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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 (a2b2) ] / 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.

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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

8

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

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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).

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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

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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

21

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.