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glob.m
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glob.m
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% constants file
% INPUT: NaF
% OUTPUT: all constants
sp= 0.01; % [s] sampling period only used to compute YiA term in dinamica.m
g= 9.806; % [m/s^2]
b= 3.13E-5; % [N.s2] thrust factor in hover
d= 7.5E-7; % [Nm.s2] drag factor in hover
%***** DEG2RAD *****
deg2rad=1.745329E-2; % [rad/deg]
rad2deg=57.295779; % [deg/rad]
%***** Number of propellers *****/
P=4;
%***** Arm length *****/
L= 0.232; %// [m]
%***** Total mass *****/
m = 0.53; %// [kg]
%***** Inertia components *****/
Ixx = 6.228E-3; %// [kg.m2]
Iyy = 6.228E-3; %// [kg.m2]
Izz = 1.121E-2; %// [kg.m2]
%***** Rotor inertia *****/
r=4; % reduction ratio
jm= 4e-7; %// [kg.m2];
jp= 6e-5; %// [kg.m2];
jr = jp+jm/r; %// [kg.m2];
%***** CoG position *****/
% h= 0.058; %// [m] VERTICAL DISTANCE between CoG and propellers plan
h= 0; %// [m] more stable with h=0 !!!
%***** constants of the linear curve Omega=f(bin) *****
slo=2.7542; % slope (of the linear curve om=f(bin))
shi=3.627; % shift
% *************************** AERODYNAMICS ***************************
% ********* Propeller *********
N=2; % number of blades
R=0.15; % [m] propeller radius
A=pi*(R^2); % [m^2] prop disk area
c=0.0394; % [m] chord
theta0=0.2618; % [rad] pitch of incidence
thetatw=0.045; % [rad] twist pitch
sigma_=(N*c)/(pi*R); % solidity ratio (rotor fill ratio) [rad^-1]
a=5.7; % Lift slope (given in literature)
Cd=0.052; % Airfoil drag coefficient -- found by tests
Ac=0.005; % [m^2] helicopter center hub area
% ********* Gaz *********
rho= 1.293; % [kg/m^3] air density
nu=1.8e-5; % [Pa.s] air viscosity @ 20deg
% ********* misc *********
w=(m/P)*g; % [N] weigth of the helicopter/number of propellers
OmegaH=sqrt(w/b); % [rad/s] prop spd at hover
OmegaMax = 600; %[rad/s]
% ********* Longitudinal Drag Coefficients *********
Cx=1.32; % estimation taken from literature [less]
Cy=1.32; % estimation taken from literature [less]
Cz=1.32; % estimation taken from literature [less]
% ***** OS4 volume *****
Vol=3.04E-4; % [m3]
PArchim = rho*g*Vol; % [N]
% ***** ANNEXE ! ******
%Ftb=0.5*Cz*A*rho*v^2 % turbulent
%Fl=16*0.3*nu*v % laminary