I am having trouble with the folloiwng MATLAB code. I am getting an error that says "unrecognized function or variable 'numericalPropogatorOptions". I have the aerospace toolbox and the aerospace blockset added. what add on do I have to download to use that function. How do I make this code work?     % Define Keplerian Elements a = 29599.8; e = 0.0001; i = 0.9774; Omega = 1.3549; w = 0; M = 0.2645; [RECI, VECI] = Kepler2RV(a, e, i, Omega, w, M); initialState = [RECI * 1e3; VECI * 1e3]; % Initial position (m) and velocity (m/s)   % Define constants mu = 3.986004418e14; % Gravitational constant (m^3/s^2) earthRadius = 6378.1363 * 1e3; % Earth radius in meters j2 = 1.08263e-3; % J2 perturbation coefficient   % Define propagator options propOptions = numericalPropagatorOptions('CentralBody', 'Earth', ... 'GravitationalParameter', mu, ... 'InitialState', initialState, ... 'OutputTimeStep', 300); % Output every 300 seconds   % Add perturbations addGravityModel(propOptions, 'Degree', 2, 'Order', 0); % J2 effect addDragModel(propOptions, 'AtmosphereModel', 'Exponential', ... 'ReferenceArea', 1.0, 'DragCoefficient', 2.2, ... 'Mass', 500); % Atmospheric drag addSolarRadiationPressureModel(propOptions, 'ReferenceArea', 1.0, ... 'Reflectivity', 1.3, 'Mass', 500); % SRP addThirdBodyGravity(propOptions, 'Moon'); % Third-body effect from the Moon addThirdBodyGravity(propOptions, 'Sun'); % Third-body effect from the Sun   % Propagate the orbit endTime = 300 * 300; % 300 steps, each of 300 seconds [t, state] = numericalPropagator(endTime, propOptions);   % Extract position and velocity for plotting position = state(:, 1:3); % Position in ECI (meters) velocity = state(:, 4:6); % Velocity in ECI (meters/second)   % Plot the orbit plot3(position(:, 1), position(:, 2), position(:, 3)); xlabel('X (m)'); ylabel('Y (m)'); zlabel('Z (m)'); title('Simulated Orbit with Perturbations in ECI'); grid on;       function [RECI, VECI] = Kepler2RV(a, e, i, Omega, w, M) mu = 398600.441799999971; % Earth's Standard Gravitational Parameter (GM) Eps = 2.22044604925031e-8; % Machine epsilon E = SolveKepler(M,e,Eps); % Eccentric Anomaly from Mean Anomaly M = 2*atan(sqrt((1+e)/(1-e))*tan(E/2)); % True Anomaly obtained from Eccentric Anomaly p = a*(1 - e^2); % Semiparameter (km) % Position Coordinates in Perifocal Coordinate System x = (p*cos(M)) / (1 + e*cos(M)); % x-coordinate (km) y = (p*sin(M)) / (1 + e*cos(M)); % y-coordinate (km) z = 0; % z-coordinate (km) vx = -(mu/p)^(1/2) * sin(M); % velocity in x (km/s) vy = (mu/p)^(1/2) * (e + cos(M)); % velocity in y (km/s) vz = 0; % velocity in z (km/s) % Transformation Matrix (3 Rotations) t_rot = [cos(Omega)*cos(w)-sin(Omega)*sin(w)*cos(i) ... (-1)*cos(Omega)*sin(w)-sin(Omega)*cos(w)*cos(i) ... sin(Omega)*sin(i); ... sin(Omega)*cos(w)+cos(Omega)*sin(w)*cos(i) ... (-1)*sin(Omega)*sin(w)+cos(Omega)*cos(w)*cos(i) ... (-1)*cos(Omega)*sin(i); ... sin(w)*sin(i) cos(w)*sin(i) cos(i)]; % Transformation Perifocal xyz to ECI RECI = t_rot*[x y z]'; VECI = t_rot*[vx vy vz]'; end       function E = SolveKepler(M,e,Eps) % Iterative Solution of Kepler's equation (M = E-e*sin(E)) %========================================================================== % Inputs: M = Mean Anomaly (rad) % e = Eccentricity % Eps = Machine epsilon % Output: E = Eccentric Anomaly (rad) %========================================================================== E0 = M; E1 = E0 - (E0 - e*sin(E0) - M)/(1 -e*cos(E0)); while (abs(E1-E0) > Eps) E0 = E1; E1 = E0 - (E0 - e*sin(E0) - M)/(1 - e*cos(E0)); end E = E1; end

please write a matlab code

How do I input this code for this MATLAB problem? Thanks!

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   %…