336 lines
8.1 KiB
C
336 lines
8.1 KiB
C
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/***************************************************************************
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TITLE: c172_aero
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----------------------------------------------------------------------------
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FUNCTION: aerodynamics model based on constant stability derivatives
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----------------------------------------------------------------------------
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MODULE STATUS: developmental
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----------------------------------------------------------------------------
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GENEALOGY: Based on data from:
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Part 1 of Roskam's S&C text
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The FAA type certificate data sheet for the 172
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Various sources on the net
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John D. Anderson's Intro to Flight text (NACA 2412 data)
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UIUC's airfoil data web site
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----------------------------------------------------------------------------
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DESIGNED BY: Tony Peden
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CODED BY: Tony Peden
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MAINTAINED BY: Tony Peden
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----------------------------------------------------------------------------
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MODIFICATION HISTORY:
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DATE PURPOSE BY
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6/10/99 Initial test release
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----------------------------------------------------------------------------
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REFERENCES:
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Aero Coeffs:
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CL lift
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Cd drag
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Cm pitching moment
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Cy sideforce
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Cn yawing moment
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Croll,Cl rolling moment (yeah, I know. Shoot me.)
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Subscripts
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o constant i.e. not a function of alpha or beta
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a alpha
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adot d(alpha)/dt
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q pitch rate
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qdot d(q)/dt
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beta sideslip angle
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p roll rate
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r yaw rate
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da aileron deflection
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de elevator deflection
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dr rudder deflection
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s stability axes
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----------------------------------------------------------------------------
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CALLED BY:
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----------------------------------------------------------------------------
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CALLS TO:
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----------------------------------------------------------------------------
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INPUTS:
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----------------------------------------------------------------------------
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OUTPUTS:
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--------------------------------------------------------------------------*/
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#include "ls_generic.h"
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#include "ls_cockpit.h"
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#include "ls_constants.h"
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#include "ls_types.h"
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#include <math.h>
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#include <stdio.h>
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#define NCL 11
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#define DYN_ON_SPEED 33 /*20 knots*/
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#ifdef USENZ
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#define NZ generic_.n_cg_body_v[2]
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#else
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#define NZ 1
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#endif
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extern COCKPIT cockpit_;
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FILE *out;
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SCALAR interp(SCALAR *y_table, SCALAR *x_table, int Ntable, SCALAR x)
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{
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SCALAR slope;
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int i=1;
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float y;
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/* if x is outside the table, return value at x[0] or x[Ntable-1]*/
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if(x <= x_table[0])
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{
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y=y_table[0];
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/* printf("x smaller than x_table[0]: %g %g\n",x,x_table[0]); */
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}
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else if(x >= x_table[Ntable-1])
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{
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y=y_table[Ntable-1];
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/* printf("x larger than x_table[N]: %g %g %d\n",x,x_table[NCL-1],Ntable-1); */
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}
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else /*x is within the table, interpolate linearly to find y value*/
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{
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while(x_table[i] <= x) {i++;}
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slope=(y_table[i]-y_table[i-1])/(x_table[i]-x_table[i-1]);
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/* printf("x: %g, i: %d, cl[i]: %g, cl[i-1]: %g, slope: %g\n",x,i,y_table[i],y_table[i-1],slope); */
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y=slope*(x-x_table[i-1]) +y_table[i-1];
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}
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return y;
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}
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void record()
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{
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fprintf(out,"%g,%g,%g,%g,%g,%g,%g,%g,%g,",Long_control,Lat_control,Rudder_pedal,Aft_trim,Fwd_trim,V_rel_wind,Dynamic_pressure,P_body,R_body);
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fprintf(out,"%g,%g,%g,%g,%g,%g,%g,%g,%g,%g,",Alpha,Cos_alpha,Sin_alpha,Alpha_dot,Q_body,Theta_dot,Sin_theta,Cos_theta,Beta,Cos_beta,Sin_beta);
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fprintf(out,"%g,%g,%g,%g,%g,%g,%g,%g\n",Sin_phi,Cos_phi,F_X_aero,F_Y_aero,F_Z_aero,M_l_aero,M_m_aero,M_n_aero);
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fflush(out);
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}
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void aero( SCALAR dt, int Initialize ) {
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static int init = 0;
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static SCALAR trim_inc = 0.0002;
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SCALAR long_trim;
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SCALAR elevator, aileron, rudder;
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static SCALAR alpha_ind[NCL]={-0.087,0,0.175,0.209,0.24,0.262,0.278,0.303,0.314,0.332,0.367};
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static SCALAR CLtable[NCL]={-0.14,0.31,1.21,1.376,1.51249,1.591,1.63,1.60878,1.53712,1.376,1.142};
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/*Note that CLo,Cdo,Cmo will likely change with flap setting so
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they may not be declared static in the future */
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static SCALAR CLadot=1.7;
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static SCALAR CLq=3.9;
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static SCALAR CLde=0.43;
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static SCALAR CLo=0;
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static SCALAR Cdo=0.031;
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static SCALAR Cda=0.13; /*Not used*/
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static SCALAR Cdde=0.06;
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static SCALAR Cma=-0.89;
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static SCALAR Cmadot=-5.2;
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static SCALAR Cmq=-12.4;
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static SCALAR Cmo=-0.062;
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static SCALAR Cmde=-1.28;
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static SCALAR Clbeta=-0.089;
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static SCALAR Clp=-0.47;
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static SCALAR Clr=0.096;
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static SCALAR Clda=0.178;
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static SCALAR Cldr=0.0147;
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static SCALAR Cnbeta=0.065;
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static SCALAR Cnp=-0.03;
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static SCALAR Cnr=-0.099;
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static SCALAR Cnda=-0.053;
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static SCALAR Cndr=-0.0657;
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static SCALAR Cybeta=-0.31;
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static SCALAR Cyp=-0.037;
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static SCALAR Cyr=0.21;
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static SCALAR Cyda=0.0;
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static SCALAR Cydr=0.187;
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/*nondimensionalization quantities*/
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/*units here are ft and lbs */
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static SCALAR cbar=4.9; /*mean aero chord ft*/
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static SCALAR b=35.8; /*wing span ft */
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static SCALAR Sw=174; /*wing planform surface area ft^2*/
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static SCALAR rPiARe=0.054; /*reciprocal of Pi*AR*e*/
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SCALAR W=Mass/INVG;
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SCALAR CLwbh,CL,cm,cd,cn,cy,croll,cbar_2V,b_2V,qS,qScbar,qSb,ps,rs;
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SCALAR F_X_wind,F_Y_wind,F_Z_wind,W_X,W_Y,W_Z;
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if (Initialize != 0)
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{
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out=fopen("flight.csv","w");
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/* Initialize aero coefficients */
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}
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record();
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/*
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LaRCsim uses:
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Cm > 0 => ANU
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Cl > 0 => Right wing down
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Cn > 0 => ANL
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so:
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elevator > 0 => AND -- aircraft nose down
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aileron > 0 => right wing up
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rudder > 0 => ANL
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*/
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if(Aft_trim) long_trim = long_trim - trim_inc;
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if(Fwd_trim) long_trim = long_trim + trim_inc;
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/*scale pct control to degrees deflection*/
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if ((Long_control+long_trim) <= 0)
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elevator=(Long_control+long_trim)*-28*DEG_TO_RAD;
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else
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elevator=(Long_control+long_trim)*23*DEG_TO_RAD;
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aileron = Lat_control*17.5*DEG_TO_RAD;
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rudder = Rudder_pedal*16*DEG_TO_RAD;
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/*check control surface travel limits*/
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/* if((elevator+long_trim) > 23)
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elevator=23;
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else if((elevator+long_trim) < -28)
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elevator=-23; */
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/*
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The aileron travel limits are 20 deg. TEU and 15 deg TED
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but since we don't distinguish between left and right we'll
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use the average here (17.5 deg)
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*/
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/* if(fabs(aileron) > 17.5)
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aileron = 17.5;
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if(fabs(rudder) > 16)
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rudder = 16; */
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/*calculate rate derivative nondimensionalization (is that a word?) factors */
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/*hack to avoid divide by zero*/
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/*the dynamic terms might be negligible at low ground speeds anyway*/
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if(V_rel_wind > DYN_ON_SPEED)
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{
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cbar_2V=cbar/(2*V_rel_wind);
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b_2V=b/(2*V_rel_wind);
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}
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else
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{
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cbar_2V=0;
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b_2V=0;
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}
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/*calcuate the qS nondimensionalization factors*/
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qS=Dynamic_pressure*Sw;
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qScbar=qS*cbar;
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qSb=qS*b;
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/*transform the aircraft rotation rates*/
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ps=-P_body*Cos_alpha + R_body*Sin_alpha;
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rs=-P_body*Sin_alpha + R_body*Cos_alpha;
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/* sum coefficients */
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CLwbh = interp(CLtable,alpha_ind,NCL,Alpha);
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CL = CLo + CLwbh + (CLadot*Alpha_dot + CLq*Theta_dot)*cbar_2V + CLde*elevator;
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cd = Cdo + rPiARe*CL*CL + Cdde*elevator;
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cy = Cybeta*Beta + (Cyp*ps + Cyr*rs)*b_2V + Cyda*aileron + Cydr*rudder;
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cm = Cmo + Cma*Alpha + (Cmq*Theta_dot + Cmadot*Alpha_dot)*cbar_2V + Cmde*(elevator+long_trim);
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cn = Cnbeta*Beta + (Cnp*ps + Cnr*rs)*b_2V + Cnda*aileron + Cndr*rudder;
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croll=Clbeta*Beta + (Clp*ps + Clr*rs)*b_2V + Clda*aileron + Cldr*rudder;
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/*calculate wind axes forces*/
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F_X_wind=-1*cd*qS;
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F_Y_wind=cy*qS;
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F_Z_wind=-1*CL*qS;
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/*calculate moments and body axis forces */
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/*find body-axis components of weight*/
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/*with earth axis to body axis transform */
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W_X=-1*W*Sin_theta;
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W_Y=W*Sin_phi*Cos_theta;
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W_Z=W*Cos_phi*Cos_theta;
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/* requires ugly wind-axes to body-axes transform */
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F_X_aero = W_X + F_X_wind*Cos_alpha*Cos_beta - F_Y_wind*Cos_alpha*Sin_beta - F_Z_wind*Sin_alpha;
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F_Y_aero = W_Y + F_X_wind*Sin_beta + F_Z_wind*Cos_beta;
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F_Z_aero = W_Z*NZ + F_X_wind*Sin_alpha*Cos_beta - F_Y_wind*Sin_alpha*Sin_beta + F_Z_wind*Cos_alpha;
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/*no axes transform here */
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M_l_aero = I_xx*croll*qSb;
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M_m_aero = I_yy*cm*qScbar;
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M_n_aero = I_zz*cn*qSb;
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}
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