EC 723 Satellite Communication Systems

Mobile - Wireless

Nov 24, 2013 (4 years and 7 months ago)

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EC 723

Satellite Communication Systems

Mohamed Khedr

http://webmail.aast.edu/~khedr

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
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

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:

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

(ISL)

Mobile User

(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

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

-

satellite

-

base station

transponder used for sending/receiving and shifting of
frequencies

transparent transponder: only shift of frequencies

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

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

connection of two mobile phones

Problems:

more complex focusing of antennas between satellites

high system complexity due to moving routers

higher fuel consumption

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

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
[years]
5-8
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.

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
°
.

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

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 satellite is said to
be visible from the
earth station if the
elevation angle
EI

is
positive. This requires
r
s

be greater than the
ratio
r
e
/cos(

), where
r
e

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