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Transcript of COMMUNICATION ENGINEERING-II ACADEMIC YEAR … YEAR/COMMUNICATION... · communication...
1
COMMUNICATION ENGINEERING-II
ACADEMIC YEAR-1017-18
UNIT-I
SATELLITE COMMUNICATION SYSTEMS
2marks
1. What is Satellite?
An artificial body that is projected from earth to orbit either earth (or) another body of solar
systems.
Types: Information satellites and Communication Satellites
2. Define Satellite Communication.
It is defined as the use of orbiting satellites to receive, amplify and retransmit data to earth
stations.
3. State Kepler’s first law.
It states that the path followed by the satellite around the primary will be an ellipse. An ellipse
has two focal points F1 and F2. The center of mass of the two body system, termed the
barycenter is always centered on one of the foci.
e = [square root of (a2– b2) ] / a.(Explain with diagram, Refer notes for kepler’s laws )
4. State Kepler’s second law.
It states that for equal time intervals, the satellite will sweep out equal areas in its orbital
plane, focused at the barycenter.
5. State Kepler’s third law.
It states that the square of the periodic time of orbit is perpendicular to the
cube of the mean distance between the two bodies.
a3= 3 / n2
2
Where, n = Mean motion of the satellite in rad/sec.
2
3 = Earth’s geocentric gravitational constant. With the n in radians per sec. the orbital
period in second is given by,
P = 2 / n
6. Define apogee.
The point farthest from the earth.
7. Define Perigee.
The point closest from the earth.
8. What is line of apsides?
The line joining the perigee and apogee through the center of the earth.
9. Define ascending node.
The point where the orbit crosses the equatorial plane going from south to north.
10. Define descending node.
The point where the orbit crosses the equatorial plane going from north to south.
11. Mention the apogee and perigee height.
r a = a(1+e)
r p = a(1+e)
h a = r a – R p
h p = r p – R p
12. Give the 3 different types of applications with respect to satellite systems.
• The largest international system (Intelsat)
• The domestic satellite system (Dom sat) in U.S.
• U.S. National oceanographic and atmospheric administrations
(NOAA)
3
13. Mention the 3 regions to allocate the frequency for satellite services.
• Region1: It covers Europe, Africa and Mangolia.
• Region2: It covers North & South Ameriaca and Greenland.
• Region3: It covers Asia, Australia and South West Pacific.
14. Give the types of satellite services.
• Fixed satellite service
• Broadcasting satellite service
• Mobile satellite service
• Navigational satellite services
• Meteorological satellite services
15. Give the advantage of geostationary orbit.
There is no necessity for tracking antennas to find the satellite positions.
16. Define look angles.
The azimuth and elevation angles of the ground station antenna are termed as look angles.
17. What are the geostationary satellites?
The satellites present in the geostationary orbit are called geostationary satellite. The
geostationary orbit is one in which the satellite appears stationary relative to the earth. It lies in
equatorial plane and inclination is ‘0’. The satellite must orbit the earth in the same direction as
the earth spin. The orbit is circular.
18. Give the two segments of basic satellite communication.
a. Earth segment (or) ground segment
b. Space segment
19.What is meant by transponder?
In a communication satellite, the equipment which provides the connecting link between the
satellite’s transmit and receive antennas is referred to as the transponder.
21. Write short notes on station keeping.
4
It is the process of maintenance of satellite’s attitude against different factors that can cause
drift with time. Satellites need to have their orbits adjusted from time to time, because the
satellite is initially placed in the correct orbit, natural forces induce a progressive drift.
22. What is meant by Pitch angle?
Movement of a spacecraft about an axis which is perpendicular to its longitudinal axis. It is
the degree of elevation or depression.
23. What is meant by frequency reuse?
The carrier with opposite senses of polarization may overlap in frequency. This technique is
known as frequency reuse.
24. What is meant by GEO?
GEO means Geostationary or Geosynchronous earth orbit. GEO satellites have a distance
of almost 36000 km to the earth. Examples are almost all TV and radio broadcast satellites, many
weather satellites and satellites operating as backbone for the telephone network.
25. What are the advantages of GEO?
Three GEO satellites are enough for a complete coverage of almost any spot on earth,
senders and receivers can use fixed antennas positions, no adjusting is needed. Therefore GEO’s
are ideal for T.V and radio broadcasting.
26. What are the applications in satellites?
Satellites can be used in the Following Areas • Weather Forecasting • Radio and TV broadcast
Satellites • Military Satellites • Satellites for Navigation
27. What are the advantages of LEO(low earth orbit)?
• Data rate is 2400 bit/s
• Packet delay is relatively low
• Smaller footprints of LEO allows frequency reuse
• Provide high elevations
5
28. Define the inclination angle and perigee?
The inclination angle is defined as the angle between the equatorial plane and the plane
described by the satellite orbit. An inclination angle of 0 degrees means that the satellite is
exactly above the equator. If the satellite does not have a circular orbit, the closest point to the
earth is called the perigee.
29. Define the elevation angle and footprint ?
The elevation angle is defined as the angle between the centre of satellite beam and the
plane tangential to the earth’s surface. The foot-print can be defined as the area on earth where
the signals of the satellite can be received.
30. What is meant by GEO?
GEO means Geostationary or Geosynchronous earth orbit. GEO satellites have a distance
of almost 36000 km to the earth. Examples are almost all TV and radio broadcast satellites, many
weather satellites and satellites operating as backbone for the telephone network.
31. What are the advantages of GEO?
Three GEO satellites are enough for a complete coverage of almost any spot on earth,
senders and receivers can use fixed antennas positions, no adjusting is needed. Therefore GEO’s
are ideal for T.V and radio broadcasting.
32. What are the applications in satellites?
Satellites can be used in the Following Areas • Weather Forecasting • Radio and TV broadcast
Satellites • Military Satellites • Satellites for Navigation
33. What are the advantages of LEO?
• Data rate is 2400 bit/s
• Packet delay is relatively low • Smaller footprints of LEO allows frequency reuse •
Provide high elevations
34. Define the inclination angle and perigee?
The inclination angle is defined as the angle between the equatorial plane and the plane
described by the satellite orbit. An inclination angle of 0 degrees means that the satellite is
exactly above the equator. If the satellite does not have a circular orbit, the closest point to the
earth is called the perigee.
6
35. Define the elevation angle and footprint?
The elevation angle is defined as the angle between the centre of satellite beam and the
plane tangential to the earth’s surface. The foot-print can be defined as the area on earth where
the signals of the satellite can be received.
UNIT I
SATELLITE COMMUNICATION SYSTEM
SATELLITE ORBITS
The orbital locations of the spacecraft in a communications satellite system play a major
role in determining the coverage and operational characteristics of the services provided by that
system. This chapter describes the general characteristics of satellite orbits and summarizes the
characteristics of the most popular orbits for communications applications.
The same laws of motion that control the motions of the planets around the sun govern
artificial earth satellites that orbit the earth. Satellite orbit determination is based on the Laws of
Motion first developed by Johannes Kepler and later refined by Newton in 1665 from his own
Laws of Mechanics and Gravitation. Competing forces act on the satellite; gravity tends to pull
the satellite in towards the earth, while its orbital velocity tends to pull the satellite away from
the earth. Fig. 1 shows a simplified picture of the forces acting on an orbiting satellite.
The gravitational force, Fin , and the angular velocity force, Fout , can be represented as
Fin = m(μ
r2)
and
Fout = m(v2
r)
where m = satellite mass; v = satellite velocity in the plane of orbit; r = distance from the
center of the earth (orbit radius); and µ = Kepler’s Constant (or Geocentric Gravitational
Constant) = 3.986004 × 105 km
3/s
2.
Note that for Fin = Fout
7
v = (μ
r)
12
This result gives the velocity required to maintain a satellite at the orbit radius r. Note
that for the discussion above all other forces acting on the satellite, such as the gravity forces
from the moon, sun, and other bodies, are neglected.
Fig. 1 Forces in a satellite
KEPLER’S LAWS
Kepler’s laws of planetary motion apply to any two bodies in space that interact through
gravitation. The laws of motion are described through three fundamental principles.
Kepler’s First Law, as it applies to artificial satellite orbits, can be simply stated as
follows: ‘the path followed by a satellite around the earth will be an ellipse, with the center of
mass of earth as one of the two foci of the ellipse.’ This is shown in Fig. 2.
Fig. 2 Kepler’s First Law
If no other forces are acting on the satellite, either intentionally by orbit control or
unintentionally as in gravity forces from other bodies, the satellite will eventually settle in an
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elliptical orbit, with the earth as one of the foci of the ellipse. The ‘size’ of the ellipse will
depend on satellite mass and its angular velocity
Kepler’s Second Law can likewise be simply stated as follows: ‘for equal time intervals,
the satellite sweeps out equal areas in the orbital plane.’ Fig. 3 demonstrates this concept.
The shaded area A1 shows the area swept out in the orbital plane by the orbiting satellite
in a one hour time period at a location near the earth. Kepler’s second law states that the area
swept out by any other one hour time period in the orbit will also sweep out an area equal to A1 .
For example, the area swept out by the satellite in a one hour period around the point farthest
from the earth (the orbit’s apogee), labeled A2 on the figure, will be equal to A1 , i.e.: A1 = A2 .
This result also shows that the satellite orbital velocity is not constant; the satellite is moving
much faster at locations near the earth, and slows down as it approaches apogee. This factor will
be discussed in more detail later when specific satellite orbit types are introduced.
Fig. 3 Kepler’s Second Law
Stated simply, Kepler’s Third Law is as follows: ‘the square of the periodic time of
orbit is proportional to the cube of the mean distance between the two bodies.’ This is quantified
as follows:
T2 = [4π2
μ] r3
where T = orbital period in s; a = distance between the two bodies, in km; µ = Kepler’s
Constant (or Geocentric Gravitational Constant) = 3.986004 × 105 km
3/s
2.
If the orbit is circular, then a = r, and
r = [μ
4π2]
13T23
This demonstrates an important result:
9
Orbit Radius = [ Constant ] × (Orbit Period)2/3
Under this condition, a specific orbit period is determined only by proper selection of the
orbit radius. This allows the satellite designer to select orbit periods that best meet particular
application requirements by locating the satellite at the proper orbit altitude. The altitudes
required to obtain a specific number of repeatable ground traces with a circular orbit are listed in
Table 1.
Table 1 Orbit altitudes for specified orbital periods
Revolutions/day Nominal period (hours) Nominal altitude (km)
1
2
3
4
6
8
24
12
8
6
4
3
36000
20200
13900
10400
6400
4200
ORBITAL PARAMETERS
Fig. 4 shows two perspectives useful in describing the important orbital parameters used to
define earth-orbiting satellite characteristics. The parameters are:
Apogee – the point farthest from earth.
Perigee – the point of closest approach to earth.
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.
Argument of Perigee, ω – the angle from ascending node to perigee, measured in the
orbital plane.
Right Ascension of the Ascending Node, Φ – the angle measured eastward, in the
equatorial plane, from the line to the first point of Aries (Y) to the ascending node.
10
The eccentricity is a measure of the ‘circularity’ of the orbit. It is determined from
e =ra − rp
ra + rp
where e = the eccentricity of the orbit; ra = the distance from the center of the earth to the apogee
point; and rp = the distance from the center of the earth to the perigee point.
Fig. 4 Earth-orbiting satellite parameters
The higher the eccentricity, the ‘flatter’ the ellipse. A circular orbit is the special case of
an ellipse with equal major and minor axes (zero eccentricity). That is:
Elliptical Orbit 0 < e < 1
Circular Orbit e = 0
The inclination angle, θi, is the angle between the orbital plane and the earth’s equatorial
plane. A satellite that is in an orbit with some inclination angle is in an inclined orbit. A satellite
that is in orbit in the equatorial plane (inclination angle = 00) is in an equatorial orbit. A satellite
that has an inclination angle of 900 is in a polar orbit. The orbit may be elliptical or circular,
depending on the orbital velocity and direction of motion imparted to the satellite on insertion
into orbit.
11
Fig. 5 shows another important characteristic of satellite orbits. An orbit in which the
satellite moves in the same direction as the earth’s rotation is called a prograde orbit. The
inclination angle of a prograde orbit is between 00
and 900. A satellite in a retrograde orbit moves
in a direction opposite (counter to) the earth’s rotation, with an inclination angle between 900
and
1800. Most satellites are launched in a prograde orbit, because the earth’s rotational velocity
enhances the satellite’s orbital velocity, reducing the amount of energy required to launch and
place the satellite in orbit.
An almost endless number of combinations of orbital parameters are available for
satellite orbits. Orbital elements defines the set of parameters needed to uniquely specify the
location of an orbiting satellite. The minimum number of parameters required is six:
Eccentricity;
Semi-Major Axis;
Time of Perigee;
Right Ascension of Ascending Node;
Inclination Angle;
Argument of Perigee.
Fig. 5 Prograde and retrograde orbits
12
These parameters will uniquely define the absolute (i.e., the inertial) coordinates of the
satellite at any time t. They are used to determine the satellite track and provide a prediction of
satellite location for extended periods beyond the current time.
Satellite orbits coordinates are specified in sidereal time rather than in solar time. Solar
time, which forms the basis of all global time standards, is based on one complete rotation of the
earth relative to the sun. Sidereal time is based on one complete rotation of the earth relative to a
fixed star reference, as shown in Fig. 6.
Fig. 6 Sidereal time
ORBITS IN COMMON USE
With all the possible combinations of orbit parameters available to the satellite designer,
an almost endless list of possible orbits can be used. Experience has narrowed down the list of
orbits in common use for communications, sensor, and scientific satellites, and they are
introduced in the following sections. We begin with the most popular orbit used for
communications satellites – the geostationary (or geosynchronous) orbit.
Geostationary Orbit
Kepler’s third law demonstrated that there is a fixed relationship between orbit radius and
the orbit period of revolution. Under this condition a specific orbit period can be determined by
proper selection of the orbit radius.
If the orbit radius is chosen so that the period of revolution of the satellite is exactly set to
the period of the earth’s rotation, one mean sidereal day, a unique satellite orbit is defined.
13
In addition, if the orbit is circular (eccentricity = 0), and the orbit is in the equatorial
plane (inclination angle = 00), the satellite will appear to hover motionless above the earth at the
subsatellite point above the equator. This important special orbit is the geostationary earth orbit
(GEO). From Kepler’s third law, the orbit radius for the GEO, rS , is found as
rs = [μ
4π2]
13T23 = [
3.986004 × 105
4π2]
13
(86164.09)23
= 42164.17 km
where T = 1 mean sidereal day = 86 164.09 s.
The geostationary height (altitude above the earth’s surface), hS , is then
hs = rs − rE
= 42164-6378
= 35786 km
where rE = equatorial earth radius = 6378 km.
The value of hS is often rounded to 36 000 km for use in orbital calculations. The
geostationary orbit is an ideal orbit that cannot be achieved for real artificial satellites because
there are many other forces besides the earth’s gravity acting on the satellite. A ‘perfect orbit’,
i.e., one with e exactly equal to zero and with θi exactly equal to 00, cannot be practically
achieved without extensive station keeping and a vast amount of fuel to maintain the precise
position required. A typical GEO orbit in use today would have an inclination angle slightly
greater than 0 and possibly an eccentricity that also exceeds 0. The ‘real world’ GEO orbit that
results is often referred to as a geosynchronous earth orbit (GSO) to differentiate it from the ideal
geostationary orbit. 1
Most current communications satellites operate in a geosynchronous earth orbit, which is
ideally suited for the transfer of communications information between two or more points on the
earth through a ‘relay’ that is fixed in space, relative to the earth. Fig. 7 shows the basic elements
of the geosynchronous earth orbit as it applies to satellite operations. The GSO location provides
a fixed path from the ground to the satellite; therefore little or no ground tracking is
required.Asatellite in GSO sees about one-third of the earth’s surface, so three GSO satellites,
placed 1200 apart in the equatorial plane, could provide global coverage, except for the pole
areas (to be discussed further later).
14
Fig. 7 GSO – Geosynchronous earth orbit
The period of revolution for the geostationary orbit is 23 hours, 56 minutes, which is the
time for the earth to complete one revolution about its axis, measured relative to the star field
reference (sidereal time). It is four minutes shorter than the 24-hour mean solar day because of
the earth’s movement around the sun.
The geosynchronous orbit does suffer from some disadvantages, even though it is the
most heavily implemented orbit for current communications systems because of its fixed earth-
satellite geometry and its large coverage area. The long path length produces a large path loss
and a significant latency (time delay) for the radiowave signal propagating to and from the
satellite. The two-way (up to the satellite and back) delay will be approximately 260 ms for a
ground station located at a mid-latitude location. This could produce problems, particularly for
voice communications or for certain protocols that cannot tolerate large latency.
The GSO cannot provide coverage to high latitude locations. The highest latitude, at
which the GSO satellite is visible, with a 10◦ earth station elevation angle, is about 70◦, North or
South latitude. Coverage can be increase somewhat by operation at higher inclination angles, but
that produces other problems, such as the need for increased ground antenna tracking, which
increases costs and system complexity.
The number of satellites that can operate in geostationary orbits is obviously limited,
because there is only one equatorial plane, and the satellites must be spaced to avoid interference
between each other. The allocation of geostationary orbital locations or slots is regulated by
international treaties through the International Telecommunications Union, in close coordination
with frequency band and service allocations, as discussed in Chapter 1. Current allocations place
satellites in the range of 2–50 apart for each frequency band and service allocation, meaning that
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only 72–180 slots are available for global use, depending on the frequency band and service
provided.
Low Earth Orbit
Earth satellites that operate well below the geostationary altitude, typically at altitudes
from 160 to 2500 km, and in near circular orbits, are referred to as low earth orbit or LEO
satellites. 2 The low earth orbit satellite has several characteristics that can be advantageous for
communications applications, as summarized on Fig. 8.
Fig. 8 LEO – Low earth orbit
The earth-satellite links are much shorter, leading to lower path losses, which result in
lower power, smaller antenna systems. Propagation delay is also less because of shorter path
distances. LEO satellites, with the proper inclinations, can cover high latitude locations,
including polar areas, which cannot be reached by GSO satellites.
A major disadvantage of the LEO satellite is its restricted operations period, because the
satellite is not at a fixed location in the sky, but instead sweeps across the sky for as little as 8 to
10 minutes from a fixed location on earth. If continuous global or wide area coverage is desired,
a constellation of multiple LEO satellites is required, with links between the satellites to allow
for point-to-point communications. Some current LEO satellite networks operate with 12, 24,
and 66 satellites to achieve the desired coverage.
The oblateness (non-spherical shape) of the earth will cause two major perturbations to
the LEO orbit. The point on the equator where the LEO satellite crosses from south to north (the
ascending node) will drift westward several degrees per day. A second effect of the earth’s
oblateness is to rotate the orientation of the major axis in the plane of the orbit, either clockwise
16
or counterclockwise. If the inclination is set to about 63◦, however, the forces that induce the
rotation will be balanced and the major axis direction remains fixed.
The LEO orbit has found serious consideration for mobile applications, because the small
power and small antenna size of the earth terminals are a definite advantage. More LEO satellites
are required to provide communications services comparable to the GSO case, but LEO satellites
are much smaller and require significantly less energy to insert into orbit, hence total life cycle
costs may be lower.
Medium Earth Orbit
Satellites that operate in the range between LEO and GSO, typically at altitudes of 10 000
to 20 000 km, are referred to as medium altitude orbit, or MEO satellites. The basic elements of
the MEO are summarized on Fig. 9.
Fig. 9 MEO – Medium earth orbit
The desirable features of the MEO include: repeatable ground traces for recurring ground
coverage; selectable number of revolutions per day; and adequate relative satellite-earth motion
to allow for accurate and precise position measurements. A typical MEO would provide one to
two hours of observation time for an earth terminal at a fixed location. MEO satellites have
characteristics that have been found useful for meteorological, remote sensing, navigation, and
position determination applications. The Global Positioning System (GPS), for example,
employs a constellation of up to 24 satellites operating in 12-hour circular orbits, at an altitude of
20184 km.
SATELLITE LAUNCH SYSTEMS
17
Background
The first launch systems to place satellites into orbits around the Earth were
developed by government agencies in the 1950s to insert satellite communication and
observation systems into low-Earth orbits (150-200 km altitude). Most of these launchers
were modelled after the intercontinental ballistic missiles of the period. In the 1960s era,
space exploration programmes associated with flights to the Moon and planets resulted in the
development of powerful rockets that were capable of inserting satellites into the geostationary
orbit, commonly referred to as the "GSO" (35 786 km altitude). The era of the extensive use of
GSO communication satellites started in the 1970s and has continued without interruption to the
present time.
Recently, considerable interest has been shown for the development of new
non-GSO communication satellites which have very different launch requirements from
GSO satellites. The technology, however, is well developed since many non-GSO satellites
with a variety of service missions (weather, earth mapping, navigation, etc.) have been
launched during the last several decades. Also, many launch systems with GSO capabilities
are able to insert several LEO satellites into low- or medium-Earth orbits with one launch
operation.
Launcher considerations
The basic requirements for the selection of a launch system are 1) lift capability to the
desired orbit; 2) availability after the satellite construction and test phase has been
completed; and 3) cost of equipment and services. Until recently, the choice has been limited
and negotiations have normally been with government agencies. Now, a new era has evolved in
which a range of launch vehicles are being offered internationally on a commercial basis
by competing private companies and government organizations. The launch industry is
expanding rapidly and new performance capabilities and services are constantly being
featured. Thus, this section should only be regarded as a guide to what may be available.
Direct contact with the suppliers will be necessary in order to obtain all the necessary
details associated with contracting for a launch system.
TYPES OF LAUNCH SYSTEMS
18
Geostationary orbit (GSO)
The predominant launch systems for GSO satellites have expendable boosters
which employ several steps for inserting a satellite into its final orbit. The first step usually
involves a few rocket firing phases which place the satellite and its attached apogee rocket motor
(ARM) into a transfer orbit with a perigee of approximately 200 km in altitude and an
apogee at the GSO altitude. At apogee, the ARM is fired to circularize the orbit into a
geosynchronous mode. Some available launch systems with these characteristics include
the ARIANE, ATLAS, DELTA, H-Series, LLV, LONG MARCH, M-Series, PROTON,
TITAN, ZENIT, among others. A brief description of the capabilities of these systems is
provided in the following sections.
There has been interest in developing reusable launchers in which the launch vehicle is
returned to Earth intact and then readied for the next launch. An example is NASA's
space transportation system (Space Shuttle), which places satellites into low-Earth orbit
from which an intermediate rocket inserts the satellite into a GSO transfer orbit. Then the
ARM can be fired to achieve the final orbit. Since the Space Shuttle carries a human crew,
its costs are too high to be practical for the many commercial communication satellites
that need to be placed into orbit. It is reserved for launching special payloads or
performing special operations that require human intervention. New initiatives have been
reported about the development of small reusable launch vehicles (Kistler Co.) for operations in
the next decade.
Non-geostationary orbits (non-GSO)
Launch systems for low-Earth orbit (LEO) satellites usually require much lower booster
capabilities than for GSO systems and have shown greater flexibility in their designs. For
example, some LEO launch systems have been carried aloft in aircraft to improve their
payload delivery capabilities.
Others are designed to launch several satellites in a particular orbit or constellation, thus
reducing the number of launches and the overall costs. The basic design or vehicle of non-
GSO launch systems are similar to that for the GSO satellites when multiple satellites or large
payloads need to be inserted in non-GSO orbits. Rocket stages may be added or deleted
depending on the payload and orbit requirements.
19
Non-GSO launch systems have enjoyed a long period of operations reaching back to the
first earth satellite (Sputnik) in 1957. New developments to increase the reliability and reduce the
cost of these systems has continued so that, at present, there are several new or modified systems
available to the communication satellite industry. A few examples of LEO type launch
systems include Atlas I (United States), Aussroc (Australia), Capricornio (Spain), Delta
Lite (United States), ESA/CNES Series (Europe), J-Series (Japan), Kosmos (Russia),
Lockheed Astria (United States), Long March CZ-1 (China), PacAstro (United States),
Pegasus (United States), Sea Launch (United States/International), Shavit (Israel), SLV
Series (India), Soyuz/Vostok (Russia), and VLS Series (Brazil), among others.
Launcher selection
A preliminary review for the selection of a launch system would entail equating
the performance capabilities against such requirements as satellite system weight to be injected
into a specific orbit, the volume available in the nose cone or housing of the rocket, the injection
accuracy for transfer orbits or final orbit insertion, and other technical factors. An equally
important set of considerations is the reliability and costs of the launch system, including launch
services. In addition, transportation costs to the launch site need to be assessed as well as related
insurance fees.
The recent commercialization of the launch industry has introduced a high level
of competition among suppliers, both for governments and private organizations. The latest
information should be obtained in this highly dynamic environment before any commitment is
made for a particular launch system. New data services, such as the "Internet", and
technical journals, such as NASA's "Transportation Systems Data Book", provide general
information about the status of many launch systems and their manufacturers or distributors. For
up-to-date technical details and costs, suppliers should be contacted directly.
Current and future launch systems
This section provides some preliminary information on some of the recently
employed satellite launch systems and some of the modified systems reported in trade journals
and reports. It is not an exhaustive summary of all the launch systems that have emerged during
20
the last decade, but a brief view of some examples of launch systems that have operational
experience.
a) Ariane Series
At present, Ariane 4 is the most prominent launch system in the international commercial
satellite communication industry. This system was developed by the European Space
Agency (ESA) and CNES, the French Space Agency, and operations are conducted by
Arianespace. The 4 Series, which has a reliability of over 90%, is capable of inserting
from 1 900 to 4 200 kg into a geostationary transfer orbit (GTO). The Ariane vehicles are
launched from Kourou, French Guyana where the latitude is approximately 5° N. A larger
version, Ariane 5, has recently completed development and has become available for
commercial launches. A schematic of Ariane 5 is shown in Fig.
b) Atlas Series
These systems represent the larger of the commercial launchers in the United
States of America. Developed in the 1960s, the Atlas is currently operated for
commercial services by the Lockheed Martin Company and some Russian aerospace
companies. The Atlas 1 and 2 versions, which had a reliability close to 90%, are capable of
inserting from 2 250 to 3 490 kgs into a GTO. Fig. shows a schematic of the Atlas launch system
and lists the technical characteristics of its subsystems and components.
c) Delta Series
These systems have a long history of reliable operations (98%) in the United States.
Currently, they are manufactured and marketed by the McDonnell Douglas Company. The
Delta II version is capable of inserting from 950 to 1 820 kgs into a GTO. A Delta III
version is currently under development to more than double the lift capability of its
predecessors. Fig. shows a summary of the Delta's growth from its start in 1960 until the
present. Also shown are the present models intended for LEO applications.
d) H-Series
The H-2 launch system, which was developed by Japan from their earlier N-Series of
vehicles, has been successful in their initial flights in inserting heavy payloads into the
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GSO and space. Lift capability to the GTO is 4 000 kgs. Fig. shows a history of the
development of the H-Series of launchers by Japan.
e) LLV Series
The Lockheed Launch Vehicle is another flexible system that utilizes small solid rocket
boosters to increase its lifting capacity. The LLV-2 and -3 versions are able to place 1 305 and 2
500 kgs into a GTO. This system is another good candidate for launching LEO satellite systems
into orbit.
f) Long March Series
This system, developed by China, includes a range of vehicles from the small CZ-1D to
the large CZ-2E. These launchers are available for commercial satellite services. The range of lift
capabilities to the GTO vary from 200 to 3 370 kgs. This series of vehicles has plans to launch
LEO satellites within a few years, where the lift capability will be greater by a factor of 2 or 3.
Fig. below summarizes the characteristics of Long March launchers.
g) M-Series 442 CHAPTER 6 Space segment
The M vehicles, which have all solid propellants, were also developed by Japan,
but for smaller payloads. The M-V model has the ability to insert 1 215 kgs into the GTO. This
series is planned for a variety of space missions.
h) Proton Series
These systems, which were designed to lift very heavy payloads into space, have a long
history of operations in the former USSR. Presently, it is being marketed by International Launch
Services, a joint venture between Krunichev of Russia and Lockheed Martin of the United States
This vehicle has a lift capability of 5 500 kgs into a GTO. Fig. shows the major hardware
components of the Proton D-1 launch system.
i) Titan Series
This launch system was developed in the United States several decades ago as a ballistic
vehicle and was subsequently revised to insert heavy satellites into orbit. The Titan IV
22
version is capable of inserting from 6 350 to 8 620 kgs into a GTO. This system is primarily
employed for United States government operations.
j) Zenit Series
This launch system was modified from earlier USSR large lift vehicles and is
presently manufactured in the Ukraine by NPO Yuznoye. It is capable of lifting 4 300 kgs into a
GTO. In a joint venture with Boeing (United States) and Kvaerner (Norway), the Sea Launch
Company plans to increase this capability to 5 400 kgs by launching the Zenit from a modified
ocean oil platform located on the equator. A schematic of the Zenit system in shown in Fig. with
a sketch of the ocean platform under development by the Sea Launch Company's joint venture
programme.
With regard to other systems mentioned in this section, the reader is advised to seek out
information from the organizations mentioned above for up-to-date information.
23
Fig.10 Schematic of Ariane 5 launch system
LOOK ANGLES
The azimuth and elevation angles are referred to as the look angles for the ES to the
satellite. Fig. 17 shows the geometry and definitions of the look angles with respect to the earth
station reference.
24
Fig. 17 GSO look angles to satellite
There are many sources available in the orbital mechanics and satellite literature
that describe the detailed development of the calculations for the GSO parameters, range,
elevation angle, and azimuth angle. Two good examples are provided in References 1 and 2.
The calculations involve spherical geometry derivations and evaluations requiring several
stages of development. There are also several software packages available for the determination
of orbital parameters, for both GSO and NGSO satellites networks. Our intent here is to
summarize the final results of the various derivations and to allow us to apply the GSO
parameters to the evaluation of free space links for communications satellite applications.
The input parameters required to determine the GSO parameters are:
lE = earth station longitude, in degrees
lS = satellite longitude, in degrees
LE = earth station latitude, in degrees
LS = satellite latitude in degrees (assumed to be 0, i.e., inclination angle = 0)
H = earth station altitude above sea level, in km
The point on the earth’s equator at the satellite longitude is called the subsatellite point
(SS). Fig. 18 clarifies the definition of earth station altitude.
Fig. 18 Earth station altitude
Longitude and latitude sign values are based on the sign convention shown in Fig. 19.
Longitudes east of the Greenwich Meridian and latitudes north of the equator are positive.
Additional parameters required for the calculations are:
Equatorial Radius: re = 6378.14 km
25
Geostationary Radius: rS = 42 164.17 km
Geostationary Height (Altitude): h GSO = rS − rE = 35 786 km
Eccentricity of the earth: ee = 0.08182
An additional parameter required for the calculation of the GSO parameters is the
differential longitude, B, defined as the difference between the earth station and satellite
longitudes: B = IE − IS
where the sign convention of Fig. 12 is followed.
Fig. 19 Sign convention for longitude and latitude
For example, for an earth station located in Washington, DC, at the longitude of 770W,
and a satellite located at a longitude of 1100W:
B = ( − 77) − ( − 110) = + 330
LINK CALCULATIONS
Introduction
Fig. 20 Satellite Uplink and Downlink
Referring to Fig. 20, the overall performance of a one-way link between two earth
stations A and B depends on the characteristics of three elements: the uplink (A to
satellite), the satellite transponder and the downlink (satellite to B). This section explains the
calculation of the overall link budget for such a one-way satellite link. Of course, such a
26
calculation can be directly extended to the case of multiple access links. As the purpose of a link
budget is to calculate the quality of a satellite communication:
• in the case of analogue, frequency modulation transmissions, this quality is evaluated
by the signal-to-noise ratio (S/N);
• in the case of digital communications, this quality is measured by the information
signal bit error ratio (BER).
However, it should be emphasized that, in practical applications, an inverse
process is generally followed: for the transmission of a given signal between two earth
stations (or even between two user terminals) with given availability and quality requirements 9
, the final purpose of a link budget is to calculate the technical design parameters needed for
the signal (type of modulation, error correction encoding, etc.) and for the earth station
and, possibly, for the space station, i.e. the satellite (G/T, E.I.R.P., etc.). These technical
parameters determine the type of equipment needed (type and size of antennas, power of the
amplifiers, modems, codecs, etc.).
This is limited to the calculation of the factors (C/N0), which do not postulate the choice
of the transmission bandwidth (B) nor of the modulation and coding processes.The various
coding and modulation techniques that the conversion of the (C/N0) into (S/N) or into
BER will permit evaluation of the final transmission performance. Some indications on the
subject will be given at the end of the section.
Note also that this is only devoted to the basic formulas. Practical cases of link
budget calculations. Some basic relations between (C/N), (C/N0), (C/T) and (Eb /N0) are recalled
in Table 2 below.
Important: In all formulas and calculations below, small letters (lower case) are
used when numerical units are implied (with the following exceptions: T for the noise
temperature, B for the bandwidth occupied by the signal, R for the digital information signal bit
rate). Capital letters are used when decibels are implied.
Uplink (C/N0)u
In accordance with the formulas in, the power level received at the input of the satellite
receiver is given by:
CU = PE · GET · GSR /LU (W)
27
where:
PE: the output power of the earth station high power amplifier (HPA)
GET: the earth station antenna transmit gain in the direction of the satellite, whence:
PE . GET: the equivalent isotropically radiated power of earth station (A) in the direction
of the satellite, i.e.: (E.I.R.P.)E
LU: the free-space attenuation in the uplink
GSR: the satellite receiving antenna gain in the direction of the transmitting earth station
A, including losses in the feeder between the antenna output and the receiver. The carrier-
to-noise density ratio in the uplink is then given by:
(C/N0) U = (E.I.R.P.)E · GSR /(LU · KTU )
= (G/T)S · (E.I.R.P.)E /LU · k –1
where:
TU : equivalent noise temperature of the uplink at the satellite receiver input (K).
(G/T)S : figure of merit of the space station (K –1 ).
Formula can be rewritten as follows:
(C/N0) U = (G/T)S · (λ2 / 4π) · (E.I.R.P.)E /(4πd
2 ) · k –1
Here (λ2 / 4π) is the effective aperture area of an isotropic antenna
As (E.I.R.P.)E /(4πd 2) = (PFD)U , the power flux-density transmitted by the earth station
antenna at the actual distance (d) of the satellite this figure is often included in the
satellite specifications as the operating flux-density at the transponder input.
In consequence, another useful method of expressing formula (27) is:
(C/N0) U = (G/T)S · (λ2 / 4π) · (PFD)U · k –1
Formulas are often expressed in decibels (i.e. (C/N0) U = 10 log 10 ((C/N0) U), as:
(C/N0) U = (G/T)S + (E.I.R.P.)E – LU + 228.6 (29)
(C/N0) U = (G/T)S + 10 log (λ2 / 4π) + (PFD) u + 228.6 (dB·Hz)
Typical examples of L u are: 199.75 dB at 6 GHz and 207.1 dB at 14 GHz (for
a distance d = 38 607 km corresponding to a GSO satellite at 30° elevation). Typical examples
of the effective aperture area of an isotropic antenna (in dB, i.e. 10 log (λ2 / 4π)) are –37
dB(m2 ) at 6 GHz and –44.37 dB(m
2 ) at 14 GHz.
28
Downlink (C/N0)D
The level of the carrier received at the input of the earth station receiver is given by:
CD = PS · GST · GER /LD (W)
where:
PS : the output power of the satellite transponder amplifier
GST : the satellite antenna transmit gain in the direction of the earth station, whence:
PS · GST : the equivalent isotropically radiated power of the satellite in the
direction of the receiving earth station, i.e.: (E.I.R.P)S
LD : the free-space attenuation in the downlink
GER : the receiving earth station antenna gain, including losses in the feeder between the
antenna output and the receiver.
Hence, the carrier-to-noise density ratio in the downlink is:
(C/N0)D = (E.I.R.P.)S · GER /(LD · KTD)
= (G/T)E · (E.I.R.P)S /LD · k –1
where:
TD : equivalent noise temperature of the downlink at earth station receiver input
(K)
(G/T)E : figure of merit of the earth station (K –1 ).
Formula is often expressed in decibels as:
(C/N0)D = (G/T)E + (E.I.R.P.)S – LD + 228.6
Typical examples of LD are: 196.20 dB at 4 GHz and 205 dB at 11 GHz.
Link budget for a transparent transponder
The overall link budget calculation depends on whether the satellite is equipped with a
conventional transponder or a regenerative transponder. In the former case, the role of the
transponder is simply to amplify the uplink signal (with minimum distortion and noise).
This is the reason why it is often called a transparent transponder 10 .
29
In the latter case, the uplink (generally digital) signal from earth station A is
demodulated in the transponder, then regenerated (often after implementing some decoding and
baseband processing), re-modulated, amplified before being downlink transmitted to earth station
B. This subsection deals with transparent transponders while will deal with regenerative
transponders.
Combined uplink and downlink (C/N0)UD
The total (C/N0)UD of the link between the earth stations A and B, including only
thermal noise contributions is the ratio of the signal power to the total thermal noise power, at
the receiver input of B.
The signal power is: CUD = CU · G · GST · GER /LD
where G is the transponder gain.
The noise spectral density is the sum of the uplink and downlink
contributions, i.e.: N0 = N 0U · G · GST · GER /LD + N0D .
Therefore:
(C
N0)UD
=CU
N0 + N0DLD/GGSTGER
Now, since: G = PS /CU and CD = PS · GST · GER /LD it follows that: G = (CD /CU) · LD
/(GST · GER) and, after simplifications:
(C
N0)UD
=CU
N0U + N0D (CU
CD⁄ )
(C
N0)UD
−1
= (C
N0)U
−1
+ (C
N0)D
−1
SATELLITE USED FOE MOBILE NETWORKS
The satellites of the INMARSAT system currently provide a range of
communications services (voice, telex, fax and data) to different users using a variety of
terminals and applications (aeronautical, maritime or land). There are two fundamental
types of user earth stations in the INMARSAT system that carry traffic, namely, the land
30
earth stations (LES) – sometimes also referred to as coast earth stations (CES) –
operating in the 6/4 GHz band, and the mobile earth station (MES) operating in the 1.6/1.5
GHz band.
The FSS feeder-link gateway to the mobile earth stations is via the INMARSAT land
earth stations. As of November 1998, there were about 40 LESs distributed around the globe,
with at least one in every continent. A land earth station need not necessarily be located on a
"coast" but it does need to be located within the coverage beam of one or more
INMARSAT satellites. The INMARSAT antenna beams are designed to cover the three major
ocean regions.
Land earth stations are owned independently by telecommunications operators. An LES
operator is often, but not always, the signatory (the organization nominated by its government to
invest in and work with INMARSAT) of the country in which the LES is located.
The parameters of typical INMARSAT earth stations are given below:
Table 2 Receive system performance
Table 3 Transmit system performance
(for one telephone voice channel and Mini-M antenna)
Basic Elements(Components) of satellite communication system
Satellite communications are comprised of 2 main components:
31
• The Satellite
The satellite itself is also known as the space segment, and is composed of three separate
units, namely the fuel system, the satellite and telemetry controls, and the transponder. The
transponder includes the receiving antenna to pick-up signals from the ground station, a broad
band receiver, an input multiplexer, and a frequency converter which is used to reroute the
received signals through a high powered amplifier for downlink. The primary role of a satellite is
to reflect electronic signals. In the case of a telecom satellite, the primary task is to receive
signals from a ground station and send them down to another ground station located a
considerable distance away from the first. This relay action can be two-way, as in the case of a
long distance phone call. Another use of the satellite is when, as is the case with television
broadcasts, the ground station's uplink is then downlinked over a wide region, so that it may be
received by many different customers possessing compatible equipment. Still another use for
satellites is observation, wherein the satellite is equipped with cameras or various sensors, and it
merely downlinks any information it picks up from its vantagepoint.
• The Ground Station.
This is the earth segment. The ground station's job is two-fold. In the case of an uplink, or
transmitting station, terrestrial data in the form of baseband signals, is passed through a baseband
processor, an up converter, a high powered amplifier, and through a parabolic dish antenna up to
an orbiting satellite. In the case of a downlink, or receiving station, works in the reverse fashion
as the uplink, ultimately converting signals received through the parabolic antenna to base band
signal.
Working of Transponder
When used for communications, a satellite acts as a repeater. Its height above the Earth means
that signals can be transmitted over distances that are very much greater than the line of sight. An
earth station transmits the signal up to the satellite. This is called the up-link and is transmitted
on one frequency. The satellite receives the signal and retransmits it on what is termed the down
link which is on another frequency.
32
Using a satellite for long distance communications
The circuitry in the satellite that acts as the receiver, frequency changer, and transmitter is
called a transponder. This basically consists of a low noise amplifier, a frequency changer
consisting a mixer and local oscillator, and then a high power amplifier. The filter on the input is
used to make sure that any out of band signals such as the transponder output are reduced to
acceptable levels so that the amplifier is not overloaded. Similarly the output from the amplifiers
is filtered to make sure that spurious signals are reduced to acceptable levels. Figures used in
here are the same as those mentioned earlier, and are only given as an example. The signal is
received and amplified to a suitable level. It is then applied to the mixer to change the frequency
in the same way that occurs in a super heterodyne radio receiver. As a result the communications
satellite receives in one band of frequencies and transmits in another.
In view of the fact that the receiver and transmitter are operating at the same time and in close
proximity, care has to be taken in the design of the satellite that the transmitter does not interfere
with the receiver. This might result from spurious signals arising from the transmitter, or the
receiver may become de-sensitised by the strong signal being received from the transmitter. The
filters already mentioned are used to reduce these effects.
33
Block diagram of a basic satellite transponder
Signals transmitted to satellites usually consist of a large number of signals multiplexed onto a
main transmission. In this way one transmission from the ground can carry a large number of
telephone circuits or even a number of television signals. This approach is operationally far more
effective than having a large number of individual transmitters.
Obviously one satellite will be unable to carry all the traffic across the Atlantic. Further capacity
can be achieved using several satellites on different bands, or by physically separating them apart
from one another. In this way the beamwidth of the antenna can be used to distinguish between
different satellites. Normally antennas with very high gains are used, and these have very narrow
beamwidths, allowing satellites to be separated by just a few degrees.
Separating satellites by position
GPS SERVICES
Definition:
34
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. The system provides
critical capabilities to military, civil, and commercial users around the world.
GPS was built with military uses in mind during the Cold War. In 1983, Korean Air flight
007 was shot down by Soviet interceptors over Kamchatka when it went off-course. All
passengers and crew aboard the civilian flight, including a sitting US congressman, were killed.
Amid the ensuing controversy, President Reagan announced that GPS would be made available
for free for civilian use to avoid such preventable disasters in the future. So in essence, it took the
political momentum from a national tragedy for it to become freely available.
GPS provides two different positioning services: the Precise Positioning Service (PPS)
and the Standard Positioning Service (SPS).
Services:
1.The Precise Positioning Service (PPS)
Precise Positioning Service (PPS) is a positioning and timing service provided by way of
authorized access to ranging signals broadcast at the GPS L1 and L2 frequencies. The L1
frequency, transmitted by all Navstar satellites, contains a coarse/acquisition (C/A) code ranging
signal, with a navigation data message, that is available for peaceful civil, commercial, and
scientific use; and a precision (P) code ranging signal with a navigation data message, that is
reserved for authorized use.
The P-code will normally be cryptographically altered to become the Y-code. The Y-
code will not be available to users that do not have valid cryptographic keys. Navstar satellites
also transmit a second P- or Y-(P(Y)-) code ranging signal with a navigation data message at the
L2 frequency.
2.Standard Positioning Service (PPS)
Standard Positioning Service (PPS) is a positioning and timing service provided by way
of ranging signals broadcast at the GPS L1 frequency. The L1 frequency, transmitted by all
satellites, contains a coarse/acquisition (C/A) code ranging signal, with a navigation data
message, that is available for peaceful civil, commercial, and scientific use.
The Standard Positioning Service is based on the Coarse/Acquisition code (C/A(t)),
which is modulated only on L1. It has a chipping-rate of 1.023 MHz, and contains 1 023 chips,
so that the code is repeated every millisecond and each chip lasts about 1 µs, meaning a chip-
width or wavelength of 293.1 metre.
Segments In GPS
35
The current GPS consists of three major segments. These are the space segment (SS), a
control segment (CS), and a user segment (US).
The U.S. Air Force develops, maintains, and operates the space and control segments.
GPS satellites broadcast signals from space, and each GPS receiver uses these signals to
calculate its three-dimensional location (latitude, longitude, and altitude) and the current time.
1. The space segment is composed of 24 to 32 satellites in medium earth orbit and also includes
the payload adapters to the boosters required to launch them into orbit.
2. The control segment is composed of a master control station (MCS), an alternate master
control station, and a host of dedicated and shared ground antennas and monitor stations.
3. The user segment is composed of hundreds of thousands of U.S. and allied military users of
the secure GPS Precise Positioning Service, and tens of millions of civil, commercial, and
scientific users of the Standard Positioning Service.
APPLICATIONS
The applications of the Global Positioning System fall into five categories: location,
navigation, timing, mapping, and tracking. Each category contains uses for the military, industry,
transportation, recreation and science.
1. Location
This category is for position determination and is the most obvious use of the
Global Positioning System. GPS is the first system that can give accurate and precise
measurements anytime, anywhere and under any weather conditions. Some examples of
applications within this category are:
1. Measuring the movement of volcanoes and glaciers.
2. Measuring the growth of mountains.
3. Measuring the location of icebergs - this is very valuable to ship captains helping them to
avoid possible disasters.
4. Storing the location of where you were - most GPS receivers on the market will allow
you to record a certain location. This allows you to find it again with minimal effort and
would prove useful in a hard to navigate place such as a dense forest.
36
2.Navigation
It is the process of getting from one location to another. This was the what the Global
Positioning System was designed for. The GPS system allows us to navigate on water, air, or
land. It allows planes to land in the middle of mountains and helps medical evacuation
helicopters save precious time by taking the best route.
3. Timing
GPS brings precise timing to the us all. Each satellite is equipped with an extremely
precise atomic clock. This is why we can all synchronize our watches so well and make sure
international events are actually happening at the same time.
4. Mapping
This is used for creating maps by recording a series of locations. The best example is
surveying where the DGPS technique is applied but with a twist. Instead of making error
corrections in real time, both the stationary and moving receivers calculate their positions using
the satellite signals. When the roving receiver is through making measurements, it then takes
them back to the ground station which has already calculated the errors for each moment in time.
At this time, the accurate measurements are obtained.
5. Tracking
The applications in this category are ways of monitoring people and things such as
packages. This has been used along with wireless communications to keep track of some
criminals. The suspect agrees to keep a GPS receiver and transmitting device with him at all
times. If he goes where he's not allowed to, the authorities will be notified. This can also be used
to track animals.