prm_rock.m

导弹控制系统的鲁棒控制设计实例

function [m,G,P,J_x,J_y,J_z,x_G,x_C] = prm_rock(t,V,H)
%
% Rocket parameters
% All distances are measured from the rocket nose
%
% Constants
g0          = 9.80665;      % m/s^2   acceleration of gravity at sea level
p0          = 101325;       % N/m^2   air pressure at sea level
%
% Construction parameters
m_0         = 45;           % kg      initial rocket mass
mu          = 0.3;          % kg/s    propellant consumption per second
P_0         = 740*g0;       % N       engine thrust at sea level
S_a         = 0.011;        % m^2     engine nozzle area
L           = 2.1;          % m       rocket length
d           = 0.12;         % m       rocket diameter
x_G_0       = 0.8;          % m       initial rocket mass center coordinate
x_C_0       = 1.05;         % m       initial rocket pressure center coordinate
tk          = 40;           % s       time of active flight
%
% Moments of inertia
J_x_0       = 8.10*10^(-2); % kg.m^2  initial rocket moment of inertia about x-axis
J_y_0       = 9.64;         % kg.m^2  initial rocket moment of inertia about y-axis
J_z_0       = 9.64;         % kg.m^2  initial rocket moment of inertia about z-axis
%
if t <= tk
   J_x = J_x_0 - 0.00054*t;
   J_y = J_y_0 - 0.0643*t;
   J_z = J_z_0 - 0.0643*t;
else
   J_x = J_x_0 - 0.00054*tk;
   J_y = J_y_0 - 0.0643*tk;
   J_z = J_z_0 - 0.0643*tk;
end 
%
[p,ro,M,Tm] = air_data(V,H);
%
% Engine thrust
if t <= tk
   m = m_0  - mu*t;       
   P = P_0 + (p0 - p)*S_a;
else
   m = m_0 - mu*tk;
   P = 0;
end   
%
% Weight
G = m*g0;
%
% Mass center coordinate
if t <= tk
  x_G = x_G_0 - 0.005*t;              % m     
else
  x_G = x_G_0 - 0.005*tk;
end   
%
% Pressure center coordinate
x_C = x_C_0;                         % m