th

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Week 1
Week 2
Week 3
Week 4
Week 5
Week 6
Week 7
Week 8
Overview
Orbits and constellations: GEO, MEO and
LEO
Satellite space segment, Propagation and satellite links , channel modelling
Satellite Communications Techniques
Satellite error correction Techniques
Multiple Access I
Multiple access II
Satellite in networks I
Week 9
Week 10
Week 11
Week 12
INTELSAT systems , VSAT networks, GPS
GEO, MEO and LEO mobile communications
INMARSAT systems, Iridium , Globalstar,
Odyssey
Presentations
Presentations
Week 13 Presentations
Week 14 Presentations
Week 15 Presentations

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Satellite Subsystems
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Telemetry, Tracking, and Control
Electrical Power and Thermal Control
Attitude Control
Communication Subsystems
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Link Budget
Modulation Techniques
Coding and Error Correction
Networking (service provisioning, multimedia constraints and QoS)
Multiple Access and On-board Processing
Applications (Internet, Mobile computing)

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Circular with center at earth’s center
Elliptical with one foci at earth’s center
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Equatorial orbit above earth’s equator
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Polar orbit passes over both poles
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Other orbits referred to as inclined orbits
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Geostationary orbit (GEO)
Medium earth orbit (MEO)
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Low earth orbit (LEO)

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Gravity depends on the mass of the earth, the mass of the satellite, and the distance between the center of the earth and the satellite
For a satellite traveling in a circle, the speed of the satellite and the radius of the circle determine the force (of gravity) needed to maintain the orbit
The radius of the orbit is also the distance from the center of the earth.
For each orbit the amount of gravity available is therefore fixed
That in turn means that the speed at which the satellite travels is determined by the orbit

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T  2  r 3
4  10 14
R^3=mu/n^2
N=2pi/T
T is the time for one full revolution around the orbit, in seconds r is the radius of the orbit, in meters, including the radius of the earth (6.38x10
6 m).

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7
7
6
7

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5
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6
T  2  r 3
4  10 14

footprint
Mobile User
Link (MUL) small cells
(spotbeams)
Inter Satellite Link
(ISL)
Gateway Link
(GWL)
GWL base station or gateway
ISDN PSTN
MUL
GSM
PSTN: Public Switched
Telephone Network
User data

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Satellites in circular orbits
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 attractive force F g centrifugal force F c
= m g (R/r)²
= m r m: mass of the satellite
 ²
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R: radius of the earth (R = 6370 km)
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 r: distance to the center of the earth g: acceleration of gravity (g = 9.81 m/s²)
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 : angular velocity (  = 2  f, f: rotation frequency)
Stable orbit
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F g
= F c r

3
( 2 gR
 f
2
)
2

satellite period [h]
24
Velocity
Km/sec
12 velocity [ x1000 km/h]
10
20
8
16
6
12
4
8
2
4 synchronous distance
35,786 km
40 x10 6 m 10 20 radius
30

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 elliptical or circular orbits complete rotation time depends on distance satellite-earth inclination: angle between orbit and equator elevation: angle between satellite and horizon
LOS (Line of Sight) to the satellite necessary for connection
 high elevation needed, less absorption due to e.g. buildings
Uplink: connection base station - satellite
Downlink: connection satellite - base station typically separated frequencies for uplink and downlink
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 transponder used for sending/receiving and shifting of frequencies transparent transponder: only shift of frequencies
 regenerative transponder: additionally signal regeneration

plane of satellite orbit satellite orbit perigee d inclination d equatorial plane

Elevation: angle e between center of satellite beam and surface minimal elevation: elevation needed at least to communicate with the satellite e

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Four different types of satellite orbits can be identified depending on the shape and diameter of the orbit:
GEO: geostationary orbit, ca. 36000 km above earth surface
LEO (Low Earth Orbit): ca. 500 - 1500 km
MEO (Medium Earth Orbit) or ICO (Intermediate
Circular Orbit): ca. 6000 - 20000 km
HEO (Highly Elliptical Orbit) elliptical orbits

Van-Allen-Belts: ionized particles
2000 - 6000 km and
15000 - 30000 km above earth surface
HEO
LEO
(Globalstar,
Irdium) earth
1000
10000
35768 km
GEO (Inmarsat)
MEO (ICO) inner and outer Van
Allen belts

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Orbit 35,786 km distance to earth surface, orbit in equatorial plane (inclination 0°)
 complete rotation exactly one day, satellite is synchronous to earth rotation fix antenna positions, no adjusting necessary satellites typically have a large footprint (up to 34% of earth surface!), therefore difficult to reuse frequencies bad elevations in areas with latitude above 60° due to fixed position above the equator high transmit power needed high latency due to long distance (ca. 275 ms)

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Orbit ca. 500 - 1500 km above earth surface visibility of a satellite ca. 10 - 40 minutes global radio coverage possible latency comparable with terrestrial long distance connections, ca. 5 - 10 ms smaller footprints, better frequency reuse but now handover necessary from one satellite to another many satellites necessary for global coverage more complex systems due to moving satellites
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Examples:
Iridium (start 1998, 66 satellites)
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Bankruptcy in 2000, deal with US DoD (free use, saving from “deorbiting”)
Globalstar (start 1999, 48 satellites)
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Not many customers (2001: 44000), low stand-by times for mobiles

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Orbit ca. 5000 - 12000 km above earth surface comparison with LEO systems: slower moving satellites less satellites needed simpler system design for many connections no hand-over needed higher latency, ca. 70 - 80 ms higher sending power needed special antennas for small footprints needed
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Example:
ICO (Intermediate Circular Orbit, Inmarsat) start ca. 2000
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Bankruptcy, planned joint ventures with Teledesic, Ellipso – cancelled again, start planned for 2003

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One solution: inter satellite links (ISL) reduced number of gateways needed forward connections or data packets within the satellite network as long as possible only one uplink and one downlink per direction needed for the connection of two mobile phones
Problems: more complex focusing of antennas between satellites high system complexity due to moving routers higher fuel consumption thus shorter lifetime
Iridium and Teledesic planned with ISL
Other systems use gateways and additionally terrestrial networks

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Mechanisms similar to GSM
Gateways maintain registers with user data
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HLR (Home Location Register): static user data
VLR (Visitor Location Register): (last known) location of the mobile station
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SUMR (Satellite User Mapping Register):
 satellite assigned to a mobile station
 positions of all satellites
Registration of mobile stations
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Localization of the mobile station via the satellite’s position requesting user data from HLR
 updating VLR and SUMR
Calling a mobile station
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 localization using HLR/VLR similar to GSM connection setup using the appropriate satellite

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Several additional situations for handover in satellite systems compared to cellular terrestrial mobile phone networks caused by the movement of the satellites
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Intra satellite handover
 handover from one spot beam to another
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 mobile station still in the footprint of the satellite, but in another cell
Inter satellite handover
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 handover from one satellite to another satellite mobile station leaves the footprint of one satellite
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Gateway handover
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Handover from one gateway to another
 mobile station still in the footprint of a satellite, but gateway leaves the footprint
Inter system handover
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Handover from the satellite network to a terrestrial cellular network
 mobile station can reach a terrestrial network again which might be cheaper, has a lower latency etc.

Iridium
# satellites 66 + 6 altitude 780
(km) coverage global min.
elevation frequencies
[GHz
(circa)] access method
ISL bit rate
8°
1.6 MS
29.2
19.5
23.3 ISL
FDMA/TDMA CDMA yes
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2.4 kbit/s
Globalstar
48 + 4
1414
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70° latitude
20°
1.6 MS
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2.5 MS
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5.1
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6.9
 no
9.6 kbit/s
# channels 4000
Lifetime
[years]
5-8 cost estimation
4.4 B$
2700
7.5
2.9 B$
ICO
10 + 2
10390
Teledesic
288 ca. 700 global
20° global
40°
2 MS
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2.2 MS
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5.2
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7
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19
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28.8
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62 ISL
FDMA/TDMA FDMA/TDMA no
4.8 kbit/s
4500
12
4.5 B$ yes
64 Mbit/s
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2/64 Mbit/s
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2500
10
9 B$

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The path followed by a satellite around the primary will be an ellipse.
An ellipse has two focal points shown as F 1 and F 2.
The center of mass of the two-body system, termed the always centered on one of the foci. barycenter, is
In our specific case, because of the enormous difference between the masses of the earth and the satellite, the center of mass coincides with the center of the earth, which is therefore always at one of the foci.
The semimajor axis of the ellipse is denoted by axis, by b. The eccentricity e is given by a, and the semiminor e
 a a

 b b

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For equal time intervals, a satellite will sweep out equal areas in its orbital plane, focused at the barycenter.

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The square of the periodic time of orbit is proportional to the cube of the mean distance between the two bodies.
The mean distance is equal to the semimajor axis a. For the satellites orbiting the earth, Kepler’s third law can be written in the form
 where n is the mean motion of the satellite in radians per second and is the earth’s geocentric gravitational constant. With a in meters, its value is

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Apogee The point farthest from earth.
Apogee height is shown as ha in Fig
Perigee The point of closest approach to earth. The perigee height is shown as i in Fig.
hp
Line of apsides The line joining the perigee and apogee through the center of the earth.
Ascending node The point where the orbit crosses the equatorial plane going from south to north.
Descending node The point where the orbit crosses the equatorial plane going from north to south.
Line of nodes The line joining the ascending and descending nodes through the center of the earth.
Inclination The angle between the orbital plane and the earth’s equatorial plane. It is measured at the ascending node from the equator to the orbit, going from east to north.
The inclination is shown as
Mean anomaly M gives an average value of the angular position of the satellite with reference to the perigee.
True anomaly is the angle from perigee to the satellite position, measured at the earth’s center. This gives the true angular position of the satellite in the orbit as a function of time.

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Prograde orbit An orbit in which the satellite moves in the same direction as the earth’s rotation. The inclination of a prograde orbit always lies between 0 and 90°.
Retrograde orbit An orbit in which the satellite moves in a direction counter to the earth’s rotation. The inclination of a retrograde orbit always lies between 90 and 180°.
Argument of perigee The angle from ascending node to perigee, measured in the orbital plane at the earth’s center, in the direction of satellite motion.
Right ascension of the ascending node To define completely the position of the orbit in space, the position of the ascending node is specified. However, because the earth spins, while the orbital plane remains stationary the longitude of the ascending node is not fixed, and it cannot be used as an absolute reference.
For the practical determination of an orbit, the longitude and time of crossing of the ascending node are frequently used. However, for an absolute measurement, a fixed reference in space is required.
The reference chosen is the first point of Aries, otherwise known as the vernal, or spring, equinox. The vernal equinox occurs when the sun crosses the equator going from south to north, and an imaginary line drawn from this equatorial crossing through the center of the sun points to the first point of Aries (symbol ). This is the line of Aries.

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Earth-orbiting artificial satellites are defined by six orbital elements referred to as the
The semimajor axis
The eccentricity
 give the shape of the ellipse.
A third, the mean anomaly orbit at a reference time known as the
A fourth, the argument of perigee  , gives the rotation of the orbit’s perigee point relative to the orbit’s line of nodes in the earth’s equatorial plane.
I e keplerian element set. a.
M , gives the position of the satellite in its epoch.
The inclination
The right ascension of the ascending node 
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Relate the orbital plane’s position to the earth.

Forces acting on a satellite in a stable orbit around the earth.
Gravitational force is inversely proportional to the square of the distance between the centers of gravity of the satellite and the planet the satellite is orbiting, in this case the earth.
F The gravitational force inward (
IN
, the centripetal force) is directed toward the center of gravity of the earth.
The kinetic energy of the satellite ( F
OUT
, the centrifugal force) is directed opposite to the gravitational force. Kinetic energy is proportional to the square of the velocity of the satellite. When these inward and outward forces are balanced, the satellite moves around the earth in a “free fall” trajectory: the satellite’s orbit.

Cartesian coordinate system
The initial coordinate system that could be used to describe the relationship between the earth and a satellite.
A Cartesian coordinate system with the geographical axes of the earth as the principal axis is the simplest coordinate system to set up.
The rotational axis of the earth is about the axis earth and geographic north pole.
Axes cx , cz cy orthogonal axes, with through the earth’s geographic equator.
The vector
, where cz r passes through the
, and cz c is the center of the are mutually cx and cy passing locates the moving satellite with respect to the center of the earth.

The orbital plane coordinate system.
In this coordinate system, the orbital plane of the satellite is used as the reference plane. The x y orthogonal axes,
0 and
0 lie in the orbital plane. The third axis, z
0
, is perpendicular to the orbital plane. The geographical z -axis of the earth (which passes through the true North Pole and the center of the earth, c ) does not lie in the same direction as the
0 axis except for satellite orbits that are exactly in the plane of the geographical equator.
z

Polar coordinate system in the plane of the satellite’s orbit.
The plane of the orbit coincides with the plane of the paper. The axis x Φ r z plane of the satellite’s orbit. The satellite’s position is described in terms of the radius from the center of the earth
0 and the angle this radius makes with the
0 axis, o
.
0 is straight out of the paper from the center of the earth, and is normal to the

Kepler’s second law of planetary motion.
A satellite is in orbit about the planet earth, E .
The orbit is an ellipse with a relatively high eccentricity, that is, it is far from being circular.
Two shaded portions of the elliptical plane in which the orbit moves, one is close to the earth and encloses the perigee while the other is far from the earth and encloses the apogee.
The perigee is the point of closest approach to the earth while the apogee is the point in the orbit that is furthest from the earth.
While close to perigee, the satellite moves in the orbit between times and sweeps out an area denoted by A
12
.
While close to apogee, the satellite moves in the orbit between times sweeps out an area denoted by A
34
. If t
1
– t
2
= t
3
– t
4
, then A
12
= A
34
.
t t
1
3 and and t
2

The orbit as it appears in the orbital plane.
The point O a circle and is the center of the earth and the point the center of the ellipse.
The two centers do not coincide unless the eccentricity, e zero (i.e., the ellipse becomes a
, of the ellipse is
=
The dimensions b a
). and b
C is are the semimajor and semiminor axes of the orbital ellipse, respectively.

The circumscribed circle and the eccentric anomaly
.
Point O is the center of the earth and point
A vertical line through the satellite intersects the circumscribed circle at point
The eccentric anomaly x
C is both the center of the orbital ellipse and the center of the circumscribed circle.
The satellite location in the orbital plane coordinate system is specified by (
A .
x
E
0
, y
0
).
A . is the angle from the
0 line joining C and axis to the

The geocentric equatorial system.
This geocentric system differs from that shown in Figure 2.1 only in that the x i axis points to the first point of
Aries. The first point of Aries is the direction of a line from the center of the earth through the center of the sun at the vernal equinox (about
March 21 in the Northern
Hemisphere), the instant when the subsolar point crosses the equator from south to north. In the above system, an object may be located by its right ascension declination d .
RA and its

Locating the orbit in the geocentric equatorial system.
The satellite penetrates the equatorial plane (while moving in the positive z direction) at the ascending node.
The right ascension of the ascending node is  and the inclination i is the angle between the equatorial plane and the orbital plane.
Angle  , measured in the orbital plane, locates the perigee with respect to the equatorial plane.

The definition of elevation (
) and azimuth (
).
The elevation angle is measured upward from the local horizontal at the earth station and the azimuth angle is measured from the true north in an eastward direction to the projection of the satellite path onto the local horizontal plane.

Zenith and nadir pointing directions.
The line joining the satellite and the center of the earth, surface of the earth and point Sub
C
, the subsatellite point.
, passes through the
The satellite is directly overhead at this point and so an observer at the subsatellite point would see the satellite at zenith (i.e., at an elevation angle of 90 ° ).
The pointing direction from the satellite to the subsatellite point is the nadir direction from the satellite.
If the beam from the satellite antenna is to be pointed at a location on the earth that is not at the subsatellite point, the pointing direction is defined by the angle away from nadir.
In general, two off-nadir angles are given: the number of degrees north (or south) from nadir; and the number of degrees east (or west) from nadir. East, west, north, and south directions are those defined by the geography of the earth.

The geometry of elevation angle calculation. The plane of the paper is the plane defined by the center of the earth, the satellite, and the earth station. The central angle is  . The elevation angle EI is measured upward from the local horizontal at the earth station.

The geometry of the visibility calculation.
The satellite is said to be visible from the earth station if the
is elevation angle positive. This requires that the orbital radius be greater than the ratio
e
/cos(  ), where is the radius of the earth and  is the central angle.
s
e

During the equinox periods around the March 21 and September 3, the geostationary plane is in the shadow of the earth on the far side of the earth from the sun. As the satellite moves around the geostationary orbit, it will pass through the shadow and undergo an eclipse period.
The length of the eclipse period will vary from a few minutes to over an hour (see Figure 2.22), depending on how close the plane of the geostationary orbit is with respect to the center of the shadow thrown by the earth.

Dates and duration of eclipses. (Source: Martin,
, Prentice Hall 1978.)

Schematic of sun outage conditions. During the equinox periods, not only does the earth’s shadow cause eclipse periods to occur for geostationary satellites, during the sunlit portion of the orbit, there will be periods when the sun appears to be directly behind the satellite. At the frequencies used by communications satellites (4 to 50 GHz), the sun appears as a hot noise source. The effective temperature of the sun at these frequencies is on the order of 10,000 K. The precise temperature observed by the earth station antenna will depend on whether the beamwidth partially, or completely, encloses the sun.