*
*  Data generated by DATCOM for 
*  TOTAL: FIAT G91R4                                                         
*
*      $Revision: 1.3 $
*
*
*     Table definitions
*     -----------------
* CL_Flaps_Table - Incremental lift coefficent in the linear-lift 
*    angle-of-attack range due to deflection of flap surface, based on
*    wing area.
* Cm_Flaps_Table - Incremental pitching-moment coefficient due to flap
*    deflection valid in the linear lift angle-of-attack range, based
*    on wing area and mean aerodynamic chord.
* Cd_Flaps_Table - Incremental induced-drag coefficient due to flap
*    deflection valid in the linear lift angle-of-attack range, based
*    on wing area.
* Cn_Aileron_Table - Incremental yawing-moment coefficient due to
*    asymmetrical deflection of ailerons, based on wing area and wing
*    span. Positive yawing moment is nose right.
* Cl_Aileron_Table - Incremental rolling-moment coefficient due to
*    asymmetrical deflection of control surface based on wing area
*    and wing span. Positive rolling moment is right wing down.
* Cd_Basic_Table - Vehicle drag coefficient, based on wing area. 
*    Positive aft.
* CL_Basic_Table - Vehicle lift coefficient, based on wing area. 
*    Positive up.
* Cm_Basic_Table - Vehicle pitching-moment coefficient, based on wing 
*    area and mean aerodynamic chord. Positive causes nose-up rotation.
* CN_Basic_Table - (body axis) Vehicle normal-force coefficient, based 
*    on wing area. Positive up through body.
* CA_Basic_Table - (body-axis) Vehicle axial-force coefficient, based 
*    on wing area. Positive aft.
* CLa_Basic_Table - Derivative of lift coefficient w.r.t alpha, numerical
*    derivatives of the lift curve, based on the wing area.
* Cma_Basic_Table - Derivative of pitching-moment coefficient w.r.t alpha,
*    derivatives of the lift curve, based on the wing area and mean 
*    aerodynamic chord.
* Cyb_Basic_Table - Derivative of side-force coefficient w.r.t. sideslip 
*    angle, based on wing area.
* Cnb_Basic_Table - Derivative of yawing-moment coefficient w.r.t. 
*    sideslip angle, based on wing area and wing span.
* Clb_Basic_Table - Derivative of rolling-moment coeficient w.r.t. 
*    sideslip angle, based on wing area and wing span.
* 
* Q_Qinf_Table - Ratio of dynamic pressure at the horiz. tail
*    to freestream dynamic pressure.
* Epslon_Table - Downwash angle at horizonal tail expressed in degrees.
*    Positive implies local AOA less than free-stream AOA.
* d_Epslon_Table - d(EPSLON)/d(ALPHA) - Derivative of downwash angle 
*    w.r.t AOA
* Cd_Elevator_Table - Incremental induced drag coefficient due to 
*    elevator deflection, based on wing area.
* CL_Elevator_Table - Incremental lift coefficient in the linear-lift
*    angle-of-attack range due to elevator deflection, based on wing 
*    area.
* Cm_Elevator_Table - Incremental pitching-moment coefficient due to 
*    elevator deflection valid in the linear-lift angle-of-attack range, 
*    based on wing area and mean aerodynamic chord.
* Cha_Elevator_Table - Elevator-surface hinge-moment derivative due to 
*    angle of attack based on elevator area and elevator chord. Positive 
*    hinge moment will tend to rotate the surface trailing edge down.
* Chd_Elevator_Table - Elevator-surface hinge-moment derivative due to 
*    elevator deflection based on elevator area and elevator chord. 
*    Positive hinge moment will tend to rotate the surface trailing 
*    edge down.
* Cd_Power_Table - Incremental drag coefficient due to power
*    effects, based on wing area.
* CL_Power_Table - Incremental lift coefficient due to power
*    effects, based on wing area.
* Cm_Power_Table - Incremental pitching moment coefficient due to power
*    effects, based on wing area.
* 
*    Dynamic Derivatives
*    -------------------
* CLq_Basic_Table - lift derivative ( dCL/d(q*cbar/2Vin) ), based on
*    wing area and mean aerodynamic chord.
* Cmq_Basic_Table - pitching moment derivative ( dCm/d(q*cbar/2Vin) ), 
*    based on wing area and mean aerodynamic chord squared.
* CLad_Basic_Table - CL alpha dot  ( dCL/d(alpha_dot*cbar/2Vin) ), 
*    based on wing area and mean aerodynamic chord.
* Cmad_Basic_Table - CM alpha dot ( dCL/d(alpha_dot*cbar/2Vin) ), 
*    based on wing area and mean aerodynamic chord squared.
* Clp_Basic_Table  - Rolling derivative ( dCL/d(p*b/2Vin) ), 
*    based on wing area and wing span squared.
* Cyp_Basic_Table  - Sideforce derivative ( dCL/d(p*b/2Vin) ), 
*    based on wing area and wing span.
* Cnp_Basic_Table  - Yawing derivative ( dCL/d(p*b/2Vin) ), 
*    based on wing area and wing span squared.
* Cnr_Basic_Table  - Yawing derivative ( dCL/d(r*b/2Vin) ), 
*    based on wing area and wing span squared.
* Clr_Basic_Table - Rolling derivative ( dCL/d(r*b/2Vin) ), 
*    based on wing area and wing span squared.
*
*    In Ground Effects (IGE)
*    -----------------------
* Note that flap IGE tables are delta values from free air, while elevator
* tables include entire aircraft, and are actual values IGE
*
* Delta_Cd_IGE_FLAPS_Table - Difference in flap drag coef. due to IGE
* Delta_CL_IGE_FLAPS_Table - Difference in flap lift coef. due to IGE
* Delta_Cm_IGE_FLAPS_Table - Difference in flap pitch mom. coef. due to IGE
* Delta_CLa_IGE_FLAPS_Table - Difference in flap lift deriv. due to IGE
* Delta_Cma_IGE_FLAPS_Table - Difference in flap pitch mom. deriv. due to IGE
* Cd_IGE_TOTAL_Table - Total drag coef IGE for aircraft for 
*    various elevator delflections
* CL_IGE_TOTAL_Table - Total lift coef IGE for aircraft for 
*    various elevator delflections
* Cm_IGE_TOTAL_Table - Total pitch mom. coef IGE for aircraft
*    for various elevator delflections
* CLa_IGE_TOTAL_Table - Total lift coef deriv. IGE for 
*    aircraft for various elevator delflections
* Cma_IGE_TOTAL_Table - Total pitch mom coef deriv.IGE for 
*    aircraft for various elevator delflections
*
*    General Notes
*    -------------
*    1. All tables are in stability axis system unless otherwise noted.
*    2. DATCOM provides predicted data only up to stall for some tables.
*       Beyond stall, if DATCOM did not provide data, the table is clamped
*       at the last valid value.
*    3. Dynamic derivatives from DATCOM are presented for clean 
*       configuration only.
*    4. CL (upper case 'L') represents lift coefficients, 
*       Cl (lower case 'L') represents rolling moment coefficients.
*


***************************************
*  Total Aircraft Data
***************************************



Cd_Basic_Table
   Alpha_At_Wing_deg

         .00	   .1657E-01
      -80.00	   .1021    
      -60.00	   .1265    
      -45.00	   .1782    
      -30.00	   .1690    
      -15.00	   .7973E-01
      -10.00	   .3844E-01
       -5.00	   .1833E-01
       -2.00	   .1522E-01
         .00	   .1657E-01
        2.00	   .2089E-01
        5.00	   .3335E-01
       10.00	   .7115E-01
       15.00	   .1187    
       30.00	   .1310    
       45.00	   .8434E-01
       60.00	   .9433E-01
       75.00	   .1023    


CL_Basic_Table
   Alpha_At_Wing_deg

         .00	   .1204    
      -80.00	  -.9079E-01
      -60.00	  -.8335    
      -45.00	  -1.331    
      -30.00	  -1.357    
      -15.00	  -.8791    
      -10.00	  -.5329    
       -5.00	  -.1954    
       -2.00	  -.6257E-02
         .00	   .1204    
        2.00	   .2515    
        5.00	   .4555    
       10.00	   .8056    
       15.00	   1.098    
       30.00	   1.058    
       45.00	   .6552    
       60.00	   .5441    
       75.00	   .3085    


Cm_Basic_Table
   Alpha_At_Wing_deg

         .00	  -.4544E-01
      -80.00	   59.66    
      -60.00	   10.85    
      -45.00	   1.844    
      -30.00	   .5573    
      -15.00	   .4704    
      -10.00	   .2957    
       -5.00	   .1169    
       -2.00	   .1778E-01
         .00	  -.4544E-01
        2.00	  -.1077    
        5.00	  -.2010    
       10.00	  -.3684    
       15.00	  -.5211    
       30.00	  -.6951    
       45.00	  -.3383    
       60.00	  -.2323    
       75.00	  -.2329E-01


CN_Basic_Table
   Alpha_At_Wing_deg

         .00	   .1204    
      -80.00	  -.1164    
      -60.00	  -.5263    
      -45.00	  -1.067    
      -30.00	  -1.260    
      -15.00	  -.8698    
      -10.00	  -.5314    
       -5.00	  -.1962    
       -2.00	  -.6784E-02
         .00	   .1204    
        2.00	   .2521    
        5.00	   .4566    
       10.00	   .8057    
       15.00	   1.092    
       30.00	   .9819    
       45.00	   .5229    
       60.00	   .3537    
       75.00	   .1786    


CA_Basic_Table
   Alpha_At_Wing_deg

         .00	   .1657E-01
      -80.00	  -.7168E-01
      -60.00	  -.6586    
      -45.00	  -.8151    
      -30.00	  -.5322    
      -15.00	  -.1505    
      -10.00	  -.5467E-01
       -5.00	   .1234E-02
       -2.00	   .1500E-01
         .00	   .1657E-01
        2.00	   .1210E-01
        5.00	  -.6476E-02
       10.00	  -.6982E-01
       15.00	  -.1696    
       30.00	  -.4156    
       45.00	  -.4037    
       60.00	  -.4240    
       75.00	  -.2716    


CLa_Basic_Table
   Alpha_At_Wing_deg

         .00	   .6443E-01
      -80.00	  -.5040E-01
      -60.00	  -.3487E-01
      -45.00	  -.1745E-01
      -30.00	   .1506E-01
      -15.00	   .5990E-01
      -10.00	   .6837E-01
       -5.00	   .6471E-01
       -2.00	   .6321E-01
         .00	   .6444E-01
        2.00	   .6653E-01
        5.00	   .6875E-01
       10.00	   .6430E-01
       15.00	   .4325E-01
       30.00	  -.1477E-01
       45.00	  -.1714E-01
       60.00	  -.1156E-01
       75.00	  -.1985E-01


Cma_Basic_Table
   Alpha_At_Wing_deg

         .00	  -.3127E-01
      -80.00	  -3.005    
      -60.00	  -1.389    
      -45.00	  -.3430    
      -30.00	  -.4579E-01
      -15.00	  -.2766E-01
      -10.00	  -.3535E-01
       -5.00	  -.3406E-01
       -2.00	  -.3218E-01
         .00	  -.3137E-01
        2.00	  -.3112E-01
        5.00	  -.3199E-01
       10.00	  -.3201E-01
       15.00	  -.2580E-01
       30.00	   .6094E-02
       45.00	   .1543E-01
       60.00	   .1050E-01
       75.00	   .1737E-01


Cyb_Basic_Table
   Alpha_At_Wing_deg

         .00	  -.1167E-01
      -80.00	  -.1167E-01
      -60.00	  -.1167E-01
      -45.00	  -.1167E-01
      -30.00	  -.1167E-01
      -15.00	  -.1167E-01
      -10.00	  -.1167E-01
       -5.00	  -.1167E-01
       -2.00	  -.1167E-01
         .00	  -.1167E-01
        2.00	  -.1167E-01
        5.00	  -.1167E-01
       10.00	  -.1167E-01
       15.00	  -.1167E-01
       30.00	  -.1167E-01
       45.00	  -.1167E-01
       60.00	  -.1167E-01
       75.00	  -.1167E-01


Cnb_Basic_Table
   Alpha_At_Wing_deg

         .00	   .2805E-02
      -80.00	   .2805E-02
      -60.00	   .2805E-02
      -45.00	   .2805E-02
      -30.00	   .2805E-02
      -15.00	   .2805E-02
      -10.00	   .2805E-02
       -5.00	   .2805E-02
       -2.00	   .2805E-02
         .00	   .2805E-02
        2.00	   .2805E-02
        5.00	   .2805E-02
       10.00	   .2805E-02
       15.00	   .2805E-02
       30.00	   .2805E-02
       45.00	   .2805E-02
       60.00	   .2805E-02
       75.00	   .2805E-02


Clb_Basic_Table
   Alpha_At_Wing_deg

         .00	  -.1477E-02
      -80.00	  -.3298E-02
      -60.00	  -.3628E-03
      -45.00	   .1848E-02
      -30.00	   .2484E-02
      -15.00	   .1622E-02
      -10.00	   .5367E-03
       -5.00	  -.5211E-03
       -2.00	  -.1094E-02
         .00	  -.1477E-02
        2.00	  -.1881E-02
        5.00	  -.2522E-02
       10.00	  -.3636E-02
       15.00	  -.4475E-02
       30.00	  -.2667E-02
       45.00	  -.3485E-04
       60.00	   .1196E-02
       75.00	   .2642E-02


CLq_Basic_Table
   Alpha_At_Wing_deg

         .00	   .1321    
      -80.00	   .1321    
      -60.00	   .1321    
      -45.00	   .1321    
      -30.00	   .1321    
      -15.00	   .1321    
      -10.00	   .1321    
       -5.00	   .1321    
       -2.00	   .1321    
         .00	   .1321    
        2.00	   .1321    
        5.00	   .1321    
       10.00	   .1321    
       15.00	   .1321    
       30.00	   .1321    
       45.00	   .1321    
       60.00	   .1321    
       75.00	   .1321    


Cmq_Basic_Table
   Alpha_At_Wing_deg

         .00	  -.1399    
      -80.00	  -.1399    
      -60.00	  -.1399    
      -45.00	  -.1399    
      -30.00	  -.1399    
      -15.00	  -.1399    
      -10.00	  -.1399    
       -5.00	  -.1399    
       -2.00	  -.1399    
         .00	  -.1399    
        2.00	  -.1399    
        5.00	  -.1399    
       10.00	  -.1399    
       15.00	  -.1399    
       30.00	  -.1399    
       45.00	  -.1399    
       60.00	  -.1399    
       75.00	  -.1399    


CLad_Basic_Table
   Alpha_At_Wing_deg

         .00	   .2725E-01
      -80.00	   .8484E-03
      -60.00	   .8484E-03
      -45.00	  -.6396E-02
      -30.00	  -.1011E-01
      -15.00	   .1505E-01
      -10.00	   .2401E-01
       -5.00	   .2575E-01
       -2.00	   .2636E-01
         .00	   .2725E-01
        2.00	   .2833E-01
        5.00	   .2960E-01
       10.00	   .2948E-01
       15.00	   .2309E-01
       30.00	  -.1203E-01
       45.00	  -.2036E-01
       60.00	  -.4816E-02
       75.00	  -.4816E-02


Cmad_Basic_Table
   Alpha_At_Wing_deg

         .00	  -.5506E-01
      -80.00	  -.1714E-02
      -60.00	  -.1714E-02
      -45.00	   .1292E-01
      -30.00	   .2043E-01
      -15.00	  -.3040E-01
      -10.00	  -.4851E-01
       -5.00	  -.5203E-01
       -2.00	  -.5326E-01
         .00	  -.5506E-01
        2.00	  -.5722E-01
        5.00	  -.5980E-01
       10.00	  -.5956E-01
       15.00	  -.4666E-01
       30.00	   .2431E-01
       45.00	   .4113E-01
       60.00	   .9729E-02
       75.00	   .9729E-02


Clp_Basic_Table
   Alpha_At_Wing_deg

         .00	  -.5323E-02
      -80.00	   .1006E-02
      -60.00	  -.3103E-03
      -45.00	  -.1353E-02
      -30.00	  -.2862E-02
      -15.00	  -.5526E-02
      -10.00	  -.5949E-02
       -5.00	  -.5384E-02
       -2.00	  -.5205E-02
         .00	  -.5323E-02
        2.00	  -.5519E-02
        5.00	  -.5743E-02
       10.00	  -.5425E-02
       15.00	  -.3633E-02
       30.00	   .8137E-03
       45.00	  -.2818E-03
       60.00	  -.2033E-02
       75.00	  -.2570E-02


Cyp_Basic_Table
   Alpha_At_Wing_deg

         .00	   .3843E-03
      -80.00	  -.8080E-02
      -60.00	  -.7664E-02
      -45.00	  -.1414E-01
      -30.00	  -.1164E-01
      -15.00	  -.6507E-02
      -10.00	  -.4010E-02
       -5.00	  -.1787E-02
       -2.00	  -.4917E-03
         .00	   .3843E-03
        2.00	   .1277E-02
        5.00	   .2649E-02
       10.00	   .5135E-02
       15.00	   .7738E-02
       30.00	   .1119E-01
       45.00	   .1029E-01
       60.00	   .1137E-01
       75.00	   .1170E-01


Cnp_Basic_Table
   Alpha_At_Wing_deg

         .00	  -.2182E-03
      -80.00	   .1246E-03
      -60.00	   .4071E-02
      -45.00	   .4205E-02
      -30.00	   .3855E-02
      -15.00	   .2044E-02
      -10.00	   .1191E-02
       -5.00	   .4977E-03
       -2.00	   .7259E-04
         .00	  -.2182E-03
        2.00	  -.5105E-03
        5.00	  -.9562E-03
       10.00	  -.1835E-02
       15.00	  -.2944E-02
       30.00	  -.4336E-02
       45.00	  -.3819E-02
       60.00	  -.3745E-02
       75.00	  -.2786E-02


Cnr_Basic_Table
   Alpha_At_Wing_deg

         .00	  -.3713E-02
      -80.00	  -.8259E-04
      -60.00	  -.5444E-03
      -45.00	  -.1545E-02
      -30.00	  -.2506E-02
      -15.00	  -.3205E-02
      -10.00	  -.3343E-02
       -5.00	  -.3518E-02
       -2.00	  -.3634E-02
         .00	  -.3713E-02
        2.00	  -.3795E-02
        5.00	  -.3924E-02
       10.00	  -.4159E-02
       15.00	  -.4363E-02
       30.00	  -.3977E-02
       45.00	  -.3068E-02
       60.00	  -.2063E-02
       75.00	  -.1042E-02


Clr_Basic_Table
   Alpha_At_Wing_deg

         .00	   .1762E-02
      -80.00	  -.9096E-03
      -60.00	  -.2509E-02
      -45.00	  -.3570E-02
      -30.00	  -.3208E-02
      -15.00	  -.1673E-02
      -10.00	  -.4123E-03
       -5.00	   .7405E-03
       -2.00	   .1353E-02
         .00	   .1762E-02
        2.00	   .2191E-02
        5.00	   .2863E-02
       10.00	   .4020E-02
       15.00	   .4885E-02
       30.00	   .3152E-02
       45.00	   .7681E-03
       60.00	   .1842E-03
       75.00	  -.2991E-03


Q_Qinf_Table
   Alpha_At_Wing_deg

         .00	   1.000    
      -80.00	   1.000    
      -60.00	   1.000    
      -45.00	   1.000    
      -30.00	   1.000    
      -15.00	   1.000    
      -10.00	   1.000    
       -5.00	   1.000    
       -2.00	   1.000    
         .00	   1.000    
        2.00	   1.000    
        5.00	   1.000    
       10.00	   1.000    
       15.00	   1.000    
       30.00	   1.000    
       45.00	   1.000    
       60.00	   1.000    
       75.00	   1.000    


Epslon_Table
   Alpha_At_Wing_deg

         .00	   1.525    
      -80.00	    .000    
      -60.00	    .000    
      -45.00	    .000    
      -30.00	  -4.312    
      -15.00	  -6.817    
      -10.00	  -4.284    
       -5.00	  -1.420    
       -2.00	   .3269    
         .00	   1.525    
        2.00	   2.777    
        5.00	   4.736    
       10.00	   8.164    
       15.00	   11.36    
       30.00	   13.73    
       45.00	   3.247    
       60.00	    .000    
       75.00	    .000    


d_Epslon_Table
   Alpha_At_Wing_deg

         .00	   .6125    
      -80.00	   .1907E-01
      -60.00	    .000    
      -45.00	  -.1437    
      -30.00	  -.2272    
      -15.00	   .3382    
      -10.00	   .5397    
       -5.00	   .5788    
       -2.00	   .5925    
         .00	   .6125    
        2.00	   .6366    
        5.00	   .6652    
       10.00	   .6625    
       15.00	   .5190    
       30.00	  -.2705    
       45.00	  -.4576    
       60.00	  -.1082    
       75.00	    .000    
         .00	
      -80.00	
      -60.00	
      -45.00	
      -30.00	
      -15.00	
      -10.00	
       -5.00	
       -2.00	
         .00	
        2.00	
        5.00	
       10.00	
       15.00	
       30.00	
       45.00	
       60.00	
       75.00	


CLa_Elevator_Table
   Alpha_At_Wing_deg

         .00	    .000    
      -80.00	    .000    
      -60.00	    .000    
      -45.00	    .000    
      -30.00	    .000    
      -15.00	    .000    
      -10.00	    .000    
       -5.00	    .000    
       -2.00	    .000    
         .00	    .000    
        2.00	    .000    
        5.00	    .000    
       10.00	    .000    
       15.00	    .000    
       30.00	    .000    
       45.00	    .000    
       60.00	    .000    
       75.00	    .000    


Cha_Elevator_Table
   Alpha_At_Wing_deg

         .00	    .000    
      -80.00	    .000    
      -60.00	    .000    
      -45.00	    .000    
      -30.00	    .000    
      -15.00	    .000    
      -10.00	    .000    
       -5.00	    .000    
       -2.00	    .000    
         .00	    .000    
        2.00	    .000    
        5.00	    .000    
       10.00	    .000    
       15.00	    .000    
       30.00	    .000    
       45.00	    .000    
       60.00	    .000    
       75.00	    .000    



