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?

I am trying to find the gross liftoff mass of a 2 stage rocket for different values of n and plot it in MATLAB. I found the total liftoff mass for stage 1 and stage 2. Is the gross liftoff mass equal to m_i1 in the code or do you have to add m_i1 and m_i2? Also, the liftoff seems to be negative for some intial values of n? That is not feasible, right?   % Constants delta_V_ideal = 9800; % m/s f_inert1 = 0.15; f_inert2 = 0.25; Isp_1 = 325; % s Isp_2 = 380; % s g0 = 9.81; % m/s^2 m_PL = 750; n = 0.3:0.01:0.7;   % Initializing variables delta_v1 = zeros(size(n)); delta_v2 = zeros(size(n)); MR_2 = zeros(size(n)); m_prop2 = zeros(size(n)); m_inert2 = zeros(size(n)); m_i2 = zeros(size(n)); MR_1 = zeros(size(n)); m_prop1 = zeros(size(n)); m_inert1 = zeros(size(n)); m_i1 = zeros(size(n));   for i = 1:length(n) delta_v1(i) = n(i)*delta_V_ideal; delta_v2(i) = (1-n(i))*delta_V_ideal;   MR_2(i) = exp(delta_v2(i)/(g0*Isp_2)); m_prop2(i) = (m_PL*(MR_2(i)-1)*(1-f_inert2))/(1-f_inert2*MR_2(i));…

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);…

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,…