Question 3 Create a mask for each frame of this video that isolates the nearly-white flame exiting the rocket. Do this by converting each frame to grayscale. Then label all pixels with intensities less than or equal to 245 as false, and all pixels with intensities greater than 245 as true.
What is the ratio of area of the flame between the last frame and the first? (Use the number of true pixels for area.)
% Load the video file ‘shuttle.avi’
videoFile = ‘shuttle.avi’;
v = VideoReader(videoFile);

% Read the first frame
firstFrame = readFrame(v);

% Move to the last frame
while hasFrame(v)
lastFrame = readFrame(v);
end

% Convert both frames to grayscale
firstGray = rgb2gray(firstFrame);
lastGray = rgb2gray(lastFrame);

% Create masks where pixel intensities > 245 are true, the rest are false
firstMask = firstGray > 245;
lastMask = lastGray > 245;

% Calculate the flame area (number of ‘true’ pixels) for both frames
areaFirst = sum(firstMask(:));
areaLast = sum(lastMask(:));

% Calculate the ratio of the flame area between the last and the first frame
if areaFirst == 0
error(‘No flame detected in the first frame.’);
else
ratio = areaLast / areaFirst;
end

% Display the results
fprintf(‘Area of the flame in the first frame: %d pixelsn’, areaFirst);
fprintf(‘Area of the flame in the last frame: %d pixelsn’, areaLast);
fprintf(‘Ratio of flame area (last frame / first frame): %.2fn’, ratio);

% Optional: Display the original frames and the masks for visual verification
figure;
subplot(2,2,1), imshow(firstFrame), title(‘First Frame’);
subplot(2,2,2), imshow(lastFrame), title(‘Last Frame’);
subplot(2,2,3), imshow(firstMask), title(‘First Frame Mask’);
subplot(2,2,4), imshow(lastMask), title(‘Last Frame Mask’);Question 3 Create a mask for each frame of this video that isolates the nearly-white flame exiting the rocket. Do this by converting each frame to grayscale. Then label all pixels with intensities less than or equal to 245 as false, and all pixels with intensities greater than 245 as true.
What is the ratio of area of the flame between the last frame and the first? (Use the number of true pixels for area.)
% Load the video file ‘shuttle.avi’
videoFile = ‘shuttle.avi’;
v = VideoReader(videoFile);

% Read the first frame
firstFrame = readFrame(v);

% Move to the last frame
while hasFrame(v)
lastFrame = readFrame(v);
end

% Convert both frames to grayscale
firstGray = rgb2gray(firstFrame);
lastGray = rgb2gray(lastFrame);

% Create masks where pixel intensities > 245 are true, the rest are false
firstMask = firstGray > 245;
lastMask = lastGray > 245;

% Calculate the flame area (number of ‘true’ pixels) for both frames
areaFirst = sum(firstMask(:));
areaLast = sum(lastMask(:));

% Calculate the ratio of the flame area between the last and the first frame
if areaFirst == 0
error(‘No flame detected in the first frame.’);
else
ratio = areaLast / areaFirst;
end

% Display the results
fprintf(‘Area of the flame in the first frame: %d pixelsn’, areaFirst);
fprintf(‘Area of the flame in the last frame: %d pixelsn’, areaLast);
fprintf(‘Ratio of flame area (last frame / first frame): %.2fn’, ratio);

% Optional: Display the original frames and the masks for visual verification
figure;
subplot(2,2,1), imshow(firstFrame), title(‘First Frame’);
subplot(2,2,2), imshow(lastFrame), title(‘Last Frame’);
subplot(2,2,3), imshow(firstMask), title(‘First Frame Mask’);
subplot(2,2,4), imshow(lastMask), title(‘Last Frame Mask’); Question 3 Create a mask for each frame of this video that isolates the nearly-white flame exiting the rocket. Do this by converting each frame to grayscale. Then label all pixels with intensities less than or equal to 245 as false, and all pixels with intensities greater than 245 as true.
What is the ratio of area of the flame between the last frame and the first? (Use the number of true pixels for area.)
% Load the video file ‘shuttle.avi’
videoFile = ‘shuttle.avi’;
v = VideoReader(videoFile);

% Read the first frame
firstFrame = readFrame(v);

% Move to the last frame
while hasFrame(v)
lastFrame = readFrame(v);
end

% Convert both frames to grayscale
firstGray = rgb2gray(firstFrame);
lastGray = rgb2gray(lastFrame);

% Create masks where pixel intensities > 245 are true, the rest are false
firstMask = firstGray > 245;
lastMask = lastGray > 245;

% Calculate the flame area (number of ‘true’ pixels) for both frames
areaFirst = sum(firstMask(:));
areaLast = sum(lastMask(:));

% Calculate the ratio of the flame area between the last and the first frame
if areaFirst == 0
error(‘No flame detected in the first frame.’);
else
ratio = areaLast / areaFirst;
end

% Display the results
fprintf(‘Area of the flame in the first frame: %d pixelsn’, areaFirst);
fprintf(‘Area of the flame in the last frame: %d pixelsn’, areaLast);
fprintf(‘Ratio of flame area (last frame / first frame): %.2fn’, ratio);

% Optional: Display the original frames and the masks for visual verification
figure;
subplot(2,2,1), imshow(firstFrame), title(‘First Frame’);
subplot(2,2,2), imshow(lastFrame), title(‘Last Frame’);
subplot(2,2,3), imshow(firstMask), title(‘First Frame Mask’);
subplot(2,2,4), imshow(lastMask), title(‘Last Frame Mask’); image processing MATLAB Answers — New Questions

​

For an Aerodynamics couse, we need to create a code that computes lifting surfaces (wing) aerodynamics. I cannot get the graph to display a nice CL vs AoA curve, but only a single value. I don’t know how I can accomplish this. As I am a novice to MAtlab, I would appreciate some pointers or hints. Thank you.
clc;
% Default airfoil or not…
Default_AF = 0;

% A/F Characteristics:
% Lift coefficient [1/rad]:
Cl_alpha = 5.79;
Cl_alphaInRadians = deg2rad(Cl_alpha);

% Sectional lift curve slope [deg]:
alpha_0 = -2.0;

% Pitching moment [1/rad]:
Cm_alpha = 0;
Cm_alphaInRadians = deg2rad(Cm_alpha);

% 2D Pitching moment:
Cm_0 = -0.025;

% Coefficient in series expansion of circulation distribution [deg]:
alpha_nl = 8;

% Wing planform Geometry [inches]:
% [root-LE,root-TE,tip-LE,tip-TE]
x1 = 0;
y1 = 0;
x2 = 200;
y2 = 0;
x3 = 0;
y3 = 1000;
x4 = 200;
y4 = 1000;

% Wing twist angle [deg]:
theta_t = 0;

% Number of spanwise coefficients:
N = 100;

% Flight Conditions:
% Free-stream Velocity [ft/sec]
V = 100;

% Density [slugs/ft^3]:
rho = 0.002344;

% Viscous Force [lb.s/ft^2]
mu = 0.0000003737;

% Angle of Attack [deg]
alpha = 5;

% Wing span:
b = sqrt((x2-x1)^2 + (y2-y1)^2);

% Wing chord:
c = sqrt((x3-x1)^2 + (y3-y1)^2);

% Wing area:
S = b*c;

% Wing semi-span:
s = b/2;

% Aspect Ratio:
AR = b^2./S;

% Applying segment length of the wings:
alpha_segment = zeros(1,N);
theta_segment = zeros(1,N);

% Taper Ratio and its range within the wing (which is 0 to 1):
for lambda_r = 0:0.25:1

% Root Chord:
c_root = (2*S) / ((1+lambda_r)*b);

% Tip Chord:
c_tip = ((2*S)/((1+lambda_r)*b)) * (1-((2*(1-lambda_r))/b)*(b/2));

% Mean Aerodynamic Chord:
MAC = (2/3)*(c_root+c_tip-(c_root*c_tip)/(c_root+c_tip));

% Taper ratio:
lambda = c_tip / c_root;

% Providing segment limitation within range:
alpha_segment(1) = alpha;
theta_segment(1) = pi/2;

% Completing the Lifting Line Theory to determine the
% matrix value of the segments:
for N = 2:N

c(N) = c_root+(c_tip-c_root)/N.*N;
alpha_segment(N) = alpha+(theta_t)/N.*N;
theta_segment = pi/(2*N):pi/(2*N):pi/2;

end

% Lifting line theory construction:
xb = 4*b./(pi*c);

% Completing the [B] matrix:
for i = 1:N

for j = 1:N

B(i,j) = (sin(j.*theta_segment(i))).*(xb(i)+j/(sin(theta_segment(i))));

end

slope(i) = (alpha_segment(i)-alpha_0)*(pi/180);

end

A = Btranspose(slope);

end

% Leading edge sweep angle
for fai = 1:N

delta_LE(fai) = fai*(A(fai)/A(1))^2; %#ok<*SAGROW>

end

% Section Lift Coefficient of Airfoil
cl = 0.5*cos(pi/b);

% Wing Lift Coefficient:
CL = pi*AR*A(1);

% Span Efficiency:
delta = sum(delta_LE);
CD_0 = 1/(1+delta);

% Induced drag coefficient:
CD_i = CL.^2 / (pi*CD_0*AR);

% Speed of sound (assuming 20 degree dry air) [ft/sec]:
C = 1125.33;

% Mach Number:
M = V / C;

% Dynamic Pressure:
q = (0.5*rho*V^2);

% Pitching Moment Coefficient:
CM = M / q*S*c;

% Subsonic Lift:
alpha_inf = 2*pi*cos(delta_LE);

% 3D Lift Curve Slope of airfoil:
CL_alpha = 2*pi*A;

% Geometric Angle of Attack of Wing:
alpha = (alpha_inf)/(1+(alpha_inf)/(pi*AR));

%———————————————————-

% Wing lift coefficient (CL) vs Angle of Attack (alpha):
figure(1);
plot(alpha,CL,’-o’)
grid on
title(‘Wing lift coefficient (CL) vs Angle of Attack (alpha)’)
ylabel(‘Wing Lift Coefficient, CL’)
xlabel(‘Angle of Attack, alpha [deg]’)

% Pitching moment coefficient(CM) vs Angle of Attack (alpha):
figure(2);
plot(alpha,CM,’-o’)
grid on
title(‘Pitching moment coefficient(CM) vs Angle of Attack (alpha)’)
ylabel(‘Pitching Moment Coefficient, CM’)
xlabel(‘Angle of Attack, alpha [deg]’)

% Wing lift coefficient (CL) vs Induced drag coefficient (CD_i):
figure(3);
plot(CD_i,CL,’-o’)
grid on
title(‘Wing lift coefficient (CL) vs Induced drag coefficient (CD_i)’)
ylabel(‘Wing Lift Coefficient, CL’)
xlabel(‘Induced Drag Coefficient, CD_i’)For an Aerodynamics couse, we need to create a code that computes lifting surfaces (wing) aerodynamics. I cannot get the graph to display a nice CL vs AoA curve, but only a single value. I don’t know how I can accomplish this. As I am a novice to MAtlab, I would appreciate some pointers or hints. Thank you.
clc;
% Default airfoil or not…
Default_AF = 0;

% A/F Characteristics:
% Lift coefficient [1/rad]:
Cl_alpha = 5.79;
Cl_alphaInRadians = deg2rad(Cl_alpha);

% Sectional lift curve slope [deg]:
alpha_0 = -2.0;

% Pitching moment [1/rad]:
Cm_alpha = 0;
Cm_alphaInRadians = deg2rad(Cm_alpha);

% 2D Pitching moment:
Cm_0 = -0.025;

% Coefficient in series expansion of circulation distribution [deg]:
alpha_nl = 8;

% Wing planform Geometry [inches]:
% [root-LE,root-TE,tip-LE,tip-TE]
x1 = 0;
y1 = 0;
x2 = 200;
y2 = 0;
x3 = 0;
y3 = 1000;
x4 = 200;
y4 = 1000;

% Wing twist angle [deg]:
theta_t = 0;

% Number of spanwise coefficients:
N = 100;

% Flight Conditions:
% Free-stream Velocity [ft/sec]
V = 100;

% Density [slugs/ft^3]:
rho = 0.002344;

% Viscous Force [lb.s/ft^2]
mu = 0.0000003737;

% Angle of Attack [deg]
alpha = 5;

% Wing span:
b = sqrt((x2-x1)^2 + (y2-y1)^2);

% Wing chord:
c = sqrt((x3-x1)^2 + (y3-y1)^2);

% Wing area:
S = b*c;

% Wing semi-span:
s = b/2;

% Aspect Ratio:
AR = b^2./S;

% Applying segment length of the wings:
alpha_segment = zeros(1,N);
theta_segment = zeros(1,N);

% Taper Ratio and its range within the wing (which is 0 to 1):
for lambda_r = 0:0.25:1

% Root Chord:
c_root = (2*S) / ((1+lambda_r)*b);

% Tip Chord:
c_tip = ((2*S)/((1+lambda_r)*b)) * (1-((2*(1-lambda_r))/b)*(b/2));

% Mean Aerodynamic Chord:
MAC = (2/3)*(c_root+c_tip-(c_root*c_tip)/(c_root+c_tip));

% Taper ratio:
lambda = c_tip / c_root;

% Providing segment limitation within range:
alpha_segment(1) = alpha;
theta_segment(1) = pi/2;

% Completing the Lifting Line Theory to determine the
% matrix value of the segments:
for N = 2:N

c(N) = c_root+(c_tip-c_root)/N.*N;
alpha_segment(N) = alpha+(theta_t)/N.*N;
theta_segment = pi/(2*N):pi/(2*N):pi/2;

end

% Lifting line theory construction:
xb = 4*b./(pi*c);

% Completing the [B] matrix:
for i = 1:N

for j = 1:N

B(i,j) = (sin(j.*theta_segment(i))).*(xb(i)+j/(sin(theta_segment(i))));

end

slope(i) = (alpha_segment(i)-alpha_0)*(pi/180);

end

A = Btranspose(slope);

end

% Leading edge sweep angle
for fai = 1:N

delta_LE(fai) = fai*(A(fai)/A(1))^2; %#ok<*SAGROW>

end

% Section Lift Coefficient of Airfoil
cl = 0.5*cos(pi/b);

% Wing Lift Coefficient:
CL = pi*AR*A(1);

% Span Efficiency:
delta = sum(delta_LE);
CD_0 = 1/(1+delta);

% Induced drag coefficient:
CD_i = CL.^2 / (pi*CD_0*AR);

% Speed of sound (assuming 20 degree dry air) [ft/sec]:
C = 1125.33;

% Mach Number:
M = V / C;

% Dynamic Pressure:
q = (0.5*rho*V^2);

% Pitching Moment Coefficient:
CM = M / q*S*c;

% Subsonic Lift:
alpha_inf = 2*pi*cos(delta_LE);

% 3D Lift Curve Slope of airfoil:
CL_alpha = 2*pi*A;

% Geometric Angle of Attack of Wing:
alpha = (alpha_inf)/(1+(alpha_inf)/(pi*AR));

%———————————————————-

% Wing lift coefficient (CL) vs Angle of Attack (alpha):
figure(1);
plot(alpha,CL,’-o’)
grid on
title(‘Wing lift coefficient (CL) vs Angle of Attack (alpha)’)
ylabel(‘Wing Lift Coefficient, CL’)
xlabel(‘Angle of Attack, alpha [deg]’)

% Pitching moment coefficient(CM) vs Angle of Attack (alpha):
figure(2);
plot(alpha,CM,’-o’)
grid on
title(‘Pitching moment coefficient(CM) vs Angle of Attack (alpha)’)
ylabel(‘Pitching Moment Coefficient, CM’)
xlabel(‘Angle of Attack, alpha [deg]’)

% Wing lift coefficient (CL) vs Induced drag coefficient (CD_i):
figure(3);
plot(CD_i,CL,’-o’)
grid on
title(‘Wing lift coefficient (CL) vs Induced drag coefficient (CD_i)’)
ylabel(‘Wing Lift Coefficient, CL’)
xlabel(‘Induced Drag Coefficient, CD_i’) For an Aerodynamics couse, we need to create a code that computes lifting surfaces (wing) aerodynamics. I cannot get the graph to display a nice CL vs AoA curve, but only a single value. I don’t know how I can accomplish this. As I am a novice to MAtlab, I would appreciate some pointers or hints. Thank you.
clc;
% Default airfoil or not…
Default_AF = 0;

% A/F Characteristics:
% Lift coefficient [1/rad]:
Cl_alpha = 5.79;
Cl_alphaInRadians = deg2rad(Cl_alpha);

% Sectional lift curve slope [deg]:
alpha_0 = -2.0;

% Pitching moment [1/rad]:
Cm_alpha = 0;
Cm_alphaInRadians = deg2rad(Cm_alpha);

% 2D Pitching moment:
Cm_0 = -0.025;

% Coefficient in series expansion of circulation distribution [deg]:
alpha_nl = 8;

% Wing planform Geometry [inches]:
% [root-LE,root-TE,tip-LE,tip-TE]
x1 = 0;
y1 = 0;
x2 = 200;
y2 = 0;
x3 = 0;
y3 = 1000;
x4 = 200;
y4 = 1000;

% Wing twist angle [deg]:
theta_t = 0;

% Number of spanwise coefficients:
N = 100;

% Flight Conditions:
% Free-stream Velocity [ft/sec]
V = 100;

% Density [slugs/ft^3]:
rho = 0.002344;

% Viscous Force [lb.s/ft^2]
mu = 0.0000003737;

% Angle of Attack [deg]
alpha = 5;

% Wing span:
b = sqrt((x2-x1)^2 + (y2-y1)^2);

% Wing chord:
c = sqrt((x3-x1)^2 + (y3-y1)^2);

% Wing area:
S = b*c;

% Wing semi-span:
s = b/2;

% Aspect Ratio:
AR = b^2./S;

% Applying segment length of the wings:
alpha_segment = zeros(1,N);
theta_segment = zeros(1,N);

% Taper Ratio and its range within the wing (which is 0 to 1):
for lambda_r = 0:0.25:1

% Root Chord:
c_root = (2*S) / ((1+lambda_r)*b);

% Tip Chord:
c_tip = ((2*S)/((1+lambda_r)*b)) * (1-((2*(1-lambda_r))/b)*(b/2));

% Mean Aerodynamic Chord:
MAC = (2/3)*(c_root+c_tip-(c_root*c_tip)/(c_root+c_tip));

% Taper ratio:
lambda = c_tip / c_root;

% Providing segment limitation within range:
alpha_segment(1) = alpha;
theta_segment(1) = pi/2;

% Completing the Lifting Line Theory to determine the
% matrix value of the segments:
for N = 2:N

c(N) = c_root+(c_tip-c_root)/N.*N;
alpha_segment(N) = alpha+(theta_t)/N.*N;
theta_segment = pi/(2*N):pi/(2*N):pi/2;

end

% Lifting line theory construction:
xb = 4*b./(pi*c);

% Completing the [B] matrix:
for i = 1:N

for j = 1:N

B(i,j) = (sin(j.*theta_segment(i))).*(xb(i)+j/(sin(theta_segment(i))));

end

slope(i) = (alpha_segment(i)-alpha_0)*(pi/180);

end

A = Btranspose(slope);

end

% Leading edge sweep angle
for fai = 1:N

delta_LE(fai) = fai*(A(fai)/A(1))^2; %#ok<*SAGROW>

end

% Section Lift Coefficient of Airfoil
cl = 0.5*cos(pi/b);

% Wing Lift Coefficient:
CL = pi*AR*A(1);

% Span Efficiency:
delta = sum(delta_LE);
CD_0 = 1/(1+delta);

% Induced drag coefficient:
CD_i = CL.^2 / (pi*CD_0*AR);

% Speed of sound (assuming 20 degree dry air) [ft/sec]:
C = 1125.33;

% Mach Number:
M = V / C;

% Dynamic Pressure:
q = (0.5*rho*V^2);

% Pitching Moment Coefficient:
CM = M / q*S*c;

% Subsonic Lift:
alpha_inf = 2*pi*cos(delta_LE);

% 3D Lift Curve Slope of airfoil:
CL_alpha = 2*pi*A;

% Geometric Angle of Attack of Wing:
alpha = (alpha_inf)/(1+(alpha_inf)/(pi*AR));

%———————————————————-

% Wing lift coefficient (CL) vs Angle of Attack (alpha):
figure(1);
plot(alpha,CL,’-o’)
grid on
title(‘Wing lift coefficient (CL) vs Angle of Attack (alpha)’)
ylabel(‘Wing Lift Coefficient, CL’)
xlabel(‘Angle of Attack, alpha [deg]’)

% Pitching moment coefficient(CM) vs Angle of Attack (alpha):
figure(2);
plot(alpha,CM,’-o’)
grid on
title(‘Pitching moment coefficient(CM) vs Angle of Attack (alpha)’)
ylabel(‘Pitching Moment Coefficient, CM’)
xlabel(‘Angle of Attack, alpha [deg]’)

% Wing lift coefficient (CL) vs Induced drag coefficient (CD_i):
figure(3);
plot(CD_i,CL,’-o’)
grid on
title(‘Wing lift coefficient (CL) vs Induced drag coefficient (CD_i)’)
ylabel(‘Wing Lift Coefficient, CL’)
xlabel(‘Induced Drag Coefficient, CD_i’) aerodynamics, lifting surfaces MATLAB Answers — New Questions

​

I was able to install MATLAB, however I am not able to sign into my MathWorks Account with my university credentials when trying to launch MATLAB for the first time. 

When I input my email address it says I must log in through my school’s portal, but nothing pops up except a blank white page in my shortcut.
 I was able to install MATLAB, however I am not able to sign into my MathWorks Account with my university credentials when trying to launch MATLAB for the first time. 

When I input my email address it says I must log in through my school’s portal, but nothing pops up except a blank white page in my shortcut.
  I was able to install MATLAB, however I am not able to sign into my MathWorks Account with my university credentials when trying to launch MATLAB for the first time. 

When I input my email address it says I must log in through my school’s portal, but nothing pops up except a blank white page in my shortcut.
   MATLAB Answers — New Questions

​

I am searching the Dynamic Voltage Restorer (DVR) block in library browser but I am not getting the block.please provide me the steps to find out the DVR block from library browser in matlab simulink.I am searching the Dynamic Voltage Restorer (DVR) block in library browser but I am not getting the block.please provide me the steps to find out the DVR block from library browser in matlab simulink. I am searching the Dynamic Voltage Restorer (DVR) block in library browser but I am not getting the block.please provide me the steps to find out the DVR block from library browser in matlab simulink. dvr MATLAB Answers — New Questions

​

I need to include the Dynamic of the robot I’m using, imported from an URDF file, to a Extended Kalman Filter as the State Transition Function in order to estimate the states of the system. I tried to put all the simscape model inside a Simulink Function, but it brings Error1 saying I can’t put the solver configuration block (from the robot’s body tree) inside a subsystem. I came back and also tried taking the solver configuration block out of the Simulink Function and brin it in using a Connection Port, but it sends now Error2 saying all the elements of the body tree must be in the same level.
Error1: "’brazo_2dof/Simulink Function/Solver Configuration’ has been placed in the function call subsystem ‘brazo_2dof/Simulink Function’ but does not support this configuration. Move the block out of the function call subsystem."
Error2: "Connections of the Physical Network cross nonvirtual boundaries. If any blocks in a Physical Network are within a nonvirtual subsystem, then all blocks of that network must be in the same nonvirtual subsystem. Either move affected blocks out of the nonvirtual subsystem, move network blocks into exactly one nonvirtual subsystem, or reconfigure affected nonvirtual subsystems to be virtual subsystems."
How else could I include the dynamics of my system, imported from the URDF, as the StateTransitionFunction?I need to include the Dynamic of the robot I’m using, imported from an URDF file, to a Extended Kalman Filter as the State Transition Function in order to estimate the states of the system. I tried to put all the simscape model inside a Simulink Function, but it brings Error1 saying I can’t put the solver configuration block (from the robot’s body tree) inside a subsystem. I came back and also tried taking the solver configuration block out of the Simulink Function and brin it in using a Connection Port, but it sends now Error2 saying all the elements of the body tree must be in the same level.
Error1: "’brazo_2dof/Simulink Function/Solver Configuration’ has been placed in the function call subsystem ‘brazo_2dof/Simulink Function’ but does not support this configuration. Move the block out of the function call subsystem."
Error2: "Connections of the Physical Network cross nonvirtual boundaries. If any blocks in a Physical Network are within a nonvirtual subsystem, then all blocks of that network must be in the same nonvirtual subsystem. Either move affected blocks out of the nonvirtual subsystem, move network blocks into exactly one nonvirtual subsystem, or reconfigure affected nonvirtual subsystems to be virtual subsystems."
How else could I include the dynamics of my system, imported from the URDF, as the StateTransitionFunction? I need to include the Dynamic of the robot I’m using, imported from an URDF file, to a Extended Kalman Filter as the State Transition Function in order to estimate the states of the system. I tried to put all the simscape model inside a Simulink Function, but it brings Error1 saying I can’t put the solver configuration block (from the robot’s body tree) inside a subsystem. I came back and also tried taking the solver configuration block out of the Simulink Function and brin it in using a Connection Port, but it sends now Error2 saying all the elements of the body tree must be in the same level.
Error1: "’brazo_2dof/Simulink Function/Solver Configuration’ has been placed in the function call subsystem ‘brazo_2dof/Simulink Function’ but does not support this configuration. Move the block out of the function call subsystem."
Error2: "Connections of the Physical Network cross nonvirtual boundaries. If any blocks in a Physical Network are within a nonvirtual subsystem, then all blocks of that network must be in the same nonvirtual subsystem. Either move affected blocks out of the nonvirtual subsystem, move network blocks into exactly one nonvirtual subsystem, or reconfigure affected nonvirtual subsystems to be virtual subsystems."
How else could I include the dynamics of my system, imported from the URDF, as the StateTransitionFunction? simscape, extended kalman filter, state transition function, multibody, estimate states MATLAB Answers — New Questions

​

When I try to open Matlab Desktop 2024a it asks me to sign in and when i type my email it says "Sign into your organization’s portal to proceed." but there is not Organization portal popping up for me to sign inWhen I try to open Matlab Desktop 2024a it asks me to sign in and when i type my email it says "Sign into your organization’s portal to proceed." but there is not Organization portal popping up for me to sign in When I try to open Matlab Desktop 2024a it asks me to sign in and when i type my email it says "Sign into your organization’s portal to proceed." but there is not Organization portal popping up for me to sign in login, sign-in MATLAB Answers — New Questions

​

Hello,
How to create tiledlayout grid in vertical order?
Let’s say I want to create a tiledlayout(3,2), matlab default order is
1 2
3 4
5 6
But I want the lay out to be:
1 4
2 5
3 6
How do I do that?
Thanks,
PingHello,
How to create tiledlayout grid in vertical order?
Let’s say I want to create a tiledlayout(3,2), matlab default order is
1 2
3 4
5 6
But I want the lay out to be:
1 4
2 5
3 6
How do I do that?
Thanks,
Ping Hello,
How to create tiledlayout grid in vertical order?
Let’s say I want to create a tiledlayout(3,2), matlab default order is
1 2
3 4
5 6
But I want the lay out to be:
1 4
2 5
3 6
How do I do that?
Thanks,
Ping tiledlayout, order MATLAB Answers — New Questions

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I have a problem with calculation of the complex integral

The integrand has a pole in z=1 and z=-1 lying on the unit circle.
My code are as following:
t = @(theta) exp(1i*theta);
f = @(t) (t^2+3*t+2*1i)/(t+4)/(t-1)/(t+1);
integrand = @(theta) f(t(theta))*1i*exp(1i*theta);
integral_result = integral(integrand, 0, 2*pi, ‘ArrayValued’, true, ‘AbsTol’, 1e-12, ‘RelTol’, 1e-12);
result = (1/(2*pi*1i)) * integral_result;
I_num=result;
disp([‘numerical solution=’,num2str(I_num)])
How do i fix it…I have a problem with calculation of the complex integral

The integrand has a pole in z=1 and z=-1 lying on the unit circle.
My code are as following:
t = @(theta) exp(1i*theta);
f = @(t) (t^2+3*t+2*1i)/(t+4)/(t-1)/(t+1);
integrand = @(theta) f(t(theta))*1i*exp(1i*theta);
integral_result = integral(integrand, 0, 2*pi, ‘ArrayValued’, true, ‘AbsTol’, 1e-12, ‘RelTol’, 1e-12);
result = (1/(2*pi*1i)) * integral_result;
I_num=result;
disp([‘numerical solution=’,num2str(I_num)])
How do i fix it… I have a problem with calculation of the complex integral

The integrand has a pole in z=1 and z=-1 lying on the unit circle.
My code are as following:
t = @(theta) exp(1i*theta);
f = @(t) (t^2+3*t+2*1i)/(t+4)/(t-1)/(t+1);
integrand = @(theta) f(t(theta))*1i*exp(1i*theta);
integral_result = integral(integrand, 0, 2*pi, ‘ArrayValued’, true, ‘AbsTol’, 1e-12, ‘RelTol’, 1e-12);
result = (1/(2*pi*1i)) * integral_result;
I_num=result;
disp([‘numerical solution=’,num2str(I_num)])
How do i fix it… numerical integration, cauthy, matlab code MATLAB Answers — New Questions

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#include "rtw_continuous.h"
#include "rtw_solver.h"
these header files are not generated in MATLAB.
Is there any options to know that.#include "rtw_continuous.h"
#include "rtw_solver.h"
these header files are not generated in MATLAB.
Is there any options to know that. #include "rtw_continuous.h"
#include "rtw_solver.h"
these header files are not generated in MATLAB.
Is there any options to know that. embedded coder MATLAB Answers — New Questions

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We have been running MATLAB R2023B Update 7 with Simulink models using S-Functions built with VS 2022 17.9.0. When we updated to VS 2022 17.10.4 we starting having MATLAB crashes as soon as Simulink starts executing models with these MEX functions. Is this a known issue, and do we know if moving to Update 8 or 9 will fix it?We have been running MATLAB R2023B Update 7 with Simulink models using S-Functions built with VS 2022 17.9.0. When we updated to VS 2022 17.10.4 we starting having MATLAB crashes as soon as Simulink starts executing models with these MEX functions. Is this a known issue, and do we know if moving to Update 8 or 9 will fix it? We have been running MATLAB R2023B Update 7 with Simulink models using S-Functions built with VS 2022 17.9.0. When we updated to VS 2022 17.10.4 we starting having MATLAB crashes as soon as Simulink starts executing models with these MEX functions. Is this a known issue, and do we know if moving to Update 8 or 9 will fix it? mex, simulink MATLAB Answers — New Questions

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