ZCG, Base area, Zero lift drag, Base drag,
*         Friction drag, Pressure drag
*
* PART

*    This command provides configuration build-up data. Conventional
*    static and dynamic stability data are output for all of the 
*    applicable basic configuration combinations shown in Table 2, 
*    page 7.  Note: This generates a lot of tables, for the body
*    alone, wing alone, ..., body-wing, body-wing-hor-tail, ...,
*    and then finally the total table, which is probably what all 
*    that we really want. This option, however, will be useful
*    when the DATCOM input data is being fine-tuned, to see the
*    effects of minor aircraft parameter changes.
*BUILD

*   Causes data generated by the program to be written to logical 
*   unit 13, which can be retained for input to the Plot Module.
*   (Looks like it dumps the data arrays out in column format. Not
*   too useful).
*PLOT





**********************  
*  Flight Conditions *
**********************

*   NMACH   Number of Mach numbers or velocities to be run, max of 20
*           Note: This parameter, along with NALT, may affect the
*           proper setting of the LOOP control parameter. 
 $FLTCON NMACH=1.0$

*   MACH    Array(20) Values of freestream Mach number   
*   VINF    Array(20) Values of freestream speed (unit: l/t)
 $FLTCON MACH(1)=0.135$



*   NALPHA  Number of angles of attack to be run, max of 20
 $FLTCON NALPHA=20.0$

*   ALSCHD  Array(20) Values of angles of attack, in ascending order
 $FLTCON ALSCHD(1)= -16.0, -8.0, -6.0, -4.0, -2.0, 0.0, 2.0,  4.0,  8.0,
      9.0, 10.0, 12.0, 14.0, 16.0, 18.0, 19.0, 20.0, 21.0, 22.0, 24.0$

*   RNNUB   Array(20) Reynolds number per unit length
*           Freestream Reynolds numbers. Each array element must 
*           correspond to the respective Mach number/freestream 
*           speed input, use LOOP=1.0
*$FLTCON RNNUB(1)=2.28E6$    ! This is calculated for us.





*   NALT    Number of atmospheric conditions to be run, max of 20
*           input as either altitude or pressure and temperature
*           Note: This parameter, along with NMACH, may affect the
*           proper setting of the LOOP control parameter. 
 $FLTCON NALT=1.0$

*   ALT     Array(20) Values of geometric altitude
*           Number of altitude and values. Note, Atmospheric conditions 
*           are input either as altitude or pressure and temperature. (MAX 20)
*   PINF    Array(20) Values of freestream Static Pressure
*   TINF    Array(20) Values of freestream Temperature
 $FLTCON ALT(1)=0.0$


*   HYPERS  =.true.  Hypersonic analysis at all Mach numbers > 1.4
*$FLTCON HYPERS=1.0$


*   STMACH  Upper limit of Mach numbers for subsonic analysis
*           (0.6<STMACH<0.99), Default to 0.6 if not input.

*   TSMACH  Lower limit of Mach number for Supersonic analysis 
*           (1.01<=TSMACH<=1.4)  Default to 1.4
 $FLTCON STMACH=0.6$
 $FLTCON TSMACH=1.4$

*   TR      Drag due to lift transition flag, for regression analysis
*           of wing-body configuration.
*           = 0.0 for no transition (default)
*           = 1.0 for transition strips or full scale flight
 $FLTCON TR=1.0$

*   WT      Vehicle Weight
 $FLTCON WT=1400.0$

*   GAMMA   Flight path angle
*$FLTCON GAMMA=$

*   LOOP    Program Looping Control
*              1 = vary altitude and mach together, default)
*              2 = vary Mach, at fixed altitude
*              3 = vary altitude, at fixed Mach
 $FLTCON LOOP=1.0$




*************************  
*  Reference Parameters *   pg 29
*************************

*   ROUGFC  Surface roughness factor, equivalent sand roughness, default
*           to 0.16e-3 inches (Natural sheet metal)
*           0.02/0.08E-3 - Polished metal or wood
*           0.16E-3  - Natural sheet metal
*           0.25E-3  - Smooth matte paint, carefully applied
*           0.40E-3  - Standard camouflage paint, average application
 $OPTINS ROUGFC=0.25E-3$


*   SREF    Reference area value of theoretical wing area used by program
*           if not input
* $OPTINS SREF=160.0$

*   CBARR   Longitudinal reference length value of theoritcal wing
*           Mean Aerodynamic Chord used by program if not input
* $OPTINS CBARR=6.75$

*   BLREF   Lateral reference length value of wing span used by program
* $OPTINS BLREF=32.0$





******************************************
* Group II     Synthesis Parameters
******************************************
*
*   page 33
*
*   XCG     Longitudinal location of cg (moment ref. center)
* I measured this to be at about 22.68, which really screwed up the 
* Cm-basic curve. Calculating it from XW and 0.26 %MAC yields 20.8.
* Sim model has 23.5 with different reference (-1.6), adjusted is 21.9.
* Keep in mind that total cg movement on Citation is about .8 feet 
 $SYNTHS XCG=5.5$ ! xnp around 5.6

*   ZCG     Vertical location of CG relative to reference plane
 $SYNTHS ZCG=0.0$ ! CHECK

*   XW      Longitudinal location of theoretical wing apex (where
*           leading edge would intersect long axis)
 $SYNTHS XW=4.5$

*   ZW      Vertical location of theoretical wing apex relative to
*           reference plane
 $SYNTHS ZW=2.0$

*   ALIW    Wing root chord incident angle measured from reference plane
 $SYNTHS ALIW=1.0$ ! CHECK 3.0 (3.75)

*   XH      Longitudinal location of theoretical horizontal tail apex. 
*           If HINAX is input, XH and ZH are evaluated at zero incidence.
 $SYNTHS XH=17.4$

*   ZH      Vertical location of theoretical horizontal tail apex
*           relative to reference plane. If HINAX is input, XH and ZH
*           are evaluated at zero incidence.
 $SYNTHS ZH=-0.3$

*   ALIH    Horizontal tail root chord incidence angle measured from
*           reference plane
 $SYNTHS ALIH=0.0$

*   XV      Longitudinal location of theoretical vertical tail apex
 $SYNTHS XV=17.8$

*   ZV      Vertical location of theoretical vertical tail apex
*           This kinda makes sense only for twin tails that are canted 
 $SYNTHS ZV=0.0$

*   XVF     Longitudinal location of theoretical ventral fin apex
*$SYNTHS XVF=20.4$
* $SYNTHS XVF=18.0$

*   ZVF     Vertical location of theoretical ventral fin apex
*           This kinda makes sense only for twin tails that are canted 
*$SYNTHS ZVF=6.2$

*   SCALE   Vehicle scale factor (multiplier to input dimensions)
 $SYNTHS SCALE=1.0$

*   VERTUP  Vertical panel above reference plane (default=true)
 $SYNTHS VERTUP=.TRUE.$

*   HINAX   Longitudinal location of horizontal tail hinge axis.
*           Required only for all-moveable horizontal tail trim option.
*$SYNTHS HINAX=17.9$



**********************************  
*  Body Configuration Parameters *
**********************************

*  page 36

*  Here is an error message output by DIGDAT concerning body geometry:

*   IN NAMELIST BODY, ONLY THE FOLLOWING COMBINATIONS OF VARIABLES CAN BE USED
*   FOR A CIRCULAR BODY, SPECIFY X AND R OR X AND S
*   FOR AN ELLIPTICAL BODY, SPECIFY X AND R OR X AND S, AND THE VARIABLE ELLIP
*   FOR OTHER BODY SHAPES X, R, S, AND P MUST ALL BE SPECIFIED

*   NX      Number of longitudinal body stations at which data is
*           specified, max of 20
 $BODY NX=13.0$

*   X       Array(20) Longitudinal distance measured from arbitray location
 $BODY X(1)=0.0,0.0,0.10,0.18,0.45,0.7,1.42,2.88,4.6,8.2,10.6,
 19.6,19.6$

*   R       Array(20) Planform half width at station Xi. See note above.
 $BODY R(1)=0.0,1.0,1.16,1.23,1.35,1.39,1.46,1.58,1.66,1.66,1.4,
 0.34,0.0$

*   ZU      Array(20) Z-coordinate at upper body surface at station Xi
*           (positive when above centerline)
*           [Only required for subsonic asymmetric bodies]
 $BODY ZU(1)=0.1,0.3,0.6,0.66,0.73,0.77,0.87,0.94,2.25,2.25,0.67,
 -0.07,-0.6$

*   ZL      Array(20) Z-coordinate at lower body surface at station Xi
*           (negative when below centerline)
*           [Only required for subsonic asymmetric bodies]
 $BODY ZL(1)=0.1,-0.5,-0.97,-1.22,-1.42,-1.57,-1.73,-1.93,-2.04,-1.9,-1.7,
 -1.04,-0.6$

*   BNOSE   Nosecone type  1.0 = conical (rounded), 2.0 = ogive (sharp point)
*           [Not required in subsonic speed regime]
* $BODY BNOSE=1.0$

*   BTAIL   Tailcone type  1.0 = conical, 2.0 = ogive, omit for lbt = 0
*           [Not required in subsonic speed regime]
* $BODY BTAIL=1.0$

*   BLN     Length of body nose
*           Not required in subsonic speed regime
* $BODY BLN=5.1$

*   BLA     Length of cylindrical afterbody segment, =0.0 for nose alone
*           or nose-tail configuration
*           Not required in subsonic speed regime
* $BODY BLA=0.0$ ! 9.8

*   DS      Nose bluntness diameter, zero for sharp nosebodies
*           [Hypersonic speed regime only]
*$BODY DS=$

*   ITYPE   1.0 = straight wing, no area rule
*           2.0 = swept wing, no area rule (default)
*           3.0 = swept wing, area rule
 $BODY ITYPE=1.0$

*   METHOD  1.0 = Use existing methods (default)
*           2.0 = Use Jorgensen method
 $BODY METHOD=1.0$






*********************************
*  Planform Variable Parameters *   pg 37-38
*********************************

*  Note : These parameters are repeated for common blocks
*          WGPLNF
*          HTPLNF
*          VTPLNF   
*          VFPLNF
*  However, not all the parameters are used in each one of the blocks.
*


**********************************
*         Wing planform variables   pg 37-38
**********************************

*   CHRDTP  Tip chord
 $WGPLNF CHRDTP=3.63$

*   SSPNOP  Semi-span outboard panel. Not required for straight 
*           tapered planform.
 $WGPLNF SSPNOP=9.30$

*   SSPNE   Semi-span exposed panel, See diagram on pg 37.
 $WGPLNF SSPNE=14.7$

*   SSPN    Semi-span theoretical panel from theoretical root chord
 $WGPLNF SSPN=16.5$

*   CHRDBP  Chord at breakpoint. Not required for straight 
*           tapered planform.
 $WGPLNF CHRDBP=5.26$

*   CHRDR   Chord root
 $WGPLNF CHRDR=5.26$

*   SAVSI   Inboard panel sweep angle
 $WGPLNF SAVSI=0.0$

*   SAVSO   Outboard panel sweep angle
 $WGPLNF SAVSO=6.0$ ! I don't understand, should be zero but then it's wrong on 3d vue

*   CHSTAT  Reference chord station for inboard and outboard panel 
*           sweep angles, fraction of chord
 $WGPLNF CHSTAT=0.25$       ! CHECK what's that

*   TWISTA  Twist angle, negative leading edge rotated down (from 
*           exposed root to tip)
 $WGPLNF TWISTA=0.0$

*   SSPNDD  Semi-span of outboard panel with dihedral
*$WGPLNF SSPNDD=$

*   DHDADI  Dihedral angle of inboard panel
 $WGPLNF DHDADI=0.0$

*   DHDADO  Dihedral angle of outboard panel. If DHDADI=DHDADO only 
*           input DHDADI
*$WGPLNF DHDADO=$

*   TYPE    1.0 - Straight tapered planform
*           2.0 - Double delta planform (aspect ratio <= 3)
*           3.0 - Cranked planform (aspect ratio > 3)
 $WGPLNF TYPE=1.0$



**********************************************
*  Wing Sectional Characteristics Parameters *
**********************************************

*  The section aerodynamic characteristics for these surfaces are
*  input using either the sectional characteristics namelists WGSCHR,
*  HTSCHR, VTSCHR and VFSCHR and/or the NACA control cards. Airfoil
*  characteristics are assummed constant for each panel of the planform.
*
*  To avoid having to input all the airfoil sectional characteristics,
*  you can specify the NACA airfoil designation.
*
*NACA x y zzzzzz
*
*   where:
*      column 1-4   NACA
*             5     any deliminator
*             6     W, H, V, or F  Planform for which the airfoil
*                                  designation applies:  Wing, Horizontal
*                                  tail, Vertical tail, or Ventral fin.
*             7     any deliminator
*             8     1,4,5,6,S      Type of airfoil section: 1-series, 
*                                  4-digit, 5-digit, 6-series, or Supersonic
*             9     any deliminator
*            10-80  Designation, columns are free format, blanks are ignored


**************************************
*       Wing Sectional Characteristics  pg 39-40
**************************************


*   Note : Only one airfoil section can be defined, so we're going to
*          with the airfoil at the root. The actual airfoil is a
*          5-series modified (23016-5), but I'm not really sure how
*          to get the -5 part in, and what it means, and where the max
*          thickness is. The tip is NACA 23012.
*        
NACA W 4 2412		! 


*   TOVC    Maximum airfoil section thickness fraction of chord
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$WGSCHR TOVC=$

*   DELTAY  Difference between airfoil ordinates at 6% and 15% chord, 
*           percent chord (% correct ???)
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$WGSCHR DELTAY=$

*   XOVC    Chord location of maximum airfoil thickness, fraction of chord
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$WGSCHR XOVC=$

*   CLI     Airfoil section design lift coefficient
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$WGSCHR CLI=$

*   ALPHAI  Angle of attack at section design lift coefficient, deg
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$WGSCHR ALPHAI=$

*   CLALPA  Airfoil section lift curve slope dCl/dAlpha, per deg (array 20)
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$WGSCHR CLALPA(1)=$

*   CLMAX   Airfoil section maximum lift cofficient (array 20)
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$WGSCHR CLMAX(1)=$

*   CMO     Section zero lift pitching moment coefficient
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$WGSCHR CMO=$

*   LERI    Airfoil leading edge radius, fraction of chord
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$WGSCHR LERI=$

*   LERO    RLE for outboard panel, fraction of chord
*           [Required input]. 
*           Not required for straight tapered planforms.
*$WGSCHR LERO=$

*   CAMBER  Cambered airfoil flag flag
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$WGSCHR CAMBER=$

*   TOVCO   t/c for outboard panel
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*           Not required for straight tapered planforms.
*$WGSCHR TOVCO=$

*   XOVCO   (x/c)max for outboard panel
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*           Not required for straight tapered planforms.
*$WGSCHR XOVCO=$

*   CMOT    Cmo for outboard panel
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*           Not required for straight tapered planforms.
*$WGSCHR CMOT=$

*   CLMAXL  Airfoil maximum lift coefficient at mach = 0.0
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$WGSCHR CLMAXL=$

*   CLAMO   Airfoil section lift curve slope at Mach=0.0, per deg
*           [Not required for subsonic speed regime. Required input 
*            for transonic speed regime, user supplied or computed if 
*            NACA card supplied]
*$WGSCHR CLAMO=$

*    TCEFF  Planform effective thickness ratio, fraction of chord
*           [Not required for subsonic speed regime. Required input 
*            for transonic speed regime, user supplied or computed if 
*            NACA card supplied]
*$WGSCHR TCEFF=$

*   KSHARP  Wave-drag factor for sharp-nosed airfoil section, not
*           input for round-nosed airfoils
*           [Not required for subsonic speed regime. Required input 
*            for transonic speed regime, user supplied or computed if 
*            NACA card supplied]
*$WGSCHR KSHARP=$

*   SLOPE   Airfoil surface slope at 0,20,40,60,80 and 100% chord, deg.
*           Positive when the tangent intersects the chord plane forward 
*           of the reference chord point
*           [Not required for subsonic speed regime. Required input 
*            for transonic speed regime, user supplied or computed if 
*            NACA card supplied]
*$WGSCHR SLOPE(1)=$

*   ARCL    Aspect ratio classification (see table 9, pg 41)
*           [Optional input]
*$WGSCHR ARCL=$

*   XAC     Section Aerodynamic Center, fraction of chord
*           [Optional input, computed by airfoil section module if airfoil
*            defined with NACA card or section coordinates]
*$WGSCHR XAC=$

*   DWASH   Subsonic downwash method flag
*           = 1.0  use DATCOM method 1
*           = 2.0  use DATCOM method 2
*           = 3.0  use DATCOM method 3
*           Supersonic, use DATCOM method 2 
*           [Optional input]
*           See figure 9 on page 41.
*$WGSCHR DWASH=$

*   YCM     Airfoil maximum camber, fraction of chord
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$WGSCHR YCM=$

*   CLD     Conical camber design lift coefficient for M=1.0 design
*           see NACA RM A55G19 (default to 0.0)
*           [Required input]
*$WGSCHR CLD=$

*   TYPEIN  Type of airfoil section coordinates input for airfoil
*           section module
*           = 1.0  upper and lower surface coordinates (YUPPER and YLOWER)
*           = 2.0 Mean line and thickness distribution (MEAN and THICK)
*           [Optional input]
*$WGSCHR TYPEIN=$

*   NPTS    Number of section points input, max = 50.0
*           [Optional input]
*$WGSCHR NPTS=$

*   XCORD   Abscissas of inputs points, TYPEIN=1.0 or 2.0, XCORD(1)=0.0
*           XCORD(NPTS)= 1.0 required.
*           [Optional input]
*$WGSCHR XCORD=$

*   YUPPER  Ordinates of upper surface, TYPEIN=1.0, fraction of chord, and
*           requires YUPPER(1)=0.0 and YUPPER(NPTS)=0.0
*           [Optional input]
*$WGSCHR YUPPER=$

*   YLOWER  Ordinates of lower surface, TYPEIN=1.0, fraction of chord,
*           and requires YLOWER(1)=0.0 and YLOWER(NPTS)=0.0
*           [Optional input]
*$WGSCHR YLOWER=$

*   MEAN    Ordinates of mean line, TYPEIN=2.0, fraction of chord, and
*           requires MEAN(1)=0.0 and MEAN(NPTS)=0.0
*           [Optional input]
*$WGSCHR MEAN=$

*   THICK   Thickness distribution, TYPEIN=2.0, fraction of chord, and
*           requires THICK(1)=0.0 and THICK(NPTS)=0.0
*           [Optional input]
*$WGSCHR THICK=$


*************************************************************************
* At this point, we want to save this case, as this is the basic fuselage
* and wing build-up. Before we can add a tail, we have to run the flaps
* and ailerons, as these always act on the aftmost horizontal surface.
*************************************************************************

*   Preserve the input data for the case to be used in the following 
*   cases. This would be useful if multiple or comparison cases
*   are built.
*SAVE





******************************
*    Ground effects parameters
******************************

*   NGH     Number of ground heights to be run, maximum of 10.
* $GRNDEF NGH=10.0$

*   GRDHT   Values of ground heights. Ground heights equal altitude
*           of reference plane relative to ground. Ground effect output
*           may be obtained at a maximum of ten different ground heights.
*           According to the DATCOM, the ground effects become neglible
*           when the ground height exceeds the wing span. Through
*           testing, there is a minimal effect up to twice the wing
*           span, so to keep our tables smooth, let's make the last
*           point 1.5b, and the output adds a point at 2b of 0.0. The
*           smallest value should NOT be 0.0, which would be the wing
*           sitting on the ground. It should be the height of the wing
*           with the aircraft sitting on the ground.
*
* $GRNDEF GRDHT=5.0,10.0,15.0,20.0,25.0,30.0,35.0,40.0,50.0,60.0$

SAVE


******************************************
*    Symetrical Flap Deflection parameters
******************************************
* DATCOM pg 47 states :
* 
*    "In general, the eight flap types defined using SYMFLP
*     (variable FTYPE) are assumed to be located on the most
*     aft lifting surface, either horizontal tail or wing if 
*     a horizontal tail is not defined."
* 
*   FTYPE   Flap type
*              1.0  Plain flaps
*              2.0  Single slotted flaps
*              3.0  Fowler flaps
*              4.0  Double slotted flaps
*              5.0  Split flaps
*              6.0  Leading edge flap
*              7.0  Leading edge slats
*              8.0  Krueger
 $SYMFLP FTYPE=2.0$ ! CHECK

*   NDELTA  Number of flap or slat deflection angles, max of 9
 $SYMFLP NDELTA=9.0$
*
*   DELTA   Flap deflection angles measured streamwise
*           (NDELTA values in array)
 $SYMFLP DELTA(1)=0.0,5.0,10.0,15.0,20.0,25.0,30.0,35.0,40.0$

*   PHETE   Tangent of airfoil trailine edge angle based on ordinates at
*           90 and 99 percent chord
 $SYMFLP PHETE=0.0522$      ! CHECK

*   PHETEP  Tangent of airfoil trailing edge angle based on ordinates at
*           95 and 99 percent chord
 $SYMFLP PHETEP=0.0391$     ! CHECK

*   CHRDFI  Flap chord at inboard end of flap, measured parallel to
*           longitudinal axis
 $SYMFLP CHRDFI=1.32$

*   CHRDFO  Flap chord at outboard end of flap, measured parallel to
*           longitudinal axis
 $SYMFLP CHRDFO=1.32$

*   SPANFI  Span location of inboard end of flap, measured perpendicular
*           to vertical plane of symmetry
 $SYMFLP SPANFI=1.70$

*   SPANFO  Span location of outboard end of flap, measured perpendicular
*           to vertical plane of symmetry
 $SYMFLP SPANFO=6.75$

*   CPRMEI  Total wing chord at inboard end of flap (translating devices 
*           only) measured parallel to longitudinal axis
*           (NDELTA values in array)
*              Single-slotted, Fowler, Double-slotted, leading-edge
*              slats, Krueger flap, jet flap
* 
*   BIG GUESS HERE, as I don't know the characteristics of how the
*   flaps move. Guess at max of 4 inches gap at full deflection.
 $SYMFLP CPRMEI(1)=8.1,8.1,8.2,8.2,8.3,8.3,8.3,8.4,8.4$

*   CPRMEO  Total wing chord at outboard end of flap (translating devices
*           only) measured parallel to longitudinal axis
*           (NDELTA values in array)
*              Single-slotted, Fowler, Double-slotted, leading-edge
*              slats, Krueger flap, jet flap
* 
*   BIG GUESS HERE, as I don't know the characteristics of how the
*   flaps move. Guess at max of 4 inches gap at full deflection.
 $SYMFLP CPRMEO(1)=3.7,3.7,3.8,3.8,3.9,3.9,3.9,4.0,4.0$

*   CAPINS                           (double-slotted flaps only)
*           (NDELTA values in array)
*$SYMFLP CAPINS$

*   CAPOUT                           (double-slotted flaps only)
*           (NDELTA values in array)
*$SYMFLP CAPOUT$

*   DOSDEF                           (double-slotted flaps only)
*           (NDELTA values in array)
*$SYMFLP DOSDEF$

*   DOBCIN                           (double-slotted flaps only)
*$SYMFLP DOBCIN

*   DOBCOT                           (double-slotted flaps only)
*$SYMFLP DOBCOT$

*   SCLD    Increment in section lift coefficient due to
*           deflecting flap to angle DELTA[i]      (optional)
*           (NDELTA values in array)
*$SYMFLP SCLD$

*   SCMD    Increment in section pitching moment coefficient due to
*           deflecting flap to angle DELTA[i]      (optional)
*           (NDELTA values in array)
*$SYMFLP SCMD$

*   CB      Average chord of the balance    (plain flaps only)
*$SYMFLP CB=$

*   TC      Average thickness of the control at hinge line
*                                           (plain flaps only)
*$SYMFLP TC=0.3$

*   NTYPE   Type of nose
*              1.0  Round nose flap
*              2.0  Elliptic nose flap
*              3.0  Sharp nose flap
 $SYMFLP NTYPE=1.0$                     ! Guess

*   JETFLP  Type of flap
*              1.0  Pure jet flap
*              2.0  IBF
*              3.0  EBF
*$SYMFLP JETFLP$

*   CMU     Two-dimensional jet efflux coefficient
*$SYMFLP CMU$

*   DELJET  Jet deflection angle
*           (NDELTA values in array)
*$SYMFLP DELJET$

*   EFFJET  EBF Effective jet deflection angle
*           (NDELTA values in array)
*$SYMFLP EFFJET$




**********************************************************************
* At this point, we are going to terminate the case so that we can get 
* the flap effects. We can't save this data, as we are
* also going to do control surfaces on the horizontal tail.
**********************************************************************

*   Terminates the reading of input cards and begins execution of
*   the case.
CASEID FLAPS: Cessna 150 L Aircraft
NEXT CASE



*************************************************************
*    Asymmetrical Control Deflection parameters : Ailerons
*************************************************************

*   STYPE   Type
*              1.0  Flap spoiler on wing
*              2.0  Plug spoiler on wing
*              3.0  Spoiler-slot-deflection on wing
*              4.0  Plain flap aileron
*              5.0  Differentially deflected all moveable horizontal tail
 $ASYFLP STYPE=4.0$

*   NDELTA  Number of control deflection angles, required for all controls,
*           max of 9
 $ASYFLP NDELTA=9.0$

*   DELTAL  Defelction angle for left hand plain flap aileron or left
*           hand panel all moveable horizontal tail, measured in
*           vertical plane of symmetry
 $ASYFLP DELTAL(1)=-21.0,-15.0,-10.0,-5.0,0.0,5.0,10.0,15.0,21.0$

*   DELTAR  Defelction angle for right hand plain flap aileron or right
*           hand panel all moveable horizontal tail, measured in
*           vertical plane of symmetry
 $ASYFLP DELTAR(1)=21.0,15.0,10.0,5.0,0.0,-5.0,-10.0,-15.0,-21.0$

*   SPANFI  Span location of inboard end of flap or spoiler control,
*           measured perpendicular to vertical plane of symmetry
 $ASYFLP SPANFI=6.75$

*   SPANFO  Span location of outboard end of flap or spoiler control,
*           measured perpendicular to vertical plane of symmetry
 $ASYFLP SPANFO=15.65$

*   PHETE   Tangent of airfoil trailing edge angle based on ordinates
*           at x/c - 0.90 and 0.99
 $ASYFLP PHETE=0.00$      ! CHECK 

*   CHRDFI  Aileron chord at inboard end of plain flap aileron,
*           measured parallel to longitudinal axis
 $ASYFLP CHRDFI=0.786$

*   CHRDFO  Aileron chord at outboard end of plain flap aileron,
*           measured parallel to longitudinal axis
 $ASYFLP CHRDFO=0.786$

*   DELTAD  Projected height of deflector, spoiler-slot-deflector
*           control, fraction of chord
*$ASYFLP DELTAD(1) $

*   DELTAS  Projected height of spoiler, flap spoiler, plug spoiler and
*           spoiler-slot-deflector control; fraction of chord
*$ASYFLP DELTAS(1) $

*   XSOC    Distance from wing leading edge to spoiler lip measured
*           parallel to streamwise wng chord, flap and plug spoilers,
*           fraction of chord
*$ASYFLP XSOC(1) $

*   XSPRME  Distance from wing leading edge to spoiler hinge line
*           measured parallel to streamwise chord, flap spoiler,
*           plug spoiler and spoiler-slot-deflector control, fraction 
*           of chord
*$ASYFLP XSPRME $

*   HSOC    Projected height of spoiler measured from and normal to
*           airfoil mean line, flap spoiler, plug spoiler and spoiler-
*           slot-reflector, fraction of chord
*$ASYFLP HSOC(1) $


**********************************************************************
* At this point, we are going to terminate the case so that we can get 
* the aileron effects.
**********************************************************************

*   Terminates the reading of input cards and begins execution of
*   the case. Case data are destroyed following execution of a case,
*   unless the SAVE card is present.
CASEID AILERONS: Cessna 150 L Aircraft
SAVE
NEXT CASE




*********************************************
*         Horizontal Tail planform variables   pg 37-38
*********************************************

*   CHRDTP  Tip chord
 $HTPLNF CHRDTP=2.13$

*   SSPNOP  Semi-span outboard panel. Not required for straight 
*           tapered planform.
*$HTPLNF SSPNOP=$

*   SSPNE   Semi-span exposed panel
 $HTPLNF SSPNE=3.6$

*   SSPN    Semi-span theoretical panel from theoretical root chord
 $HTPLNF SSPN=3.9$

*   CHRDBP  Chord at breakpoint
*$HTPLNF CHRDBP=$

*   CHRDR   Chord root
 $HTPLNF CHRDR=4.00$

*   SAVSI   Inboard panel sweep angle
 $HTPLNF SAVSI=0.0$ ! CHECK

*   CHSTAT  Reference chord station for inboard and outboard panel 
*           sweep angles, fraction of chord
 $HTPLNF CHSTAT=0.25$ ! CHECK

*   TWISTA  Twist angle, negative leading edge rotated down (from 
*           exposed root to tip)
 $HTPLNF TWISTA=-3.0$		! Guess

*   SSPNDD  Semi-span of outboard panel with dihedral
*$HTPLNF SSPNDD=$

*   DHDADI  Dihedral angle of inboard panel
 $HTPLNF DHDADI=0.0$		! Average value used

*   DHDADO  Dihedral angle of outboard panel. If DHDADI=DHDADO only 
*           input DHDADI
*$HTPLNF DHDADO=$

*   TYPE    1.0 - Straight tapered planform
*           2.0 - Double delta planform (aspect ratio <= 3)
*           3.0 - Cranked planform (aspect ratio > 3)
 $HTPLNF TYPE=1.0$

*   SHB     Portion of fuselage side area that lies between Mach lines 
*           originating from leading and trailing edges of horizontal 
*           tail exposed root chord (array 20).
*           Only required for supersonic and hypersonic speed regimes.
*$HTPLNF SHB(1)=$

*   SEXT     Portion of extended fueslage side area that lies between 
*            Mach lines originating from leading and trailing edges of 
*            horizontal tail exposed root chord (array 20)
*           Only required for supersonic and hypersonic speed regimes.
*$HTPLNF SEXT(1)=$

*   RLPH    Longitudinal distance between CG and centroid of Sh(B) 
*           positive aft of CG
*           Only required for supersonic and hypersonic speed regimes.
*$HTPLNF RLPH(1)=$


NACA H 4 0012	! Citation is 0010 at root, 0008 at tip



******************************************
*         Vertical Tail planform variables   pg 37-38
******************************************

*   CHRDTP  Tip chord
 $VTPLNF CHRDTP=2.13$

*   SSPNOP  Semi-span outboard panel
*$VTPLNF SSPNOP=$

*   SSPNE   Semi-span exposed panel
 $VTPLNF SSPNE=3.6$

*   SSPN    Semi-span theoretical panel from theoretical root chord
 $VTPLNF SSPN=3.9$

*   CHRDBP  Chord at breakpoint
*$VTPLNF CHRDBP=1.16$

*   CHRDR   Chord root
 $VTPLNF CHRDR=4.00$

*   SAVSI   Inboard panel sweep angle
 $VTPLNF SAVSI=5.0$ ! CHECK

*   SAVSO   Outboard panel sweep angle
* $VTPLNF SAVSO=$

*   CHSTAT  Reference chord station for inboard and outboard panel 
*           sweep angles, fraction of chord
 $VTPLNF CHSTAT=0.0$

*   TYPE    1.0 - Straight tapered planform
*           2.0 - Double delta planform (aspect ratio <= 3)
*           3.0 - Cranked planform (aspect ratio > 3)
 $VTPLNF TYPE=1.0$

*   SVWB    Portion of exposed vertical panel area that lies between 
*           Mach lines emanating from leading and trailing edges of 
*           wing exposed root chord (array 20)
*           Only required for supersonic and hypersonic speed regimes.
*$VTPLNF SVWB(1)=$

*   SVB     Area of exposed vertical panel not influenced by wing or 
*           horizontal tail (array 20)
*           Only required for supersonic and hypersonic speed regimes.
*$VTPLNF SVB(1)=$

*   SVHB    Portion of exposed vertical panel area that lies between Mach 
*           lines emanating from leading and and trailing edges of 
*           horizontal tail exposed root chord (array 20)
*           Only required for supersonic and hypersonic speed regimes.
*$VTPLNF SVHB(1)=$


NACA V 4 0012	! Citation is 0012 at root, 0008 at tip



******************************************
*         Vertical Fin planform variables     pg 37-38
******************************************

*   CHRDTP  Tip chord
* $VFPLNF CHRDTP=0.0$    ! Will this value work?

*   SSPNOP  Semi-span outboard panel
*$VFPLNF SSPNOP=$

*   SSPNE   Semi-span exposed panel
* $VFPLNF SSPNE=1.9$

*   SSPN    Semi-span theoretical panel from theoretical root chord
* $VFPLNF SSPN=3.4$

*   CHRDBP  Chord at breakpoint
*$VFPLNF CHRDBP=$

*   CHRDR   Chord root
* $VFPLNF CHRDR=13.0$

*   SAVSI   Inboard panel sweep angle
* $VFPLNF SAVSI=80.0$

*   CHSTAT  Reference chord station for inboard and outboard panel 
*           sweep angles, fraction of chord
* $VFPLNF CHSTAT=0.0$

*   DHDADO  Dihedral angle of outboard panel. If DHDADI=DHDADO only 
*           input DHDADI
* $VFPLNF DHDADO=0.0$

*   DHDADO  Dihedral angle of outboard panel. If DHDADI=DHDADO only 
*           input DHDADI
*$VFPLNF DHDADO=$

*   TYPE    1.0 - Straight tapered planform
*           2.0 - Double delta planform (aspect ratio <= 3)
*           3.0 - Cranked planform (aspect ratio > 3)
* $VFPLNF TYPE=1.0$

*   SVWB    Portion of exposed vertical panel area that lies between 
*           Mach lines emanating from leading and trailing edges of 
*           wing exposed root chord (array 20)
*           Only required for supersonic and hypersonic speed regimes.
*$VFPLNF SVWB(1)=$

*   SVB     Area of exposed vertical panel not influenced by wing or 
*           horizontal tail (array 20)
*           Only required for supersonic and hypersonic speed regimes.
*$VFPLNF SVB(1)=$

*   SVHB    Portion of exposed vertical panel area that lies between Mach 
*           lines emanating from leading and and trailing edges of 
*           horizontal tail exposed root chord (array 20)
*           Only required for supersonic and hypersonic speed regimes.
*$VFPLNF SVHB(1)=$




*************************************************
*       Horizontal Tail Sectional Characteristics  pg 39-40
*************************************************

*   Note : a lot of tails are 009 airfoils. It's probably a good guess
*          unless you know otherwise.
*

*   TOVC    Maximum airfoil section thickness fraction of chord
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$HTSCHR TOVC=$

*   DELTAY  Difference between airfoil ordinates at 6% and 15% chord, 
*           percent chord (% correct ???)
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$HTSCHR DELTAY=$

*   XOVC    Chord location of maximum airfoil thickness, fraction of chord
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$HTSCHR XOVC=$

*   CLI     Airfoil section design lift coefficient
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$HTSCHR CLI=$

*   ALPHAI  Angle of attack at section design lift coefficient, deg
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$HTSCHR ALPHAI=$

*   CLALPA  Airfoil section lift curve slope dCl/dAlpha, per deg (array 20)
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$HTSCHR CLALPA(1)=$

*   CLMAX   Airfoil section maximum lift cofficient (array 20)
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$HTSCHR CLMAX(1)=$

*   CMO     Section zero lift pitching moment coefficient
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$HTSCHR CMO=$

*   LERI    Airfoil leading edge radius, fraction of chord
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$HTSCHR LERI=$

*   LERO    RLE for outboard panel, fraction of chord
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$HTSCHR LERO=$

*   CAMBER  Cambered airfoil flag flag
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$HTSCHR CAMBER=$

*   TOVCO   t/c for outboard panel
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$HTSCHR TOVCO=$

*   XOVCO   (x/c)max for outboard panel
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$HTSCHR XOVCO=$

*   CMOT    Cmo for outboard panel
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$HTSCHR CMOT=$

*   CLAMO   Airfoil section lift curve slope at Mach=0.0, per deg
*           [Not required for subsonic speed regime. Required input 
*            for transonic speed regime, user supplied or computed if 
*            NACA card supplied]
*$HTSCHR CLAMO=$

*    TCEFF  Planform effective thickness ratio, fraction of chord
*           [Not required for subsonic speed regime. Required input 
*            for transonic speed regime, user supplied or computed if 
*            NACA card supplied]
*$HTSCHR TCEFF=$

*   KSHARP  Wave-drag factor for sharp-nosed airfoil section, not
*           input for round-nosed airfoils
*           [Not required for subsonic speed regime. Required input 
*            for transonic speed regime, user supplied or computed if 
*            NACA card supplied]
*$HTSCHR KSHARP=$

*   ARCL    Aspect ratio classification (see table 9, pg 41)
*           [Optional input]
*$HTSCHR ARCL=$

*   XAC     Section Aerodynamic Center, fraction of chord
*           [Optional input, computed by airfoil section module if airfoil
*            defined with NACA card or section coordinates]
*$HTSCHR XAC=$

*   YCM     Airfoil maximum camber, fraction of chord
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$HTSCHR YCM=$

*   CLD     Conical camber design lift coefficient for M=1.0 design
*           see NACA RM A55G19 (default to 0.0)
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$HTSCHR CLD=$

*   TYPEIN  Type of airfoil section coordinates input for airfoil
*           section module
*           = 1.0  upper and lower surface coordinates (YUPPER and YLOWER)
*           = 2.0 Mean line and thickness distribution (MEAN and THICK)
*           [Optional input]
*$HTSCHR TYPEIN=$

*   NPTS    Number of section points input, max = 50.0
*           [Optional input]
*$HTSCHR NPTS=$

*   XCORD   Abscissas of inputs points, TYPEIN=1.0 or 2.0, XCORD(1)=0.0
*           XCORD(NPTS)= 1.0 required.
*           [Optional input]
*$HTSCHR XCORD=$

*   YUPPER  Ordinates of upper surface, TYPEIN=1.0, fraction of chord, and
*           requires YUPPER(1)=0.0 and YUPPER(NPTS)=0.0
*           [Optional input]
*$HTSCHR YUPPER=$

*   YLOWER  Ordinates of lower surface, TYPEIN=1.0, fraction of chord,
*           and requires YLOWER(1)=0.0 and YLOWER(NPTS)=0.0
*           [Optional input]
*$HTSCHR YLOWER=$

*   MEAN    Ordinates of mean line, TYPEIN=2.0, fraction of chord, and
*           requires MEAN(1)=0.0 and MEAN(NPTS)=0.0
*           [Optional input]
*$HTSCHR MEAN=$

*   THICK   Thickness distribution, TYPEIN=2.0, fraction of chord, and
*           requires THICK(1)=0.0 and THICK(NPTS)=0.0
*           [Optional input]
*$HTSCHR THICK=$






***********************************
*    Elevator Deflection parameters
***********************************
* DATCOM pg 47 states :
* 
*    "In general, the eight flap types defined using SYMFLP
*     (variable FTYPE) are assumed to be located on the most
*     aft lifting surface, either horizontal tail or wing if 
*     a horizontal tail is not defined."
* 
* So, does this mean that a flap cannot be added to a 
* wing-body-tail configuration??? Additionally, does this mean
* that the elvator is treated as a flap?

*   FTYPE   Flap type
*              1.0  Plain flaps
*              2.0  Single slotted flaps
*              3.0  Fowler flaps
*              4.0  Double slotted flaps
*              5.0  Split flaps
*              6.0  Leading edge flap
*              7.0  Leading edge slats
*              8.0  Krueger
 $SYMFLP FTYPE=1.0$

*   NDELTA  Number of flap or slat deflection angles, max of 9
 $SYMFLP NDELTA=9.0$
*
*   DELTA   Flap deflection angles measured streamwise
*           (NDELTA values in array)
 $SYMFLP DELTA(1)=-26.0,-17.0,-10.0,-5.0,0.0,5.0,9.0,13.0,17.0$

*   PHETE   Tangent of airfoil trailine edge angle based on ordinates at
*           90 and 99 percent chord
 $SYMFLP PHETE=0.0522$      ! value in SPROB.IN for plain flaps

*   PHETEP  Tangent of airfoil trailing edge angle based on ordinates at
*           95 and 99 percent chord
 $SYMFLP PHETEP=0.0523$     ! value in SPROB.IN for plain flaps

*   CHRDFI  Flap chord at inboard end of flap, measured parallel to
*           longitudinal axis
 $SYMFLP CHRDFI=1.50$      ! scaled from three-view drawing

*   CHRDFO  Flap chord at outboard end of flap, measured parallel to
*           longitudinal axis
 $SYMFLP CHRDFO=1.02$     ! scaled from three-view drawing

*   SPANFI  Span location of inboard end of flap, measured perpendicular
*           to vertical plane of symmetry
 $SYMFLP SPANFI=0.8$       ! scaled from three-view drawing

*   SPANFO  Span location of outboard end of flap, measured perpendicular
*           to vertical plane of symmetry
 $SYMFLP SPANFO=4.6$       ! scaled from three-view drawing

*   CPRMEI  Total wing chord at inboard end of flap (translating devices 
*           only) measured parallel to longitudinal axis
*           (NDELTA values in array)
*              Single-slotted, Fowler, Double-slotted, leading-edge
*              slats, Krueger flap, jet flap
* 
*$SYMFLP CPRMEI(1)=$

*   CPRMEO  Total wing chord at outboard end of flap (translating devices
*           only) measured parallel to longitudinal axis
*           (NDELTA values in array)
*              Single-slotted, Fowler, Double-slotted, leading-edge
*              slats, Krueger flap, jet flap
* 
*$SYMFLP CPRMEO(1)=$

*   CAPINS                           (double-slotted flaps only)
*           (NDELTA values in array)
*$SYMFLP CAPINS$

*   CAPOUT                           (double-slotted flaps only)
*           (NDELTA values in array)
*$SYMFLP CAPOUT$

*   DOSDEF                           (double-slotted flaps only)
*           (NDELTA values in array)
*$SYMFLP DOSDEF$

*   DOBCIN                           (double-slotted flaps only)
*$SYMFLP DOBCIN

*   DOBCOT                           (double-slotted flaps only)
*$SYMFLP DOBCOT$

*   SCLD    Increment in section lift coefficient due to
*           deflecting flap to angle DELTA[i]      (optional)
*           (NDELTA values in array)
*$SYMFLP SCLD$

*   SCMD    Increment in section pitching moment coefficient due to
*           deflecting flap to angle DELTA[i]      (optional)
*           (NDELTA values in array)
*$SYMFLP SCMD$

*   CB      Average chord of the balance    (plain flaps only)
 $SYMFLP CB=0.84$ ! CHECK

*   TC      Average thickness of the control at hinge line
*                                           (plain flaps only)
 $SYMFLP TC=0.3$          ! CHECK

*   NTYPE   Type of nose
*              1.0  Round nose flap
*              2.0  Elliptic nose flap
*              3.0  Sharp nose flap
 $SYMFLP NTYPE=1.0$

*   JETFLP  Type of flap
*              1.0  Pure jet flap
*              2.0  IBF
*              3.0  EBF
*$SYMFLP JETFLP$

*   CMU     Two-dimensional jet efflux coefficient
*$SYMFLP CMU$

*   DELJET  Jet deflection angle
*           (NDELTA values in array)
*$SYMFLP DELJET$

*   EFFJET  EBF Effective jet deflection angle
*           (NDELTA values in array)
*$SYMFLP EFFJET$






**************************************
*       Vertical Tail Sectional Characteristics  pg 39-40
**************************************

*   Note : a lot of tails are 009 airfoils. It's probably a good guess
*          unless you know otherwise.
*

*   TOVC    Maximum airfoil section thickness fraction of chord
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$VTSCHR TOVC=$

*   XOVC    Chord location of maximum airfoil thickness, fraction of chord
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$VTSCHR XOVC=$

*   CLALPA  Airfoil section lift curve slope dCl/dAlpha, per deg (array 20)
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$VTSCHR CLALPA(1)=$

*   LERI    Airfoil leading edge radius, fraction of chord
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$VTSCHR LERI=$

*   LERO    RLE for outboard panel, fraction of chord
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$VTSCHR LERO=$

*   TOVCO   t/c for outboard panel
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$VTSCHR TOVCO=$

*   XOVCO   (x/c)max for outboard panel
*           [Required input, user supplied or computed by airfoil
*            section module if airfoil defined with NACA card or
*            section coordinates]
*$VTSCHR XOVCO=$

*    TCEFF  Planform effective thickness ratio, fraction of chord
*           [Not required for subsonic speed regime. Required input 
*            for transonic speed regime, user supplied or computed if 
*            NACA card supplied]
*$VTSCHR TCEFF=$

*   KSHARP  Wave-drag factor for sharp-nosed airfoil section, not
*           input for round-nosed airfoils
*           [Not required for subsonic speed regime. Required input 
*            for transonic speed regime, user supplied or computed if 
*            NACA card supplied]
*$VTSCHR KSHARP=$

*   ARCL    Aspect ratio classification (see table 9, pg 41)
*           [Optional input]
*$VTSCHR ARCL=$

*   TYPEIN  Type of airfoil section coordinates input for airfoil
*           section module
*           = 1.0  upper and lower surface coordinates (YUPPER and YLOWER)
*           = 2.0 Mean line and thickness distribution (MEAN and THICK)
*           [Optional input]
*$VTSCHR TYPEIN=$

*   NPTS    Number of section points input, max = 50.0
*           [Optional input]
*$VTSCHR NPTS=$

*   XCORD   Abscissas of inputs points, TYPEIN=1.0 or 2.0, XCORD(1)=0.0
*           XCORD(NPTS)= 1.0 required.
*           [Optional input]
*$VTSCHR XCORD=$

*   YUPPER  Ordinates of upper surface, TYPEIN=1.0, fraction of chord, and
*           requires YUPPER(1)=0.0 and YUPPER(NPTS)=0.0
*           [Optional input]
*$VTSCHR YUPPER=$

*   YLOWER  Ordinates of lower surface, TYPEIN=1.0, fraction of chord,
*           and requires YLOWER(1)=0.0 and YLOWER(NPTS)=0.0
*           [Optional input]
*$VTSCHR YLOWER=$

*   MEAN    Ordinates of mean line, TYPEIN=2.0, fraction of chord, and
*           requires MEAN(1)=0.0 and MEAN(NPTS)=0.0
*           [Optional input]
*$VTSCHR MEAN=$

*   THICK   Thickness distribution, TYPEIN=2.0, fraction of chord, and
*           requires THICK(1)=0.0 and THICK(NPTS)=0.0
*           [Optional input]
*$VTSCHR THICK=$


*******
*  Experimental data input block comes next, but I decided not to
*  type all of this stuff in at this point. It is all defined on page 45.
*******




*****************************************************
*   Propulsion parameters for Propeller Power Effects
*****************************************************

*   AIETLP  Angle of incidence of engine thrust axis, deg
* $PROPWR AIETLP=0.0$

*   NENGSP  Number of engines (1 or 2 only)
* $PROPWR NENGSP=1.0$

*   THSTCP  Thrust coefficient 2T/PV^2 Sref

*   PHALOC  Axial location of propeller hub
* $PROPWR PHALOC=0.0$

*   PHVLOC  Vertical location of propeller hub
* $PROPWR PHVLOC=0.0$

*   PRPRAD  Propeller radius
* $PROPWR PRPRAD=5.75$

*   ENGFCT  Empirical normal force factor
*           Not required if blade widths are input.

*   BWAPR3  Blade width at 0.3 propeller radius
*           Not required if empirical normal force factor is input.
*$PROPWR BWAPR3$

*   BWAPR6  Blade width at 0.6 propeller radius
*           Not required if empirical normal force factor is input.
*$PROPWR BWAPR6$

*   BWAPR9  Blade width at 0.9 propeller radius
*           Not required if empirical normal force factor is input.
*$PROPWR BWAPR9$

*   NOPBPE  Number of propeller blades per engine
* $PROPWR NOPBPE=2.0$

*   BAPR75  Blade angle at 0.75 propeller radius
*$PROPWR BAPR75$       !  18.0 ?     ! Do I have to input this?

*   YP      Lateral location of engine
* $PROPWR YP=0.0$

*   CROT    .true.  Counter rotation propeller, 
*           .false. Non counter rotating
*                       Is this clockwise/counterclockwise?




*********************************
*    Jet Power Effects parameters  pg 51
*********************************
*************************************************
*     Low-Aspect Ratio Wing, Wing-Body parameters
*************************************************
***************************************
*     Transverse-Jet Control parameters
***************************************
**************************************************
*     Flap Control at Hypersonic Speeds parameters
**************************************************
*****************************
*     Control Tabs parameters
*****************************
* DATCOM User's manual, pg 68/69


*   TTYPE   Type
*              1.0  Tab Control
*              2.0  Trim Tab
*              3.0  Both
*$CONTAB TTYPE$

*   CFITC   Control tab inboard chord
*$CONTAB CFITC$

*   CFOTC   Control tab outboard chord
*$CONTAB CFOTC$

*   BITC    Control tab inboard span location
*$CONTAB BITC$

*   BOTC    Control tab outboard span location
*$CONTAB BOTC$

*   CFITT   Trim tab inboard chord
*$CONTAB CFITT$

*   CFOTT   Trim tab outboard chord
*$CONTAB CFOTT$

*   BITT    Trim tab inboard span location
*$CONTAB BITT$

*   BOTT    Trim tab outboard span location
*$CONTAB BOTT$

*   B1      Ch(delta) for control tab
*$CONTAB B1$

*   B2      Ch(delta) for trim tab
*$CONTAB B2$

*   B3      Ch(alpha) for control tab
*$CONTAB B3$

*   B4      Ch(alpha) for trim tab
*$CONTAB B4$

*   D1      see table 11, pg 70-71
*$CONTAB D1$

*   D2      see table 11, pg 70-71
*$CONTAB D2$

*   D3      see table 11, pg 70-71
*$CONTAB D3$

*   GCMAX   Maximum Stick gearing
*$CONTAB GCMAX$

*   KS      Tab spring effectiveness
*$CONTAB KS$

*   RL      Aerodynamic boost link ratio (RL >  0). 
*           To input RL  infinity, set RL < 0
*$CONTAB RL$

*   BGR     Control tab gear ratio
*$CONTAB BGR$

*   DELR    -DELTAtc(max)/DELTAc(max) for a maximum control deflection
*           The value of DELR is positive because DELTAtc(max) and
*           DELTAc(max) will have opposite signs.
*           When R(L)   0, DELR = 1.0
*$CONTAB DELR$

* DUMP ALL
* DAMP


CASEID TOTAL: Cessna 150 L Aircraft
* NEXT CASE

*   End of File
