src/FDM/LaRCsim/cherokee_aero.c

   1 /***************************************************************************
   2
   3   TITLE:        Cherokee_aero
   4
   5 ----------------------------------------------------------------------------
   6
   7   FUNCTION:     Linear aerodynamics model
   8
   9 ----------------------------------------------------------------------------
  10
  11   MODULE STATUS:        developmental
  12
  13 ----------------------------------------------------------------------------
  14
  15   GENEALOGY:
  16
  17 ----------------------------------------------------------------------------
  18
  19   MODIFICATION HISTORY:
  20
  21
  22 ----------------------------------------------------------------------------
  23
  24   REFERENCES:
  25
  26                 Based upon book:
  27                                 Barnes W. McCormick,
  28                                 "Aerodynamics, Aeronautics and Flight Mechanics",
  29                                 John Wiley & Sons,1995, ISBN 0-471-11087-6
  30
  31         any suggestions, corrections, aditional data, flames, everything to
  32         Gordan Sikic
  33         gsikic@public.srce.hr
  34
  35 This source is not checked in this configuration in any way.
  36
  37
  38 ----------------------------------------------------------------------------
  39
  40   CALLED BY:
  41
  42 ----------------------------------------------------------------------------
  43
  44   CALLS TO:
  45
  46 ----------------------------------------------------------------------------
  47
  48   INPUTS:
  49
  50 ----------------------------------------------------------------------------
  51
  52   OUTPUTS:
  53
  54 --------------------------------------------------------------------------*/
  55
  56
  57
  58
  59 #include <float.h>
  60 #include <math.h>
  61 #include "ls_types.h"
  62 #include "ls_generic.h"
  63 #include "ls_cockpit.h"
  64 #include "ls_constants.h"
  65
  66
  67
  68
  69
  70 void cherokee_aero()
  71 /*float ** Cherokee (float t, VectorStanja &X, float *U)*/
  72 {
  73         static float
  74                 Cza  = -19149.0/(146.69*146.69*157.5/2.0*0.00238),
  75                 Czat = -73.4*4*146.69/0.00238/157.5/5.25,
  76                 Czq  = -2.655*4*2400.0/32.2/0.00238/157.5/146.69/5.25,
  77                 Cma  = -21662.0 *2/146.69/0.00238/157.5/146.69/5.25,
  78                 Cmat = -892.4 *4/146.69/0.00238/157.5/146.69/5.25,
  79                 Cmq  = -2405.1 *4/0.00238/157.5/146.69/5.25/5.25,
  80                 Czde = -1050.49 *2/0.00238/157.5/146.69/146.69,
  81                 Cmde = -12771.9 *2/0.00238/157.5/146.69/146.69/5.25,
  82                 Clb  = -12891.0/(146.69*146.69*157.5/2.0*0.00238)/30.0,
  83                 Clp  = -0.4704,
  84                 Clr  = 0.1665,
  85                 Cyb  = -1169.8/(146.69*146.69*157.5/2.0*0.00238),
  86                 Cyp  = -0.0342,
  87                 Cnb  = 11127.2/(146.69*146.69*157.5/2.0*0.00238)/30.0,
  88                 Cnp  = -0.0691,
  89                 Cnr  = -0.0930,
  90                 Cyf  = -14.072/(146.69*146.69*157.5/2.0*0.00238),
  91                 Cyps = 89.229/(146.69*146.69*157.5/2.0*0.00238),
  92                 Clf  = -5812.4/(146.69*146.69*157.5/2.0*0.00238)/30.0,  //%Clda ?
  93                 Cnf  = -853.93/(146.69*146.69*157.5/2.0*0.00238)/30.0,  //%Cnda ?
  94                 Cnps = -1149.0/(146.69*146.69*157.5/2.0*0.00238)/30.0,  //%Cndr ?
  95                 Cyr  = 1.923/(146.69*146.69*157.5/2.0*0.00238),
  96
  97                 Cx0 = -0.4645/(157.5*0.3048*0.3048),
  98
  99                 Cz0 = -0.11875,
 100                 Cm0 =  0.0959,
 101
 102                 Clda = -5812.4/(146.69*146.69*157.5/2.0*0.00238)/30.0, // Clf
 103                 Cnda = -853.93/(146.69*146.69*157.5/2.0*0.00238)/30.0, // Cnf
 104                 Cndr = -1149.0/(146.69*146.69*157.5/2.0*0.00238)/30.0, // Cnps
 105
 106 /*
 107         Possible problems: convention about positive control surfaces offset
 108 */
 109                 elevator = 0.0, // 20.0 * 180.0/57.3 * Long_control
 110                 aileron  = 0.0, // 30.0 * 180.0/57.3 * Lat_control
 111                 rudder   = 0.0, // 30.0 * 180.0/57.3 * Rudder_pedal,
 112
 113
 114 //              m = 2400/32.2,          // mass
 115                 S = 157.5,                      // wing area
 116                 b = 30.0,                       // wing span
 117                 c = 5.25,                       // main aerodynamic chrod
 118
 119 //              Ixyz[3] = {1070.0*14.59*0.3048*0.3048, 1249.0*14.59*0.3048*0.3048, 2312.0*14.59*0.3048*0.3048},
 120 //              Fa[3],
 121 //              Ma[3],
 122 //              *RetVal[4] = {&m, Ixyz, Fa, Ma};
 123
 124
 125 //      float
 126                 V = 0.0, // V_rel_wind
 127                 qd = 0.0, // Density*V*V/2.0,                   //dinamicki tlak
 128
 129                 Cx,Cy,Cz,
 130                 Cl,Cm,Cn,
 131                 p,q,r;
 132
 133
 134 /* derivatives are defined in "wind" axes so... */
 135                 p =  P_body*Cos_alpha + R_body*Sin_alpha;
 136                 q =  Q_body;
 137                 r = -P_body*Sin_alpha + R_body*Cos_alpha;
 138
 139
 140
 141                 Cz = Cz0 + Cza*Alpha + Czat*(Alpha_dot*c/2.0/V) + Czq*(q*c/2.0/V) + Czde * elevator;
 142                 Cm = Cm0 + Cma*Alpha + Cmat*(Alpha_dot*c/2.0/V) + Cmq*(q*c/2.0/V) + Cmde * elevator;
 143
 144                 Cx = Cx0 - (Cza*Alpha)*(Cza*Alpha)/(PI*5.71*0.6);
 145                 Cl = Clb*Beta + Clp*(p*b/2.0/V) + Clr*(r*b/2.0/V) + Clda * aileron;
 146
 147                 Cy = Cyb*Beta + Cyr*(r*b/2.0/V);
 148                 Cn = Cnb*Beta + Cnp*(p*b/2.0/V) + Cnr*(r*b/2.0/V) + Cndr * rudder;
 149
 150 /* back to body axes */
 151         {
 152                 float
 153                         CD = Cx,
 154                         CL = Cz;
 155
 156                         Cx = CD - CL*Sin_alpha;
 157                         Cz = CL;
 158         }
 159
 160 /* AD forces and moments   */
 161                 F_X_aero = Cx*qd*S;
 162                 F_Y_aero = Cy*qd*S;
 163                 F_Z_aero = Cz*qd*S;
 164
 165                 M_l_aero = (Cl*Cos_alpha - Cn*Sin_alpha)*b*qd*S;
 166                 M_m_aero = Cm*c*qd*S;
 167                 M_n_aero = (Cl*Sin_alpha + Cn*Cos_alpha)*b*qd*S;
 168 }
 169
 170
 171