Home » Posts tagged 'Earth topographic map'

# Tag Archives: Earth topographic map

## Satellite Ground Track, GPS BII-10

The Global Positioning System (GPS) is a space-based satellite navigation system that provides location and time information in all weather conditions, anywhere on or near the Earth where there is an unobstructed line of sight to four or more GPS satellites. In this example we implement algorithm to plot GPS ground track. Same algorithm could be used for any satellite.

```clear all;
clc;
% Earth topographic map
figure(1);
xwidth = 820;
ywidth = 420;
hFig = figure(1);
set(gcf,'PaperPositionMode','auto')
set(hFig, 'Position', [100 100 xwidth ywidth])
hold on;
grid on;
axis([0 360 -90 90]);
contour(0:359,-89:90,topo,[0 0],'b')
axis equal
box on
set(gca,'XLim',[0 360],'YLim',[-90 90], ...
'XTick',[0 60 120 180 240 300 360], ...
'Ytick',[-90 -60 -30 0 30 60 90]);
image([0 360],[-90 90],topo,'CDataMapping', 'scaled');
colormap(topomap1);
ylabel('Latitude [deg]');
xlabel('Longitude [deg]');
title('GPS BII-10 ground track');

R_e = 6378;        % Earth's radius
mu = 398600;       % Earth’s gravitational parameter [km^3/s^2]
J2 = 0.0010836;
we = 360*(1 + 1/365.25)/(3600*24);      % Earth's rotation [deg/s]
% GPS BII-10 Orbital Parametres
rp    =  19781  + R_e;       % [km] Perigee Radius
ra    =  20582  + R_e;       % [km] Apogee Radius
theta =  25;                 % [deg] True anomaly
RAAN  =  229.9128 ;          % [deg] Right ascension of the ascending node
i     =  54.4303 ;           % [deg] Inclination
omega =  335.2539  ;         % [deg] Argument of perigee

a = (ra+rp)/2;               % Semimajor axis
e = (ra -rp)/(ra+rp) ;       % Eccentricity
h = (mu*rp*(1 + e))^0.5;     % Angular momentum
T = 2*pi*a^1.5/mu^0.5;       % Period
dRAAN = -(1.5*mu^0.5*J2*R_e^2/((1-e^2)*a^3.5))*cosd(i)*180/pi;
domega = dRAAN*(2.5*sind(i)^2 - 2)/cosd(i);
% Initial state
[R0 V0] = Orbital2State( h, i, RAAN, e,omega,theta);
[ alfa0 ,delta0 ] = R2RA_Dec( R0 );
scatter(alfa0,delta0,'*k');

ind = 1;
eps = 1E-9;
dt = 20;        % time step [sec]
ti = 0;

while(ti <= 3*T);
E = 2*atan(tand(theta/2)*((1-e)/(1+e))^0.5);
M = E  - e*sin(E);
t0 = M/(2*pi)*T;
t = t0 + dt;
M = 2*pi*t/T;
E = keplerEq(M,e,eps);
theta = 2*atan(tan(E/2)*((1+e)/(1-e))^0.5)*180/pi;
RAAN  = RAAN  +  dRAAN*dt ;
omega = omega + domega*dt;
[R V] = Orbital2State( h, i, RAAN, e,omega,theta);
% Considering Earth's rotation
fi_earth = we*ti;
Rot = [cosd(fi_earth), sind(fi_earth),0;...
-sind(fi_earth),cosd(fi_earth),0;0,0,1];
R = Rot*R;
[ alfa(ind) ,delta(ind) ] = R2RA_Dec( R );
ti = ti+dt;
ind = ind + 1;
end
scatter(alfa,delta,'.r');
text(280,-80,'smallsats.org','Color',[1 1 1], 'VerticalAlignment','middle',...
'HorizontalAlignment','left','FontSize',14 );```

## Satellite Ground Track, MOLNIYA 1-93

Molniya satellite systems were military communications satellites used by the Soviet Union. The satellites used highly eccentric elliptical orbits of 63.4 degrees inclination and orbital period of about 12 hours. This type of orbits allowed satellites to be visible to polar regions for long periods.

In this example we will show how to plot ground track for satellites.

```clear all;
clc;
% Earth topographic map
figure(1);
xwidth = 820;
ywidth = 420;
hFig = figure(1);
set(gcf,'PaperPositionMode','auto')
set(hFig, 'Position', [100 100 xwidth ywidth])
hold on;
grid on;
axis([0 360 -90 90]);
contour(0:359,-89:90,topo,[0 0],'b')
axis equal
box on
set(gca,'XLim',[0 360],'YLim',[-90 90], ...
'XTick',[0 60 120 180 240 300 360], ...
'Ytick',[-90 -60 -30 0 30 60 90]);
image([0 360],[-90 90],topo,'CDataMapping', 'scaled');
colormap(topomap1);
ylabel('Latitude [deg]');
xlabel('Longitude [deg]');
title('MOLNIYA 1-93 satellite ground track');

R_e = 6378;        % Earth's radius
mu = 398600;       % Earth’s gravitational parameter [km^3/s^2]
J2 = 0.0010836;
we = 360*(1 + 1/365.25)/(3600*24);      % Earth's rotation [deg/s]
% MOLNIYA 1-93 Orbital Elements /Source: AFSPC
rp    =  1523.9 + R_e;       % [km] Perigee Radius
ra    =  38843.1 + R_e;      % [km] Apogee Radius
theta =  0;                  % [deg] True anomaly
RAAN  =  141.4992;           % [deg] Right ascension of the ascending node
i     =  64.7;               % [deg] Inclination
omega =  253.3915 ;          % [deg] Argument of perigee

a = (ra+rp)/2;               % Semimajor axis
e = (ra -rp)/(ra+rp) ;       % Eccentricity
h = (mu*rp*(1 + e))^0.5;     % Angular momentum
T = 2*pi*a^1.5/mu^0.5;       % Period
dRAAN = -(1.5*mu^0.5*J2*R_e^2/((1-e^2)*a^3.5))*cosd(i)*180/pi;
domega = dRAAN*(2.5*sind(i)^2 - 2)/cosd(i);
% Initial state
[R0 V0] = Orbital2State( h, i, RAAN, e,omega,theta);
[ alfa0 ,delta0 ] = R2RA_Dec( R0 );
scatter(alfa0,delta0,'*k');

ind = 1;
eps = 1E-9;
dt = 20;        % time step [sec]
ti = 0;

while(ti <= 2.5*T);
E = 2*atan(tand(theta/2)*((1-e)/(1+e))^0.5);
M = E  - e*sin(E);
t0 = M/(2*pi)*T;
t = t0 + dt;
M = 2*pi*t/T;
E = keplerEq(M,e,eps);
theta = 2*atan(tan(E/2)*((1+e)/(1-e))^0.5)*180/pi;
RAAN  = RAAN  +  dRAAN*dt ;
omega = omega + domega*dt;
[R V] = Orbital2State( h, i, RAAN, e,omega,theta);
% Considering Earth's rotation
fi_earth = we*ti;
Rot = [cosd(fi_earth), sind(fi_earth),0;...
-sind(fi_earth),cosd(fi_earth),0;0,0,1];
R = Rot*R;
[ alfa(ind) ,delta(ind) ] = R2RA_Dec( R );
ti = ti+dt;
ind = ind + 1;
end
scatter(alfa,delta,'.r');
text(280,-80,'smallsats.org','Color',[1 1 1], 'VerticalAlignment','middle',...
'HorizontalAlignment','left','FontSize',14 );```