171 lines
4.4 KiB
C
171 lines
4.4 KiB
C
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/***************************************************************************
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TITLE: Cherokee_aero
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----------------------------------------------------------------------------
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FUNCTION: Linear aerodynamics model
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----------------------------------------------------------------------------
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MODULE STATUS: developmental
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----------------------------------------------------------------------------
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GENEALOGY:
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----------------------------------------------------------------------------
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MODIFICATION HISTORY:
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----------------------------------------------------------------------------
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REFERENCES:
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Based upon book:
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Barnes W. McCormick,
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"Aerodynamics, Aeronautics and Flight Mechanics",
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John Wiley & Sons,1995, ISBN 0-471-11087-6
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any suggestions, corrections, aditional data, flames, everything to
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Gordan Sikic
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gsikic@public.srce.hr
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This source is not checked in this configuration in any way.
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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 <float.h>
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#include <math.h>
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#include "ls_types.h"
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#include "ls_generic.h"
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#include "ls_cockpit.h"
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void aero()
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/*float ** Cherokee (float t, VectorStanja &X, float *U)*/
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{
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static float
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Cza = -19149.0/(146.69*146.69*157.5/2.0*0.00238),
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Czat = -73.4*4*146.69/0.00238/157.5/5.25,
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Czq = -2.655*4*2400.0/32.2/0.00238/157.5/146.69/5.25,
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Cma = -21662.0 *2/146.69/0.00238/157.5/146.69/5.25,
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Cmat = -892.4 *4/146.69/0.00238/157.5/146.69/5.25,
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Cmq = -2405.1 *4/0.00238/157.5/146.69/5.25/5.25,
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Czde = -1050.49 *2/0.00238/157.5/146.69/146.69,
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Cmde = -12771.9 *2/0.00238/157.5/146.69/146.69/5.25,
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Clb = -12891.0/(146.69*146.69*157.5/2.0*0.00238)/30.0,
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Clp = -0.4704,
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Clr = 0.1665,
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Cyb = -1169.8/(146.69*146.69*157.5/2.0*0.00238),
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Cyp = -0.0342,
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Cnb = 11127.2/(146.69*146.69*157.5/2.0*0.00238)/30.0,
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Cnp = -0.0691,
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Cnr = -0.0930,
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Cyf = -14.072/(146.69*146.69*157.5/2.0*0.00238),
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Cyps = 89.229/(146.69*146.69*157.5/2.0*0.00238),
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Clf = -5812.4/(146.69*146.69*157.5/2.0*0.00238)/30.0, //%Clda ?
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Cnf = -853.93/(146.69*146.69*157.5/2.0*0.00238)/30.0, //%Cnda ?
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Cnps = -1149.0/(146.69*146.69*157.5/2.0*0.00238)/30.0, //%Cndr ?
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Cyr = 1.923/(146.69*146.69*157.5/2.0*0.00238),
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Cx0 = -0.4645/(157.5*0.3048*0.3048),
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Cz0 = -0.11875,
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Cm0 = 0.0959,
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Clda = -5812.4/(146.69*146.69*157.5/2.0*0.00238)/30.0, // Clf
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Cnda = -853.93/(146.69*146.69*157.5/2.0*0.00238)/30.0, // Cnf
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Cndr = -1149.0/(146.69*146.69*157.5/2.0*0.00238)/30.0, // Cnps
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/*
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Possible problems: convention about positive control surfaces offset
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*/
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elevator = 0.0, // 20.0 * 180.0/57.3 * Long_control
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aileron = 0.0, // 30.0 * 180.0/57.3 * Lat_control
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rudder = 0.0, // 30.0 * 180.0/57.3 * Rudder_pedal,
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// m = 2400/32.2, // mass
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S = 157.5, // wing area
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b = 30.0, // wing span
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c = 5.25, // main aerodynamic chrod
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// 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},
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// Fa[3],
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// Ma[3],
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// *RetVal[4] = {&m, Ixyz, Fa, Ma};
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// float
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V = 0.0, // V_rel_wind
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qd = 0.0, // Density*V*V/2.0, //dinamicki tlak
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Cx,Cy,Cz,
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Cl,Cm,Cn,
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p,q,r;
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/* derivatives are defined in "wind" axes so... */
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p = P_body*Cos_alpha + R_body*Sin_alpha;
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q = Q_body;
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r = -P_body*Sin_alpha + R_body*Cos_alpha;
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Cz = Cz0 + Cza*Alpha + Czat*(Alpha_dot*c/2.0/V) + Czq*(q*c/2.0/V) + Czde * elevator;
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Cm = Cm0 + Cma*Alpha + Cmat*(Alpha_dot*c/2.0/V) + Cmq*(q*c/2.0/V) + Cmde * elevator;
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Cx = Cx0 - (Cza*Alpha)*(Cza*Alpha)/(M_PI*5.71*0.6);
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Cl = Clb*Beta + Clp*(p*b/2.0/V) + Clr*(r*b/2.0/V) + Clda * aileron;
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Cy = Cyb*Beta + Cyr*(r*b/2.0/V);
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Cn = Cnb*Beta + Cnp*(p*b/2.0/V) + Cnr*(r*b/2.0/V) + Cndr * rudder;
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/* back to body axes */
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{
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float
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CD = Cx,
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CL = Cz;
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Cx = CD - CL*Sin_alpha;
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Cz = CL;
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}
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/* AD forces and moments */
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F_X_aero = Cx*qd*S;
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F_Y_aero = Cy*qd*S;
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F_Z_aero = Cz*qd*S;
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M_l_aero = (Cl*Cos_alpha - Cn*Sin_alpha)*b*qd*S;
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M_m_aero = Cm*c*qd*S;
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M_n_aero = (Cl*Sin_alpha + Cn*Cos_alpha)*b*qd*S;
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}
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