Aircraft-Designer/mainExample.m at master · iwknow/Aircraft-Designer · GitHub

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% Project Nickel Squirrel

% ============== Design Spec ==============
Spec = struct( ...
    'StallSpeed',117, ...               % ft/s
    'Range', 1.32E6, ...                % ft
    'Endurance', 14400, ...             % sec
    'Payload', 400, ...                 % lb
    'ClimbRate', 30, ...                % ft/s
    'CruisingSpeed',234 ...             % ft/s
    );
% ============== World Constants ==============
WorldConstants = struct(...
    'AirDensity',0.0019, ...               % Air density [slug/ft^3]
    'kinematicViscosity', 1.529e-4,...      % kinematic Viscosity [ft^2/s]
    'airSpecificHeatRatio', 1.4,...         % airSpecificHeatRatio [Null]
    'airSpecificGasConstant', 1716,...      % airSpecificGasConstant [ft lb/slug oR]
    'airTemperature',527.67);               % airTemperature [R]

% Parameters;
W_to = 3000;                            %takeoff weight [lb]

% ============== build components ==============
% fuselage
L = 20.044283750055012;             % fuselarge length [ft]
D = 4.5;              % fuselarge deepth [ft]
fuselage = Fuselage(W_to, L, D);
% wing
span = 30.188353302222050;          % chord length
cla = 6.500000000000000;          % 2D cl alpha [/rad]
c_mac = -0.1;      % 2D moment coefficient C_m(1/4)
s = 1.10630648766614e+02;            % wing area
e = 1/(1.05+0.007 * pi * span/(s/span));            % Oswald efficiency number
taper = 0.4;       % Taper Ratio
t = 0.18;           % maximum thickness of wing over chord (t/c)m
stallAngle = 15/180*pi;     % stall angle
xcRatio = 0.3;      % chordwise location of the airfoil maximum thickness point (x/c)m [Raymer 283]
wing = Wing(W_to,cla,c_mac,s,span, t, xcRatio,taper , stallAngle,e);
% fuel system
fuelCapacity = 90;     % [lb]
initFuel = 90;         % [lb]
fuelSys = FuelSystem(fuelCapacity,initFuel);
% Horizontal Tail
b = 15.858803965044864;                      % tail span [ft]
cla = 6.2;                  % 2D cl alpha
s = 50;                     % wing area
t = 0.12;                   % maximum thickness of wing over chord (t/c)m
xcRatio = 0.3;              % chordwise location of the airfoil maximum thickness point (x/c)m [Raymer 283]
e = 1;                      % Oswald efficiency number
it = 1/180*pi;             % incidence angle [/rad]
tau = 0.35;                  % elevator area persentage
de = -0/180*pi;                     % initial elevator angle [rad]
max_de = 30/180*pi;         % max abs of elevator angle [rad]
htail = HTail(W_to,b,cla,s, t,xcRatio, e, it,tau, de, max_de);
% Vertical Tail
s = 15;                      % area in ft^2
b = 5;                      % span in ft
t = 0.12;                     % maximum thickness of wing over chord (t/c)m
xcRatio = 0.3;      % chordwise location of the airfoil maximum thickness point (x/c)m [Raymer 283]
vtail = VTail(W_to,s,b,t, xcRatio);
% Engine
propD = 5.3;              % propeller diameter [ft]
engine = Engine(propD);

% ============== init Aircraft ==============
testUAV = AirCraft(W_to, fuselage,wing,fuelSys, htail,vtail,engine);
% Use the Following Flight Condition
testUAV.AoA = 2/180*pi;
testUAV.alt = 3300;
testUAV.v = 150;
testUAV.payloadWeight = 400;        % here the weight must be the payload the UAV is going to pickup
% Use the Following Layout
% Easy-Layout
FuselageNosetoWingFrontRatio = 0.1;
TailtoFuselageFrontRatio = 0.91;
FueltoFuselageFrontRatio = 0.45;
PayloadtoFuselageFrontRatio = 0.22;

fuselageToChord = testUAV.fuselage.L/testUAV.wing.getChord;
testUAV.h_acw = 1/4;
testUAV.h_fuselage = fuselageToChord * (1/2 - FuselageNosetoWingFrontRatio) ;
testUAV.h_act = testUAV.h_fuselage + fuselageToChord * (TailtoFuselageFrontRatio - 1/2);
testUAV.h_acvt = testUAV.h_fuselage + fuselageToChord * (TailtoFuselageFrontRatio - 1/2);
testUAV.h_engine = 1/4;
testUAV.h_fuelSys = testUAV.h_fuselage + fuselageToChord * (FueltoFuselageFrontRatio - 1/2);
testUAV.h_payload = testUAV.h_fuselage + fuselageToChord * (PayloadtoFuselageFrontRatio - 1/2);

% Automatically match weigt
while abs(testUAV.W_to - testUAV.getWeight) > 5
    testUAV.W_to = testUAV.getWeight;
    testUAV.fuselage.W_to = testUAV.getWeight;
    testUAV.wing.W_to = testUAV.getWeight;
    testUAV.hTail.W_to = testUAV.getWeight;
    testUAV.vTail.W_to = testUAV.getWeight;
end

% ============== test Aircraft ==============
% let aircraft cruise. this step is not necessary for test aircraft.
% this is just for getting cruising data.
[testUAV.v,testUAV.AoA,testUAV.hTail.de,testUAV.engine.throttle] = testUAV.cruise(WorldConstants, Spec);
% put aircraft into wintunnel and get test result.
result = windTunnel(testUAV,WorldConstants,Spec);


% ============== Draw Graphs ==============
% gt = GraphTool(testUAV,WorldConstants,70:300);
%  gt.drawTrimVsVelocity();
%  gt.drawDragVsVelocity();
%  gt.drawPowerVsVelocity()
%  gt.drawClVsCd();