diff --git a/classes/components/Para.m b/classes/components/Para.m
index e5faa8b44d352fb2ad3dc1307590f479c00a0eff..d62907e8fffabd3b97f9ab7010233accc13f1040 100644
--- a/classes/components/Para.m
+++ b/classes/components/Para.m
@@ -28,7 +28,7 @@ classdef Para < matlab.mixin.Heterogeneous & Component
varIn = []
end
obj@Component(mission, varIn);
-
+
%% Forcing creation of heterogeneous array
if any(size(obj) > 1)
switch class(obj)
diff --git a/classes/components/Parafoil.m b/classes/components/Parafoil.m
index c79ac4ac1ff56f5804463748f37b1ea68acf7f5a..7ccbb5f1a62951c7ba6caeb3ec436f6eef851907 100644
--- a/classes/components/Parafoil.m
+++ b/classes/components/Parafoil.m
@@ -13,28 +13,33 @@ classdef Parafoil < Para
properties
name = ''
surface double % [m^2] Surface
- deltaSMax double % [-] aerodynamic control coefficient - symmetric, max value
+ deltaSMax double % [-] Aerodynamic control coefficient - symmetric, max value
- mass double % [kg] Parachute Mass
+ mass double % [kg] Parafoil Mass
inertia % [kg*m^2] 3x3 Inertia Matrix (the diagonal is made of the inertia vector values)
-
- semiWingSpan double % [m] semiwingspan
- MAC double % [m] mean aero chord
+ finalAltitude double % [m] Final altitude of the parafoil
+ semiWingSpan double % [m] Parafoil semiwingspan
+ MAC double % [m] Parafoil mean aero chord
cd0 double % [-] aerodynamic control coefficient - asymetric, CD0
- cdAlpha2 double % [-] aerodynamic control coefficient - asymetric, CDAlpha2
- cdDeltaA double % [-] aerodynamic control coefficient - asymetric, CDDeltaA
+ cdAlpha double % [-] aerodynamic control coefficient - asymetric, CDAlpha
+ cdSurface double % [-] aerodynamic control coefficient - asymetric, CDSurface
+
cl0 double % [-] aerodynamic control coefficient - asymetric, CL0
clAlpha double % [-] aerodynamic control coefficient - asymetric, CLAlpha
- clDeltaA double % [-] aerodynamic control coefficient - asymetric, CLDeltaA
- clP double % [-] aerodynamic control coefficient - asymetric, Clp
- clPhi double % [-] aerodynamic control coefficient - asymetric, ClPhi
- cm0 double % [-] aerodynamic control coefficient - asymetric, Cm0
- cmAlpha double % [-] aerodynamic control coefficient - asymetric, CmAlpha
- cmQ double % [-] aerodynamic control coefficient - asymetric, Cmq
- cnR double % [-] aerodynamic control coefficient - asymetric, Cnr
- cnDeltaA double % [-] aerodynamic control coefficient - asymetric, CnDeltaA
+ clSurface double % [-] aerodynamic control coefficient - asymetric, CLSurface
+
+ cLP double % [-] aerodynamic control coefficient - asymetric, Clp
+ cLPhi double % [-] aerodynamic control coefficient - asymetric, ClPhi
+ cLSurface
+
+ cM0 double % [-] aerodynamic control coefficient - asymetric, Cm0
+ cMAlpha double % [-] aerodynamic control coefficient - asymetric, CmAlpha
+ cMQ double % [-] aerodynamic control coefficient - asymetric, Cmq
+
+ cNR double % [-] aerodynamic control coefficient - asymetric, Cnr
+ cNSurface double % [-] aerodynamic control coefficient - asymetric, CnDeltaA
end
properties(SetAccess = private)
@@ -52,7 +57,13 @@ classdef Parafoil < Para
obj = obj(:, :, 1);
end
+ function set.inertia(obj, inertia)
+ obj.inertia = inertia;
+ obj.updateAll();
+ end
+
function updateAll(obj)
+ if isempty(obj.inertia), return; end
obj.inverseInertia = obj.inertia\eye(3);
end
end
diff --git a/functions/eventFunctions/eventParaCut.m b/functions/eventFunctions/eventParaCut.m
index 150286bc2534444f30cf40142b751ef06ef4de91..d3055c3e8b61dd23f8ec5c7bef6d0736919bd900 100644
--- a/functions/eventFunctions/eventParaCut.m
+++ b/functions/eventFunctions/eventParaCut.m
@@ -48,6 +48,5 @@ value = altitude - rocket.parachutes(para, stage).finalAltitude;
isterminal = 1;
direction = 0;
-
end
diff --git a/functions/odeFunctions/descentParafoil.m b/functions/odeFunctions/descentParafoil.m
index 1a328c55fa82631a0cf1e562d9d60b9f89d9b374..e45670d91e5219894e40511af4eea2a0cd3b7d78 100644
--- a/functions/odeFunctions/descentParafoil.m
+++ b/functions/odeFunctions/descentParafoil.m
@@ -1,10 +1,11 @@
-function [dY,parout] = descentParafoil(t, Y, rocket, environment, wind, descentData, wrapper)
+function [dY,parout] = descentParafoil(t, Y, rocket, environment, wind, ~, descentData, wrapper)
arguments
t
Y
rocket Rocket
environment Environment
wind % WindCustom {mustBeA(wind, {'WindCustom', 'WindMatlab'})}
+ ~
descentData struct
wrapper = [] %DataWrapper = DataWrapper.empty
end
@@ -75,17 +76,18 @@ inertia = parafoil.inertia; % 3x3 inertia matr
inverseInertia = parafoil.inverseInertia; % 3x3 inverse inertia matrix
% aerodynamic coefficients
-CD0 = parafoil.cd0; CDAlpha2 = parafoil.cdAlpha2;
-CL0 = parafoil.cl0; CLAlpha = parafoil.clAlpha;
-Cm0 = parafoil.cm0; CmAlpha = parafoil.cmAlpha; Cmq = parafoil.cmQ;
-Cnr = parafoil.cnR;
-Clp = parafoil.clP; ClPhi = parafoil.clPhi;
+cd0 = parafoil.cd0; cdAlpha = parafoil.cdAlpha;
+cl0 = parafoil.cl0; clAlpha = parafoil.clAlpha;
+cM0 = parafoil.cM0; cMAlpha = parafoil.cMAlpha; cMQ = parafoil.cMQ;
+cNR = parafoil.cNR;
+cLP = parafoil.cLP; ClPhi = parafoil.cLPhi;
% aerodynamic control coefficients - asymmetric
-CnDeltaA = parafoil.cnDeltaA;
-CDDeltaA = parafoil.cdDeltaA;
-CLDeltaA = parafoil.clDeltaA;
-ClDeltaA = parafoil.clDeltaA; % !!!
+cNSurface = parafoil.cNSurface;
+cdSurface = parafoil.cdSurface;
+cLSurface = parafoil.cLSurface;
+clSurface = parafoil.clSurface;
+
% aerodynamic control coefficients - symmetric
deltaSMax = parafoil.deltaSMax; % max value
@@ -138,8 +140,8 @@ deltaANormalized = deltaA / deltaSMax;
%% FORCES
qFactor = 0.5*rho*nNorm; % [Pa * s / m] dynamic pressure/vNorm
-lift = qFactor * (CL0 + alpha * CLAlpha + deltaANormalized * CLDeltaA) * [wr; 0; -ur];
-drag = qFactor * (CD0 + alpha^2 * CDAlpha2 + deltaANormalized * CDDeltaA) * [ur; vr; wr];
+lift = qFactor * (cl0 + alpha * clAlpha + deltaANormalized * cLSurface) * [wr; 0; -ur];
+drag = qFactor * (cd0 + alpha^2 * cdAlpha + deltaANormalized * cdSurface) * [ur; vr; wr];
forceAero = (lift - drag)*surface;
@@ -150,9 +152,9 @@ Fg = dcm*[0; 0; mass*g]; % [N] force due to the gravity in body frame
F = Fg + forceAero; % [N] total forces vector
%% MOMENT
-Mx = (eul(1) * ClPhi *nNorm + 0.5 * semiWingSpan* Clp * p + deltaANormalized * ClDeltaA *nNorm) * semiWingSpan;
-My = (alpha * CmAlpha *nNorm+ 0.5 * MAC * Cmq * q + Cm0*nNorm) * MAC;
-Mz = (r * 0.5 * semiWingSpan* Cnr + deltaANormalized * CnDeltaA*nNorm) * semiWingSpan; % bizzarre
+Mx = (eul(1) * ClPhi *nNorm + 0.5 * semiWingSpan* cLP * p + deltaANormalized * clSurface *nNorm) * semiWingSpan;
+My = (alpha * cMAlpha *nNorm+ 0.5 * MAC * cMQ * q + cM0*nNorm) * MAC;
+Mz = (r * 0.5 * semiWingSpan* cNR + deltaANormalized * cNSurface*nNorm) * semiWingSpan; % bizzarre
momentAero = [Mx; My; Mz] * qFactor * surface;
%% derivatives computations
@@ -170,7 +172,7 @@ OM = [ 0 -p -q -r ;
dQQ = 1/2*OM*Q';
%% angular acceleration
-angAcc = inverseInertia*(momentAero - cross(omega, inertia .* omega));
+angAcc = inverseInertia*(momentAero - cross(omega, inertia * omega));
% %% actuator dynamics
%
@@ -205,9 +207,6 @@ dY(1:3) = vNED;
dY(4:6) = bodyAcc;
dY(7:9) = angAcc;
dY(10:13) = dQQ;
-% dY(14) = ddeltaA;
-
-dY = dY';
if nargout == 2 || ~isempty(wrapper)
parout.state = [];
diff --git a/missions/2024_Lyra_Portugal_October/config/paraConfig.m b/missions/2024_Lyra_Portugal_October/config/paraConfig.m
index 24ff5c0a1a827f43fc34bab0787f04d4eda068aa..0946cf824bd685d013748ffd4b96eb760d8581a4 100644
--- a/missions/2024_Lyra_Portugal_October/config/paraConfig.m
+++ b/missions/2024_Lyra_Portugal_October/config/paraConfig.m
@@ -76,16 +76,24 @@ para(2, 2).surface = 2.04; % [m^2] parafoil surface - vel
para(2, 2).inertia = [0.42, 0, 0.03;
0, 0.4, 0;
0.03, 0, 0.053]; % [kg m^2] [3x3] inertia matrix parafoil
-para(2, 2).cd0 = 0.25;
-para(2, 2).cdAlpha2 = 0.12;
-para(2, 2).cdDeltaA = 0.01;
-para(2, 2).cl0 = 0.091;
-para(2, 2).clAlpha = 0.9;
-para(2, 2).clDeltaA = -0.0035;
-para(2, 2).clP = -0.84;
-para(2, 2).clPhi = -0.1;
-para(2, 2).cm0 = 0.35;
-para(2, 2).cmAlpha = -0.72;
-para(2, 2).cmQ = -1.49;
-para(2, 2).cnR = -0.27;
-para(2, 2).cnDeltaA = 0.0115;
\ No newline at end of file
+
+para(2, 2).finalAltitude = 0;
+
+para(2, 2).cd0 = 0.25;
+para(2, 2).cdAlpha = 0.12;
+para(2, 2).cdSurface = 0.01;
+
+para(2, 2).cl0 = 0.091;
+para(2, 2).clAlpha = 0.9;
+para(2, 2).clSurface = -0.0035;
+
+para(2, 2).cLP = -0.84;
+para(2, 2).cLPhi = -0.1;
+para(2, 2).cLSurface = -0.0035;
+
+para(2, 2).cM0 = 0.35;
+para(2, 2).cMAlpha = -0.72;
+para(2, 2).cMQ = -1.49;
+
+para(2, 2).cNR = -0.27;
+para(2, 2).cNSurface = 0.0115;
\ No newline at end of file