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Two-line element set (TLE)

 

clear all;
mu = 398600; %  Standard gravitational parameter for the earth
% TLE file name
fname = 'tle.txt';
% Open the TLE file and read TLE elements
fid = fopen(fname, 'rb');
L1c = fscanf(fid,'%24c%',1);
L2c = fscanf(fid,'%71c%',1);
L3c = fscanf(fid,'%71c%',1);
fprintf(L1c);
fprintf(L2c);
fprintf([L3c,'\n']);
fclose(fid);
% Open the TLE file and read TLE elements
fid = fopen(fname, 'rb');
L1 = fscanf(fid,'%24c%*s',1);
L2 = fscanf(fid,'%d%6d%*c%5d%*3c%*2f%f%f%5d%*c%*d%5d%*c%*d%d%5d',[1,9]);
L3 = fscanf(fid,'%d%6d%f%f%f%f%f%f%f',[1,8]);
fclose(fid);

epoch = L2(1,4)*24*3600;        % Epoch Date and Julian Date Fraction
Db    = L2(1,5);                % Ballistic Coefficient
inc   = L3(1,3);                % Inclination [deg]
RAAN  = L3(1,4);                % Right Ascension of the Ascending Node [deg]
e     = L3(1,5)/1e7;            % Eccentricity
w     = L3(1,6);                % Argument of periapsis [deg]
M     = L3(1,7);                % Mean anomaly [deg]
n     = L3(1,8);                % Mean motion [Revs per day]

% Orbital elements

a = (mu/(n*2*pi/(24*3600))^2)^(1/3);     % Semi-major axis [km]

% Calculate the eccentric anomaly using Mean anomaly
err = 1e-10;            %Calculation Error
E0 = M; t =1;
itt = 0;
while(t)
       E =  M + e*sind(E0);
      if ( abs(E - E0) < err)
          t = 0;
      end
      E0 = E;
      itt = itt+1;
end

% Six orbital elements
OE = [a e inc RAAN w E];
fprintf('\n a [km]   e      inc [deg]  RAAN [deg]  w[deg]    E [deg] \n ')
fprintf('%4.2f  %4.4f   %4.4f       %4.4f     %4.4f    %4.4f', OE);
ISS (ZARYA)             
1 25544U 98067A   12341.40243075  .00007074  00000-0  12274-3 0  7812
2 25544  51.6485 346.0291 0016039  10.8890 135.6421 15.51944191804768

 a [km]   e      inc [deg]  RAAN [deg]  w[deg]    E [deg] 
 6789.18  0.0016   51.6485       346.0291     10.8890    135.6432

Published with MATLAB® 7.10

 

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