AEE 4263 HW 7

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Matlab code please

use this matrix num sir (AD 200107)

Question: For The Given 4-DOF Robot: A. Assign Appropriate Frames For The Denavit-... For the given 4-DOF robot: a. Assign appropriate frames for the Denavit-Hartenberg representation. b. Fill out the parameters table. c. Write an equation in terms of A matrices that shows how UTH can be calculated. U # d. a a A1 0-1 A2 1-2 Аз 2 -3 A4 3- H

For the Following question Graph all 4 : [I just need all 4 graphs and please explain and make clean solution] Position vs time Velocity vs time Acceleration vs time Force vs time [For your convenience, I have solved the numerical solutions for the problem] (Please Look at the picture since it is much cleaner) Question : A 550 kilogram mass initially at rest acted upon by a force of F(t) = 50et Newtons. What are the acceleration, speed, and displacement of the mass at  t = 4 second ? a =(50 e^t)/(550 ) [N/kg] v = ∫_0^t▒(50 e^t )dt/(550 )= v_0  +(50 e^t-50)/550=((e^t- 1))/11 x = ∫_0^t▒(e^t- 1)dt/(11 )= x_0  +(e^t- t - 1)/(11 ) a(4s)=(50*54.6)/550= 4.96[m/s^2 ] v(4s)=((e^4-1))/11= 4.87[m/s] x(4s)=((e^4- 4 - 1))/11= 4.51 [m]

I need help with my MATLAB code. There is an error in the following code. The error says my orbitaldynamics function must return a column vector. Can you help me fix it? I tried changing the semi colon to a comma in dstate_dt = [v, a_total]; but it still gives me error   mu_earth = 398600.4418; % Earth's gravitational parameter (km^3/s^2) R_earth = 6378.137; % Earth's radius (km) C_d = 0.3; % Drag coefficient (assumed) A = 0.023; % Cross-sectional area of ISS (km^2) m = 420000; % Mass of ISS (kg)       % Initial conditions: position and velocity (ISS state vector) % ISS initial state vector (km and km/s) - sample data state_ISS =[-2.1195e+03, 3.9866e+03, 5.0692e+03, -5.3489, -5.1772, 1.8324];   % Time span for 10 revolutions T_orbit = 2 * pi * sqrt((norm(state_ISS(1:3))^3) / mu_earth); time_span = [0, 10 * T_orbit];     % Step 3: Numerical integration using ODE solver options = odeset('RelTol', 1e-12, 'AbsTol', 1e-12); [t, state] = ode45(@orbitalDynamics, time_span, state_ISS,…

I could really use some assistance on part b,c,d,e,f

I need help with my MATLAB code. I am trying to plot the semi major axis and eccentricity with the following code with rungge kutta method. I have two different nnn's in the code which produce two different step sizes. So, I am wondering why both give the same exact plots for semi-major axis and eccentricity. The orbital dynamics fucntion considers J2, drag, SRP, and third body effects.   % Orbital elements obtained from TLE inc = 63.5000118; % Inclination (degs) raan = 19.999977; % RAAN (degs) ecc = 0.001; % Eccentricity argp = 120.02165342; % argument of perigee (degs) v = 0; % True anomaly (degs)   a = 6584.697271; % Semi-major axis (km)   mu = 3.986e5; period = (2*pi*sqrt((a^3)/mu));     % Calculating the state vector in eci frame [x, y, z, vx, vy, vz] = kep2cart(a, ecc, inc, raan, argp, v); Y = [x, y, z, vx, vy, vz];   options = odeset('RelTol', 1e-12, 'AbsTol', 1e-12);   nnn1 = 800; t_step1 = period/nnn1; m6 = 50; h_step6 = t_step1/m6; t_total6 = 0:h_step6:10*period;   [t6,…

I created a MATLAB code to plot the orbit of a satellite at an altitude of 300 km (LEO). I wanted to propagate my code for 2 revolutions but instead I used (180*60) because it takes 90 minutes to make a full rotation around the earth in LEO. Is that correct? Also, if I wanted to use revolutions how would I propagate it for 2 revs? Would my t = [0 4*pi].

a 02 02 R₁ ground d R3 b For the 3Bar Inverted Crank Slider shown, 1. Starting with the position equation R=Au, where A is the rotation matrix about the z axis and u is a point on the position vector, derive the velocity equations in both algebraic and matrix form to solve for w3 and b. What are the vector position and velocity equations? Your matrix equation must be in the form [ BH] 2. From part 1 derive the matrix equation to solve for @3 and b. B-[%]-[ ]] B: 3. With the values R₂ = 2 in, R3 = 10 in, R₁ = 8 in, 0₂ = 40°, 03 = 185° and w2= 10 rad/sec. solve for @3 and b. B -7--0 Do not forget units.

I am trying to code in MATLAB the equations of motion for malankovich orbitlal elements. But, I am having a problem with the B matrix. Since f matrix is 7x1 and a_d matrix has to be 3x1, the B matrix has to be 7x3. I don't know how that is possible. Can you break down the B matrix for me and let me know what size it is?

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(b): Let us first consider controlling the orbit of deputy spacecraft to rendezvous with chief spacecraft. Define x = [r] and x = x = R to represent the deputy orbital state and its target (= chief orbit) in Cartesian coordinates, respectively. The control input is thruster acceleration, u € R³, in the ECI frame. Denote the relative state by dx = x-x. Table 2 summarize the initial orbital elements. Table 2: Keplerian orbital elements at epoch (t = 0) for deputy and chief about Earth (ECI frame) Orbital element Deputy Unit Chief semi-major axis ad = 11500 ac 10000 km eccentricity inclination ed = 0.15 id=35 ee = 0.3 i = 50 degree right ascension of ascending node d = 50 Ως = 50 degree argument of periapsis true anomaly at epoch Wd Vd= 0 = 40 We = 40 degree Ve=0 degree (b.1): Derive the error dynamics of our system in ECI frame under the influence of u. (b.2): Consider a candidate Lyapunov function V = ½dr¹ K₁dr+dv₁dv, where K₁ = K, and K, > 0. Discuss the positive definiteness of V, and…

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I am trying to plot an orbit in MATLAB. There is something wrong with my code because the final values I get are incorrect. The code is shown below. The correct values are in the image.     mu = 3.986*10^5; % Earth's gravitational parameter [km^3/s^2] % Transforming orbital elements to cartesian coordinate system for LEOa_1 = 6782.99;e_1 = 0.000685539;inc_1 = 51.64;v_1 = 5;argp_1 = 30;raan_1 = 10; [x_1, y_1, z_1, vx_1, vy_1, vz_1] = kep2cart(a_1, e_1, inc_1, raan_1, ...                                            argp_1, v_1); Y_1 = [x_1, y_1, z_1, vx_1, vy_1, vz_1];   % time_span for two revolutions (depends on the orbit)t1 = [0 (180*60)];  % Setting tolerancesoptions = odeset('RelTol',1e-12,'AbsTol',1e-12); % Using ODE45 to numerically integrate for LEO[t_1, state_1] = ode45(@OrbitProp, t1, Y_1, options);       function dYdt = OrbitProp(t, Y)     mu = 3.986*10^5; % Earth's gravitational parameter [km^3/s^2]   % State Vector x = Y(1); % [km] y = Y(2); % [km] z = Y(3); % [km] vx = Y(4);…

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Given: Mass = M, Radius = R, Moment of Inertia = I, Distance =  H, Gravity = g, Initial timing = 0 Unknown =  How long will the [delta t] process take? (time) Please keep in mind that almost every variable is unknown! We are given no numbers and are instructed to generate a solution using the variables [ "R, M, g, I and H." ]  Target =  Find an equation that, if we could substitute numbers, would determine how long the process would take.  Newton's second law and torque seem to be relevant based on how the problem is written. * Also please show drawing based on what we are calculating

I want to run the SGP4 propagator for the ISS (ID = 25544) I got from spacetrack.org in MATLAB. I don't know where to get the inputs of the function. Where do I get the inFile and outFile that is mentioned in the following function. % Purpose: % This program shows how a Matlab program can call the Astrodynamic Standard libraries to propagate % satellites to the requested time using SGP4 method. % % The program reads in user's input and output files. The program generates an % ephemeris of position and velocity for each satellite read in. In addition, the program % also generates other sets of orbital elements such as osculating Keplerian elements, % mean Keplerian elements, latitude/longitude/height/pos, and nodal period/apogee/perigee/pos. % Totally, the program prints results to five different output files. % % % Usage: Sgp4Prop(inFile, outFile) % inFile : File contains TLEs and 6P-Card (which controls start, stop times and step size) % outFile : Base name for five output files   %…

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In MATLAB write out a program to integrate the equations of motion of a rigid body. The inertia matrix is given by I = [125 0 0; 0 100 0; 0 0 75] which is a diagonal, where diag operator provides a matrix with given elements placed on its diagonal. Consider three cases where the body rotates 1 rad/sec about each principal axis. Integrate the resulting motion and study the angular rates and the resulting attitude (use any attitude coordinates). For each principal axis case, assume first that a pure spin about the principal axis is performed, and then repeat the simulation where a small 0.1 rad/sec motion is present about another principal axis. Discuss the stability of each motion. The code should produce a total of 6 simulations results when it is ran.

ihài 25 Q3) Let vector A = 5i +2j, Vector B = -3i %3D -5j, and vector C = A+ B. (a) Write %3D vector C in component form. (b) Draw a coordinate system and on it show the all vectors. (c) What are the ?magnitude and direction of vector C

хоос» 0 Python 3 (p Define a python function to determine the 2nd moment of area I, under the parabola, with respect to the centroidal-axis. The parabolic curve is defined as z = ky2, where k = b/h². The centroidal axis has a given distance d with respect to z axis. Hints: Use the parallel axis theorem. Find the 2nd moment of area against z axis first and then use the formula I₂ = I₁ + d²A to calculate I, where d is a given value (no need to derive). Note that this question requires you to derive the equation first. $ 4 l) | Idle Mem: 260.77/12288.00 MB Q F4 R y 1 1 0 do 5 [ ]: # Complete the function given the variables b,h,d and return the value as "Area","Iz" and "Ic" # This is not about numerical integration, we only need the analytical solution. # Don't change the predefined content, only fill your code in the region "YOUR CODE" Mode: Command % 9 F5 0 T . Markdown v z = ky² 2 6 C F6 Validate Y & 7 8: F7 U * 8 DII F8 I 19 F9 b O ) Ln 1, Col 1 English (United States)…

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I created an orbit of the ISS in MATLAB. I want to validate my data. How and where can I get actual data of the ISS orbit? I would need the time history of the position vectors?

Saved em Set 02 - Chapter 2 Part 1 Required information of 2 Consider the following vectorial representations. Add the three vectors shown to form a resultant vector R , and report your result using polar vector representation. Take Y= 3 m. 2 m ok 3 m 30° 162 2 (Enter a positive value for the phase angle.) 4.84 m @ ces