EC 723
Satellite Communication Systems
Mohamed Khedr
http://webmail.aast.edu/~khedr
Grades
Load
Percentage
Date
Midterm Exam
30%
Week of 3
December 2007
Final Exam
30%
Participation
10%
Report and
presentation
30%
Starting week
11
th
Textbook and website
Textbook:
non specific
Website: http://webmail.aast.edu/~khedr
Syllabus
Tentatively
Week 1
Overview
Week 2
Orbits and constellations: GEO, MEO and
LEO
Week 3
Satellite space segment, Propagation and
satellite links ,
channel modelling
Week 4
Satellite Communications Techniques
Week 5
Satellite error correction Techniques
Week 6
Multiple Access I
Week 7
Multiple access II
Week 8
Satellite in networks I
Week 9
INTELSAT systems , VSAT networks, GPS
Week 10
GEO, MEO and LEO mobile communications
INMARSAT systems, Iridium , Globalstar,
Odyssey
Week 11
Presentations
Week 12
Presentations
Week 13
Presentations
Week 14
Presentations
Week 15
Presentations
Satellite Components
Satellite Subsystems
Telemetry, Tracking, and Control
Electrical Power and Thermal Control
Attitude Control
Communication Subsystems
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)
Classification of Satellite Orbits
Circular or elliptical orbit
Circular with center at earth’s center
Elliptical with one foci at earth’s center
Orbit around earth in different planes
Equatorial orbit above earth’s equator
Polar orbit passes over both poles
Other orbits referred to as inclined orbits
Altitude of satellites
Geostationary orbit (GEO)
Medium earth orbit (MEO)
Low earth orbit (LEO)
Satellite Orbits
Equatorial
Inclined
Polar
Here’s the Math…
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
Let’s look in a Physics Book…
From what we have deduced so far, there
has to be an equation that relates the orbit
and the speed of the satellite:
T
2
r
3
4
10
14
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).
R^3=mu/n^2
N=2pi/T
The Most Common Example
“Height” of the orbit = 22,300 mile
That is 36,000km = 3.6x10
7
m
The radius of the orbit is
3.6x10
7
m + 6.38x10
6
m = 4.2x10
7
m
Put that into the formula and …
The Geosynchronous Orbit
The answer is T = 86,000 sec (rounded)
86,000 sec = 1,433 min = 24hours (rounded)
The satellite needs 1 day to complete an
orbit
Since the earth turns once per day, the
satellite moves with the surface of the earth.
Assignment
How long does a Low Earth Orbit Satellite
need for one orbit at a height of 200miles =
322km = 3.22x10
5
m
Do this:
Add the radius of the earth, 6.38x10
6
m
Compute T from the formula
Change T to minutes or hours
T
2
r
3
4
10
14
base station
or gateway
Classical satellite systems
Inter Satellite Link
(ISL)
Mobile User
Link (MUL)
Gateway Link
(GWL)
footprint
small cells
(spotbeams)
User data
PSTN
ISDN
GSM
GWL
MUL
PSTN: Public Switched
Telephone Network
Basics
Satellites in circular orbits
attractive force F
g
= m g (R/r)²
centrifugal force F
c
= m r
²
m: mass of the satellite
R: radius of the earth (R = 6370 km)
r: distance to the center of the earth
g: acceleration of gravity (g = 9.81 m/s²)
: angular velocity (
= 2
f, f: rotation frequency)
Stable orbit
F
g
= F
c
3
2
2
)
2
(
f
gR
r
Satellite period and orbits
10
20
30
40 x10
6
m
24
20
16
12
8
4
radius
satellite
period [h]
velocity [ x1000 km/h]
synchronous distance
35,786 km
12
10
8
6
4
2
Velocity
Km/sec
Basics
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
transponder used for sending/receiving and shifting of
frequencies
transparent transponder: only shift of frequencies
regenerative transponder: additionally signal regeneration
Inclination
inclination
d
d
satellite orbit
perigee
plane of satellite orbit
equatorial plane
Elevation
Elevation:
angle
e
between center of satellite beam
and surface
e
minimal elevation:
elevation needed at least
to communicate with the satellite
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
Orbits I
Orbits II
earth
km
35768
10000
1000
LEO
(Globalstar,
Irdium)
HEO
inner and outer Van
Allen belts
MEO (ICO)
GEO (Inmarsat)
Van

Allen

Belts:
ionized particles
2000

6000 km and
15000

30000 km
above earth surface
Geostationary satellites
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)
not useful for global coverage for small mobile phones
and data transmission, typically used for radio and TV
transmission
LEO systems
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
Examples:
Iridium (start 1998, 66 satellites)
Bankruptcy in 2000, deal with US DoD (free use,
saving from “deorbiting”)
Globalstar (start 1999, 48 satellites)
Not many customers (2001: 44000), low stand

by times for mobiles
MEO systems
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
Example:
ICO (Intermediate Circular Orbit, Inmarsat) start ca. 2000
Bankruptcy, planned joint ventures with Teledesic, Ellipso
–
cancelled again,
start planned for 2003
Routing
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
Localization of mobile stations
Mechanisms similar to GSM
Gateways maintain registers with user data
HLR (Home Location Register): static user data
VLR (Visitor Location Register): (last known) location of the mobile station
SUMR (Satellite User Mapping Register):
satellite assigned to a mobile station
positions of all satellites
Registration of mobile stations
Localization of the mobile station via the satellite’s position
requesting user data from HLR
updating VLR and SUMR
Calling a mobile station
localization using HLR/VLR similar to GSM
connection setup using the appropriate satellite
Handover in satellite systems
Several additional situations for handover in satellite systems compared to
cellular terrestrial mobile phone networks caused by the movement of the
satellites
Intra satellite handover
handover from one spot beam to another
mobile station still in the footprint of the satellite, but in another cell
Inter satellite handover
handover from one satellite to another satellite
mobile station leaves the footprint of one satellite
Gateway handover
Handover from one gateway to another
mobile station still in the footprint of a satellite, but gateway leaves the
footprint
Inter system handover
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.
Overview of LEO/MEO systems
Iridium
Globalstar
ICO
Teledesic
# satellites
66 + 6
48 + 4
10 + 2
288
altitude
(km)
780
1414
10390
ca. 700
coverage
global
70° latitude
global
global
min.
elevation
8°
20°
20°
40°
frequencies
[
GHz
(circa)]
1.6 MS
29.2
19.5
23.3 ISL
1.6 MS
2.5 MS
5.1
6.9
2 MS
2.2 MS
5.2
7
19
28.8
62 ISL
access
method
FDMA/TDMA
CDMA
FDMA/TDMA
FDMA/TDMA
ISL
yes
no
no
yes
bit rate
2.4
kbit/s
9.6
kbit/s
4.8
kbit/s
64
Mbit/s
2/64
Mbit/s
# channels
4000
2700
4500
2500
Lifetime
[years]
58
7.5
12
10
cost
estimation
4.4 B$
2.9 B$
4.5 B$
9 B$
Kepler’s First Law
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
barycenter,
is
always centered on one of the foci.
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
a,
and the semiminor
axis, by
b.
The eccentricity
e
is given by
b
a
b
a
e
Kepler’s Second Law
For equal time intervals, a satellite will sweep out equal areas in its
orbital plane, focused at the barycenter.
Kepler’s Third Law
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
Definition of terms for earth

orbiting satellite
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
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
i
in Fig.
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.
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.
Definition of terms for
earth

orbiting satellite
Six Orbital Elements
Earth

orbiting artificial satellites are defined by six orbital elements
referred to as the
keplerian element set.
The semimajor axis
a.
The eccentricity
e
give the shape of the ellipse.
A third, the mean anomaly
M
, gives the position of the satellite in its
orbit at a reference time known as the
epoch.
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.
The inclination
I
The right ascension of the ascending node
Relate the orbital plane’s position to the earth.
NASA
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.
The gravitational force inward (
F
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.
Forces acting on a satellite in a stable orbit around the earth.
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
cz
, where
c
is the center of the
earth and
cz
passes through the
geographic north pole.
Axes
cx
,
cy
, and
cz
are mutually
orthogonal axes, with
cx
and
cy
passing
through the earth’s geographic equator.
The vector
r
locates the moving satellite
with respect to the center of the earth.
Cartesian coordinate system
In this coordinate system, the
orbital plane of the satellite is
used as the reference plane. The
orthogonal axes,
x
0
and
y
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
z
0
axis except for satellite orbits that
are exactly in the plane of the
geographical equator.
The orbital plane coordinate system.
The plane of the orbit coincides with the plane of the paper. The axis
z
0
is
straight out of the paper from the center of the earth, and is normal to the
plane of the satellite’s orbit. The satellite’s position is described in terms of
the radius from the center of the earth
r
0
and the angle this radius makes
with the
x
0
axis,
Φ
o
.
Polar coordinate system in the plane of the satellite’s orbit.
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
t
1
and
t
2
and sweeps out an area denoted by
A
12
.
While close to apogee, the satellite moves in the orbit between times
t
3
and
sweeps out an area denoted by
A
34
. If
t
1
–
t
2
=
t
3
–
t
4
, then
A
12
=
A
34
.
Kepler’s second law of planetary motion.
The orbit as it appears in the
orbital plane.
The point
O
is the center of
the earth and the point
C
is
the center of the ellipse.
The two centers do not
coincide unless the
eccentricity,
e
, of the ellipse is
zero (i.e., the ellipse becomes
a circle and
a
=
b
).
The dimensions
a
and
b
are
the semimajor and semiminor
axes of the orbital ellipse,
respectively.
Point
O
is the center of the
earth and point
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 (
x
0
,
y
0
).
A vertical line through the
satellite intersects the
circumscribed circle at point
A
.
The eccentric anomaly
E
is the
angle from the
x
0
axis to the
line joining C and
A
.
The circumscribed circle and the eccentric anomaly
E
.
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
RA
and its
declination
d
.
The geocentric equatorial system.
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 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.
The definition of elevation (
EI
) and azimuth (
Az
).
The line joining the satellite and the center of the earth,
C
, passes through the
surface of the earth and point
Sub
, the subsatellite point.
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.
Zenith and nadir pointing directions.
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 satellite is said to
be visible from the
earth station if the
elevation angle
EI
is
positive. This requires
that the orbital radius
r
s
be greater than the
ratio
r
e
/cos(
), where
r
e
is the radius of the
earth and
is the
central angle.
The geometry of the visibility calculation.
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,
Communications Satellite Systems
, 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.
Thank you
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