Elements of satellite technology and communication

7/28/2019 Elements of satellite technology and communication

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Elements of satellite technology and communicationB Y O D D G U T T E B E R G

1 Introduction and

overviewMore than 35 years ago, on 4 October1957, the worlds first artificial satellitewas launched by the Soviet Union. Thesatellite, which became known as Sput-nik-I, opened up a new era of practicaluse of the outer space. The satellitesweight was 84 kg and it circled 1400times around the earth before burning upin the atmosphere after 93 days. Thelaunch of Sputnik-I was followed by theUnited States Explorer-I in January1958. Even though these satellites wereprimarily not intended for communica-

tions, they demonstrated that this wastechnically and economically feasible.

The use of artificial satellites in earthorbits is now a well established and inte-grated part of the worlds telecommuni-cations network. The evolution of thesatellite technology together with morepowerful launchers have made the satel-lites suitable, not only for long distancecommunications, but also for nationalcommunications, television broadcastingand mobile services.

A satellite communication system isdivided into two major parts, the earth

segment and the space segment. Thesatellite and its control station form thespace segment, while the earth segmentcomprises the traffic and traffic controlstations, see figure 1.

The satellite control station (TT&C sta-

tion) maintains the satellite in orbit. Itkeeps control of the satellite status(Telemetry), orbital information (Track-ing), and performs orbital attitudemanoeuvres and configures the commu-nication system (Command).

The communication part of the satellitesystem is simply a radio system withonly one relay station, the satellite. Thesignals are transmitted on a carrier fre-quency,fA, from an earth station (trafficstation), received in the satellite, ampli-fied, shifted in frequency tof 'Aand trans-mitted back to the receiving traffic sta-tion. The satellite transmits and receiveson radio frequencies mainly in themicrowave band, i.e. 3 - 30 GHz. Onthese frequencies it is feasible to useparabolic reflector antennas. Such anten-nas concentrate the radio signals in asmall cone; thus it is possible to illumi-nate the whole earth or part of the earth,see figure 2. The illuminated part iscalled the coverage area, and most of thetransmitted power from the satellite isconcentrated in this area. The larger thetransmitting satellite antenna is, thesmaller the coverage area becomes.

The up-link signals have only one desti-

nation, the orbiting satellite, whichmeans that it is required to transmitenergy only to the satellite. The transmit-ting earth station therefore normally con-

sists of a large antenna and a high power

amplifier. Since it is expensive and com-plicated to generate an equally highpower in the satellite, the output powerfrom the satellite is generally muchsmaller. Due to this, the receiving earthstations have to use larger receivingantennas; up to 30 metres in diameter, inthe early days of satellite communica-tions.

However, due to larger satellites, i.e.higher output power, and to better receiv-ing technology, the earth stations can usesmaller and cheaper antennas, e.g. satel-lite TV broadcasting.

2 A brief historical review

The first person to suggest the use of thegeostationary orbit for communicationsatellites, was the English physicistArthur C Clark (born 1917). He pub-lished his article Extra-TerrestrialRelays in the Wireless World Magazinein October 1945. In this article a satellitesystem for broadcasting of television,was described. At that time there was adiscussion going on about how to dis-tribute television. The present technologywas not mature for the construction of

reliable and unmanned satellites. In 1945no-one could predict the rapid develop-ment within electronics, e.g. the inven-tion of the transistor in 1947.

Satellite

Traffic

station A

Satellite control

station (TT & C)

Traffic

station B

up-li

nk do

wn-lin

k

fA

fB

fB

fA

Figure 1 The main building blocks in a two-way satellite com-munication system

Figure 2 A geostationary satellite illuminating the coveragearea on earth

621.396.946


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The Space Age started with the laun-

ching of the first artificial satellite,Sputnik-I, in 1957. It is worth whilenoticing that the first trans-Atlantictelephone cable was stretched thesame year. In the following years seve-ral satellites were launched, both forscientific purposes and for other speci-fic applications. Communication expe-riments with passive reflecting satel-lites, Echo-I and -II (1960), were alsocarried out.

An operational system with passivesatellites was never realised, but theproject made essential contributions tothe development of the earth stationtechnology.

The first active telecommunicationsatellites were launched in the early1960s, Courier (1960), Telstar-I and -II(1962 and 1963) and Relay-I and -II(1962 and 1964).

They were all put into low earth orbits,

the height varying between 1,000 and8,000 km. Telstar-I was the mostimportant of these satellites, figure 3.

This satellite transmitted televisiondirectly from USA to England andFrance for the first time. The Nordiccountries started early to make use ofsatellite communication. In 1964 theNordic countries built a commonreceive earth station at R, outsideGothenburg. It was an experimentalearth station which participated inNASAs experiments with the Relaysatellites.

The first satellite in the geostationaryorbit was Syncom-II in 1963. Thelaunch of Syncom-I the same year wasa failure. This was an important break-through for commercial satellite com-munication. Several important experi-ments were carried out. One of them

Launch data

Date:

Vehicle:

10 July 1962

Thor-Delta

Orbital data

Orbit:

Inclination:

Apogee:

Perigee:

Period:

Elliptical

44.8 degrees

5653 km

936 km

157.8 min

Satellite data

Weight:

Diameter:

No of transponders:

Bandwidth:Output power:

77 kg

0,9 meter (spherical)

1

50 MHz3 Watts

Figure 3 The Telstar satellite, main data.The satellite was designed and built by BellTelephone Laboratories and launched by NASA

INTELSAT I

(Early Bird)

1965

240

40

1

INTELSAT II

1967

240

85

3

INTELSAT II I

1968

1 200

150

5

INTELSAT IV

+ IV A

1971

6 000

790

7 + 5

INTELSAT V

1980

12 000

1 040

13

INTELSAT VI

1989

36 000

2 230

6

INTELSAT VI I

1993

18 000

1 400

12

Telephonechannels

Weight inorbit (kg)

No. ofsatellites

Figure 4 The INTELSAT satellites. Technical development


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was to examine whether or not geosta-

tionary satellites could be a part of thepublic telephone network. The problemwas the time delay of the radio signal.With a mismatch at the other end of theconnection, one would hear oneself half asecond later. This was thought to be amajor obstacle for a two-way telephoneconversation. Effective echo cancellerswere then developed. This reduced theproblem to such a degree that the tele-phone users accepted a connection viasatellite.

In 1965, the first commercial geostation-ary satellite, Early Bird, was placed in aposition over the Atlantic Ocean. Thisbecame the first satellite in the interna-tional organisation for telecommunica-tion satellites, INTELSAT. A prelimi-nary agreement was made in 1964between 11 countries, but the final agree-ment was not signed until 1973. Thisorganisation has been the leading one inthe development of communication satel-lites. Over 100 countries are members,and there are almost 600 earth stations in170 countries. INTELSAT has launched35 geostationary satellites and only sixhave been unsuccessful.

Figure 4 shows the development of the

INTELSAT system.

The first Nordic INTELSAT earth stationwas built in 1971 at Tanum in Bohusln.This earth station communicates with

geostationary satellites over the Atlantic

and Indian Oceans. The first Europeancommercial communication satellite,EUTELSAT-I, was launched in 1983.

3 Applications

The most important application of satel-lite communication has up till now beeninter-continental telephone and TV trans-missions. This was the first applicationand is still the most important. Large andexpensive earth stations with antennas upto 20 m in diameter are employed. TheINTELSAT-A system is an example ofthis.

The rapid development in the satellitetechnology and the use of more powerfullaunch vehicles has led to the use ofsatellite systems in more restricted areas.As an example the Norwegian Telecomemploys satellites as a part of its domes-tic telecommunication network andoffers satellite connections between theNorwegian mainland and the oil installa-tions in the North Sea, figures 5 and 6, aswell as to the Arctic islands of Svalbard(the NORSAT-A system). This systemtransmits via one of the satellites in theINTELSAT system.

The main station in the NORSAT systemis located at Eik in the south western partof Norway. It was put into operation in1976 and was the first earth station inEurope intended for domestic traffic.

Svalbard was connected to the terrestrial

public network via this system in 1979with an earth station at Isfjord Radio, fig-ure 7. The TV broadcasting to Svalbardwas established via one of the EUTEL-SAT satellites in 1984.

Regional systems can make use of satel-lites with global, semi-global or spot cov-erage. This can be achieved by largersatellite antennas. The EuropeanEUTELSAT system is a satellite systemwith several application areas. The devel-opment of this system started in 1970.The system intended to

- transmit telephony traffic between

large national earth stations

- exchange TV programmes (Eurovi-sion)

- cover the need for special services (asbusiness communication)

- distribute TV programmes to cablenetworks.

The first satellite system for maritimemobile communication, MARISAT,became operative in 1976. This system isnow replaced by the INMARSAT sys-tem. The first operational European

coastal earth station was put into serviceat Eik in Norway in 1981. This stationwill now expand to cover satellite com-munication with aeroplanes

ISFJORD RADIO (Spitzbergen)

EIK(main station)

STATFJORD

FRIGG

EKOFISK

ca 200 km

ca20

0km

ca300

km

Oslo

Stavanger

Bergen

Trondheim

Troms

INTELSAT-IV

Figure 5 The NORSAT-A system started in 1975. The initial sys-tem used leased capacity on an INTELSAT-IV satellite

Figure 6 NORSAT-A earth station on an oil rig in the North Sea


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Figure 7 The earth station at Isfjord Radio, Spitzbergen, 78North. At this station the maximum elevation angletowards a geostationary satellite is 3.1

Figure 9 Nittedal earth station outside Oslo, Norway

Coastalearth station

Large vessels Small vessels

Large aircraftsLand mobile

(trucks)

Land mobile(cars/persons)

Small aircrafts

Necessary reduction in terminal cost

Increasing number of terminals

Figure 8 Mobile satellite communication system


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(INMARSAT AERO). The evolution

points at communication with evensmaller units such as trucks, privatecars and individuals, figure 8.

As the satellites become more power-ful it is possible to make simpler earthstation equipment. Then the satelliteterminals can be located closer to theuser. This is done in the VSAT sys-tems which are mainly used for busi-ness communication. The Norwegian

Telecom has established an earth sta-tion at Nittedal, 1987, for Europeandata communication via the EUTEL-SAT system, figure 9, as well as a traf-fic control station at Eik for businesscommunication, the NORSAT-B sys-tem. The NORSAT-B system offersconnections between unmanned earthstations in Europe.

The last application area is broadcast-

ing of television and sound. This is apoint-to-multipoint system where wemake full use of the advantage of satel-lite communication. One satellite ingeostationary orbit will cover 97%ofall households in Norway. Today morethan 100 television programmes aretransmitted via satellite to Europeancountries. Arthur C Clarkes originalidea from 1945 has become a reality.Figure 10 shows the total involvementof the Norwegian Telecom in satellitecommunication.

4 Orbital considerations

A satellites period of revolution isdetermined by the distance from theearth. The farther from earth, thelonger the period. The nearest satel-

lites, e.g. the space shuttle, has an

orbital period of approximately 1.5hours. An example of a satellite with along period of revolution is the moon.Due to the long distance from earth(ca 385,000km) the period is about27days.

The orientation of the orbit withrespect to the earths equatorial planemay also be different, i.e. inclinedorbits.

The movements of a satellite are de-termined by the laws of Newton (1642

- 1727) and Kepler (1571 - 1630).Applied to satellites Keplers laws are:

1 The satellite orbit is an ellipse withthe earth at one focus.

EIK NORSAT - A INMARSAT NORSAT - B

INTELSAT INMARSAT EUTELSAT TELE-X THOR OIL RIGS (1976, 1984)

SHIPS(1982)

SVALBARD (1979, 1984)

NITTEDAL EUTELSAT TELE-X THOR

Satellite broadcastingSatellite business communication

(1984)(1986)

(1986)(1989)(1992)

(1976)(1982)(1986, 1989)

Figure 10 The Norwegian satellite involvement; main earth stations and utilized satellites


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9

v

Fc

Fg

m

M

r

= velocity of satellite

= centrifugal force

= gravitational force

= mass of satellite

= mass of earth

= orbital radius

v

FcF

gm

M

r

Perigee

Earth

equator

Satellite

Equatorial

Plane

Descendingnode

(North Pole)

Z

Y

X

i

Ascendingnode

36000k

m

Satellite 3

Satellite 1 Satellite 2

12 0o

120o

120o

Earth

Geosta

tion

ary

orb

it

North Pole

VA

36 000 km

200 km

VP

low earth orbit

ellipticaltransferorbit

geosynchronousorbit

Earth

Equatorial

plane

inclination

N

transfer

orbital

plane

Vernal equinox(21 March)

N

N

N N

Winter solstice(21 Dec)

Autumnal equinox(21 Sept)

Sommer solstice(21 June)

X

Interseeting line between the earthsequatorial plane and the ecliptic

The ecliptic(the earths orbitaround the sun)

Sun

PerigeeApogee

Earth

Satellite

v

b

b

a a

c

p

FF

r

Figure 11 Forces acting on a satellite in a circular orbit

Figure 13 The geocentric equatorial coordinate system. Thez-axis coincides with the rotational axis of earth.

The x,y-plane cuts through the earths equator andis called the equatorial plane. The x-axis is thedirection from the centre of earth through the centreof sun at the vernal equinox ( 21 March), seefigure 14

Figure 14 Definition of the x-axis in the geocentric equatorialcoordinate system

Figure 12 The elliptical satellite orbit with major and minorsemiaxis a and b and eccentricity e =c/a. Thesemilatus rectum p =h2/ where h =v r (theangular momentum per unit mass)

Figure 15 Three satellites in the geostationary orbit willalmost cover the whole earth

Figure 16 a) and b) Orbits for launching of geosynchronous satelli te. The incli-nation of the transfer orbital plane is equal to or larger than the lati-tude of the launch site


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10

where

p=semi-latus rectum =

e= =eccentricity (o


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5 Launching

For a satellite to achieve the geostation-ary orbit, it must be accelerated to 3.075km/s at a height of 35 786 km with zeroinclination. Theoretically, one can placethe satellite directly into the geostation-ary orbit in one operation. However, con-siderations regarding costs and launchingcapabilities call for a multistage launchvehicle (2 - 4 stages). The launching ofthe satellite represents a major part of thetotal investments in a satellite system.

The conventional, and most economical,method of launching a satellite is basedon the use of the Hohmann transfer orbit,

figure 16. For the following reasons, thelaunch site should be near equator:

(i)To make maximum use of the surface

velocity of the earth

(ii) To achieve minimum inclination.This will give minimum velocityincrement for the correction of theinclination. The minimum inclinationachievable is equal to the latitude ofthe launch site.

The launch sequence is as follows: Thesatellite is placed into low earth orbitwith an altitude of about 200 km, e.g. bythe space shuttle. There must be twovelocity increments; one for injectioninto the elliptical transfer orbit (Vp ) by

the perigee kick motor and one for injec-tion into the geosynchronous orbit (VA)by the apogee kick motor. The transferorbit has an altitude of the perigee andapogee corresponding to the transfer andgeosynchronous orbit altitudes, respec-tively.

Alternatively, one can put the satellitedirectly into the elliptical transfer orbit,by a multistage launcher, e.g. Ariane, seefigure 17. The idea of using a multistagelauncher is to get rid of the unnecessarymass, to achieve the sufficient velocity.

Correction of the inclination has to be

done at one of the nodes of the transferorbit, in order to achieve an equatorialcircular synchronous orbit. The velocityincrement is depending on the inclinationand on the velocity of the satellite. Thelower the satellite velocity, the more eco-nomical the manoeuvre. The circulariza-tion manoeuvre should therefore be car-ried out at apogee. This implies that theorientation of the transfer orbit has to besuch that the line between the apogee andperigee is in the equatorial plane, cfr. fig-ure 13. The altitude of the apogee has tobe equal to the altitude of GEO orbit. Themagnitude (V

A) and direc-

1 1

Figure 17 Multistage launcher ArianeFigure 18 Velocity diagram on a plane normal to the line of

nodes

Apogee

(Ascending node)

i = inclination

VS

VA

VAVSVA

= satellite velocity at apogee

= synchronous velocity = 3,075 km/s

= required velocity increment

= direction of impulse

VA

Figure 19 Apparent movement of satellitein an inclined synchronous orbit

Nodes(equator crossing)

i

i

S

W

Lattitude

N

E Longitude

Despunforwardthermal barrier

Forwardsolarpolar

Solarpanelextensiondrive (3)

AFTsolarpanel

Solarpanel

extension rack (3)

Primarythermalradiator

AFTthermalbarrier

Spinning/despun lock (4)

Radialthruster (2)

Axialthruster (2)

Reflector

Antennadeployment

and positioningmechanism

Antenna supportbeam

Transmit/receivefeed horn andassembly

Auxiliaryshelf

Despunpayload

compartment

Telemetry andcommand antennas

BAPTA

Earth sensor (2)

Spinningsection

Propulsiontank (4)

Apogee motor

Figure 20 Exploded view of the THOR satellite


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tion () of the velocity vector can be

calculated from figure 18. If the geo-synchronous orbit has some inclina-tion, the satellite will apparently movein a figure-of-eight with respect to thenodes, figure 19. The maximum excur-sion from equator in the north orsouth direction will be equal to theinclination.

6 The spacecraft

The main purpose of a communicationsatellite is to receive and transmitradio signals within the coverage areaon the earth. This means that thespacecraft should be a reliable andstable platform for the communicationsystem. The satellite is generally divid-ed into two main modules:

1 The communication module (or thepayload) which includes

- repeaters (or transponders)

- antennas

2 The service module (or the bus orplatform) comprising the followingsubsystems

- attitude and orbit control (AOCS)

- telemetry, tracking and command(TT&C)

- power supply.

Figure 20 shows an exploded view of atypical spacecraft.

6.1Attitude and orbital controlsubsystem

The objective of the attitude controlsubsystem is to maintain the commu-nication antennas correctly pointedtowards the earth, and the solar cellscorrectly pointed towards the sun. The

movements of the satellite about itscentre of mass can be described byrotations about the three orthogonalreference axes: roll (x), pitch (y) andyaw (z), see figure 21.

There are two different methods ofcontrolling the attitude: spin stabilisa-tion and body-fixed stabilisation (3-axis stabilisation), see figure 22. In thefirst type, the satellite spins around itsmain axis of symmetry. The rotation isnormally about 60 rpm. This will pro-duce an angular momentum in a fixeddirection. For a geostationary satellite,the spin axis (pitch) has to be parallelwith the earths axis of rotation. Tomaintain the pointing of the communi-cation antennas towards the earth, the

12

N

Sun

S

Earth

yaw

roll

Orbit

X

X

Y

pitch

line

normal to orbitalplane (south)

in orbitalplane

local vertical

Solarpanel

Antennapointingtowardsearth Despun

antenna

section

Spinningsection

Solarpanel

Pitch axis

Pitch axis

Solarpanel

Antennapointingtowards

earth

Reactionwheels

Figure 21 Definition of the reference axes, roll, pitch and yaw

Figure 22 a) and b) The two concepts of stabilization


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platform containing the antennas has

to rotate in the opposite direction (de-spun). Pitch correction is done byvarying the angular velocity of the des-pin motor, since the rate of change ofangular momentum is proportional tothe torque. Yaw and roll correctionsare made by thrusters mounted atappropriate places on the body. A typi-cal spin-stabilised satellite is shown infigure 20.

To keep the gyroscopic stiffness in abody-fixed stabilised satellite, one mayuse a momentum wheel inside thesatellite. This gives rather moderate

antenna pointing. Instead, it is usual touse three reaction wheels spinningaround the three main axes. While themomentum wheel is spinning at highvelocity (about 10 000 rpm), the reac-tion wheels are spinning in both direc-tion around zero speed. By varying thespeed and direction of the reactionwheels, controlling torques can beapplied around all axes. When thewheels reach their maximum speed,they must be unloaded by operatingthrusters on the satellites body.A typical body-stabilised satellite isshown in figure 2. For operating the

torque units, it is necessary to obtainthe attitude of the satellite. This isdone by sensors on board the space-craft sensing the direction to the sunand earth. As seen from the geostatio-nary orbit the sun subtends only 0.5.It is therefore a good source for ob-taining attitude reference. The sunsensor consists of photo cells which

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Earth

equator

bolometerbeams

t1

t2

despuncommunication

antenna

Communication

(spot) antenna

Commandomni antenna

Telemetryomni antenna

Commandreceiver

DecoderCommand

output

Telemetrytransmitter

EncoderTelemetry

input

Ranging

tones

Figure 23 Earth contour detection by scanning infrared sensors (bolometers) = spin rate of satellite2 = angle between the two bolometer beams

= error anglet1 andt2 =time between the horizon crossings

Figure 24 The satellites TT&C system

Earth

N

S

Orbital plane23,4

o

23,4o

Sun in

autumn

and spring

equinoxes

Sun in

summer

solstice

Sun in

winter

solstice

Figure 25 Apparent motion of the sun relative to earth


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produce an electric current when illu-

minated by the sun.The earth, seen from the satellite, hasa much larger view angle (17.3) andits centre cannot directly be mea-sured. However, using infrared detec-tors (bolometers), one can measurethe edge of the earth, due to the diffe-rence in radiation from the cold skyand the warm earth. Infrared detectorscan be used both day and night. If twobolometers are mounted on a spinningspacecraft, as shown in figure 23, thepointing error () can be calculatedmeasuring the time between the twohorizon crossings (t

1and t

2), knowing

the spin rate () and the angle be-tween the bolometer beams (2).

On a 3-axis stabilised satellite scan-ning mirrors have to be used, sincethe satellite itself does not spin.

A higher degree of pointing accuracymay be obtained by using a radio bea-con on the earth. In such a system thesatellite antenna is directly locked tothe transmitted radio signal from theearth beacon.

6.2Telemetry, tracking and

command (TT&C)The satellite is supervised and control-led via a dedicated earth station,

TT&C station, which in turn is connec-ted to the satellite control centre(SCC). The main tasks of the space-craft management is to control theorbit and attitude of the satellite, moni-tor the health status, remaining pro-pulsion fuel and transponder configu-ration, together with steering of anten-nas.

The satellites TT&C system is shownin figure 24.

The telemetry subsystem collects datafrom various sensors on board thespacecraft, such as fuel tanks pres-sure, voltages/ currents in differentsubsystems, sighting from infraredand sunlight detectors, temperature,etc. The information rate is usuallyless than 1kbit/ s. The data are trans-mitted on a low-power telemetry car-rier using an omnidirectional satelliteantenna during the transfer phaseand/or a spot beam antenna when thesatellite is on station.

The command subsystem is used forremote control of the different func-tions in the satellite. This may includeattitude and orbital manoeuvres, swit-

ching transponders on and off, stee-

ring antennas, firing the apogee boostmotor in transfer phase, etc. Whenreceiving a command, the commandsubsystem generates a verification sig-nal, which is sent back via the tele-metry link. After checking, an executesignal is transmitted to the satellite.

This prevents inadvertent commands.As in the telemetry link the omniantenna is used during the transferphase, while spot-beams are usedwhen the satellite is in its geostatio-nary position.

The tracking or ranging subsystemdetermines the orbit of the spacecraft.

There is a number of techniques thatcan be used. One method is to mea-sure to pointing of the TT&C earthstation antenna towards the satellite,together with the slant range to thesatellite, i.e. the range vector. Therange to the satellite can be measuredby modulating the command carrierwith multiple low frequency tones(tone ranging). These tones aredemodulated and remodulated on thetelemetry carrier and received in the

TT&C ground station. By comparingthe phase between the transmitted

and received tones, the range can becalculated. The highest tone gives thebest accuracy, but several lower tonesresolve the ambiguity.

6.3Power supply

Solar energy is the only externalenergy source for orbiting satellites.

The solar radiation intensity is about1.4kW/m2. The sun energy is conver-ted to electrical energy by means ofphotovoltic cells (solar panels). Theefficiency of the solar panels is about10-15%. On the spin-stabilised satelli-tes the solar panels form the exteriorof the spacecrafts body, see figures 20and 22.

One disadvantage of the spin-stabilis-ed satellite is that only 1/ 3 of totalsolar array is exposed to the sunlight.On the body-stabilised satellite thesolar panels are mounted on twodeployable wings, as shown in figure2 and 22. The satellite moves aroundthe earth in 24 hours. For interceptingmaximum solar flux, the solar panelwings have to make one turn per day.

They are driven by two separatemotors.

A measure of efficiency is the ratio ofproduced electrical power to the massof satellite. For a spinning satellite this

ratio is about 10 Watts/kg, whereas

for the 3-axis stabilised satellite theratio is about 50 Watts/kg. Thismeans that if we need a satellite withhigh RF output power, we shouldselect a body stabilised satellite. Thedisadvantage is that they are morecomplex.

On-board batteries are needed to pro-duce power during the launch phaseand when the satellite enters the sha-dow of the earth (eclipse). This hap-pens around the equinoxes, and themaximum duration is approximately70 minutes. This is due to the fact thatthe earths equatorial plane is inclinedwith an angle of 23.4 with the direc-tion to the sun, see figure 25. Duringthe summer and winter seasons thesatellite is out of the earths shadow,figure 26, whereas during the equino-xes the satellite passes through theearths shadow once a day for 42 days,figure 27, i.e. the satellite will experi-ence 84 eclipses per year.

6.4The communication module

The communication subsystem is theprimary system in the satellite. Allother functions in the satellite can be

considered as supporting activities forthis system. It consists of the receiv-ers, transmitters and the communica-tions antennas. One receive/ transmitchain is called a transponder. A satel-lite with 5 transponders is shown sche-matically in figure 28. The buildingblocks are:

- A wide band receiver and a downconverter

- An input multiplexer (IMUX)

- 5 channelised sections including thehigh power amplifiers

- An output multiplexer (OMUX).

The wide band receiver/ converteroperates in the whole up-link band,e.g. in one of the common bands

- C-band (5.9 - 6.4 GHz)

- Ku-band (14 - 14.5 GHz)

- Ka-band (17.3 - 18.1 GHz)

or other bands allocated to satellitecommunication.

The frequency conversion could eitherbe single, as in figure 28, or dual. Inthe latter case the first mixer convertsthe frequency down to an intermediatefrequency (IF). After the channel

14


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amplification on IF the signal is con-

verted up to the down-link frequency,before the last power amplification.This is often done because it is mucheasier to make filters, amplifiers andequalisers at an intermediate fre-quency. The intermediate frequency

can be chosen independently of the

up- and down-link frequencies.The input multiplexer separates theup-link band into individual channelswith a certain bandwidth. For examplecan a 500 MHz wide up-link band beseparated into 12 channels with chan-nel bandwidth of 36 MHz. The chan-

nel bandwidths may also be unequal.

The channel amplifier/ attenuator setsthe gain (gain-step) of the transponderin order to control the input back-off ofthe TWTAs. This can be done fromthe TT&C station by the up-link com-mand system.

The last high-power amplifier (HPA) isusually a travelling wave tube ampli-fier (TWTA). This amplifier establis-hes the level of the output power,which may be in the range of 10 - 200Watts depending on the application.

The TWTA operates near saturation.The input/ output characteristic isshown in figure 29. The output multi-plexer combines all the channelsbefore signals are transmitted to theearth by the antenna.

6.5The satellite communicationantenna

Antennas are used for receiving andtransmitting modulated radio frequentsignals from and to the earth. Themost commonly used antenna in themicrowave frequency band is the para-bolic reflector antenna. The antenna ischaracterised by its radiation patternand the maximum gain, that is the

antennas capability to concentrate theenergy in one direction.

The design of the satellite antenna isdepending on required coverage area,

15

Figure 27 The satellite passes through the earth shadow at the equinoxes once a day

Figure 26 During summer and winter the satellite is always illuminated

Day

Sun

Geostationary orbitEarth shadow

Night

Earth shadow

Night

Day

Sun

Geostationary orbit

Up-link

frequency f1

Receiving

antenna

input filter

Lownoise

amplifier

mixer

Preampli-fier

Local oscillator(f=f1-f2)

A

B

C

D

E

Transmit

antenna

Down-link

frequency f2

Channelfilter

AutomaticGain Control/Gain Setting

Spare

TravellingWave Tube

amplifier

TWT

TWT

TWT

TWT

A B

CD E

Wide band low

noise receiver withdown converter

Input

multiplexer

Channel

amplifiers

High power

amplifiers

Output

multiplexer

Figure 28 A satellite with 5 transponders and single frequency conversion


13/18

see figure 30. This determines the

beam width of the antenna, usuallytaken as the half-power beam width(or 3 dB beam width). The 3 dB beamwidth is given by

(degrees) (6.1)

where

K=65 - 75, depending on theantenna aperture field distribu-tion

=wavelength

D =diameter of antenna.

The gain of the antenna is proportionalto its area, and is given by

(6.2)

where

=antenna efficiency (0.5 - 0.8)

A =area of antenna

or in dB:

(6.3)Obviously the gain and the beamwidth of a reflector antenna are rela-ted. The gain is approximately givenby

(6.4)

where

3dB is given in degrees.

Should the service area be the wholeearth, the required beam width wouldbe 17.4 as seen from the geostatio-nary orbit. This corresponds to a satel-lite antenna with 20 dB gain. By in-creasing the size of the satellite anten-na only a small area of the earth is illu-minated. These narrow beams are cal-led spot beams. If we should coveran area on earth with a diameter of 300km, the beam width of the antennashould only be 0.5, with a correspon-ding gain of 50 dB. This means that weneed less power from the HPA in thespot beam to achieve the same fluxdensity on earth as in the global beam.

To express the transmitted power of asatellite, the Equivalent Isotropic Radi-ated Power (EIRP) is used. This issimply the product of output power

G =30000

3dB2

G(dB) = 10log D

2

G = 4

2A =

D

2

3dB = K

D

16

Output power (dBW)

Pma xSaturated output

Output back-off

Input back-off

Input power (dBW)

P inat saturat ion

45 cm 65 cm 210 cm

EIRP: 49 dBW54 dBW

antennadiameter

39 dBW

N Coverage area = 3dB

equator

Subsatellitepoint17,

4o

Side lobes

Main lobe

Antenna radiation pattern

Figure 29 Input/output characteristic of a TWTA

Figure 31 The preliminary EIRP contours of the THOR satellite and the correspon-ding receiving antenna diameter for TV reception with good quality

Figure 30 Coverage area of the satellite antenna. For global coverage the requiredbeam width is 17.4


14/18

fromHPA(Pt) and satellite antenna

gain (Gt):EIRP =Pt Gt (6.5)

So instead of plotting coverage con-tours (or gain contours) we can plotEIRP contours. An example is given infigure 31.

A satellite can have more than oneantenna. Modern satellites employseveral antennas, both with global,hemispheric and spot coverage.

As described, the communication pay-load overall can be characterised bythe following parameters:

- the coverage area

- the maximum EIRP

- the input power flux density for satu-ration of the output high poweramplifier

- the channel arrangement or fre-quency plan.

7 The satellite link

7.1The link geometry

In order to communicate with a satel-lite, it is necessary to know the correctpointing of the earth station antenna.We shall limit ourselves to geostatio-

nary satellites. In that case, as previ-

ously mentioned, the satellite will notmove with respect to an earth station,and the antenna pointing can be fixed.

The pointing of the antenna is definedby two angles, azimuth () and eleva-tion ().The azimuth is the anglefrom true north to the projection of theline to the satellite in the local horizon-tal plane. The elevation is the angleabove the local horizontal plane to thesatellite.

Knowing the satellites position, i.e.the longitude, and the earth stationposition, i.e. longitude and latitude, we

can calculate the earth station anten-nas pointing angles. The geometry isshown in figure 32.

With the terms given in figure 32, theazimuth, elevation and distance tosatellite are given by:

(7.1)

(7.2)

(7.3)d= Rs

1 coscos( )2

cos

= arctgcoscos Re

Rs

1 coscos( )2

= arctg tg

sin

The maximum latitude coverage of a

geostationary satellite (=0, =0)will be

=81.3 N or S

7.2 Link analysis

The purpose of a satellite system is toprovide reliable transmission with aspecified quality of the received signal.

The transmitted information has to bemodulated on an RF carrier. In analo-gue systems, where frequency modu-lation (FM) is the dominating modula-tion method, the signal-to-noise ratio

(S/ N) after the demodulator is a mea-sure of signal quality. In digital satel-lite links the measure of quality is thebit error rate (BER). The modulationmethod most often used in digital sys-tem is phase shift keying (PSK).

In both analogue and digital systemsthere is a unique relationship betweenthe carrier-to-noise ratio (C/ N) andthe signal-to-noise (S/N) ratio or thebit error rate (BER). Given the modu-lation method, the performance of atotal link is generally specified interms of a minimum C/ N in a certain

percentage of time. In order to estab-lish the link quality, we therefore needto calculate the carrier power (C) andthe noise power (N) at the receivingstation.

max =

2 arcsin

Re

Rs

17

E

Re

Rs

SRO R

e/cos cos

cos cos

/2--

cos =

=

N

S

Equator

O

Zenith

EEarth Station

Re

Rs

d

Satellite

Q S

R

Localhorizontal

plane(ERQ)

Sub

satellite

point

Longitude of

earth station

Lattitude of

earth station

Figure 32 a) and b) Calculations of azimuth, elevation and distance to a GEO satellite


15/18


16/18

Instead of giving the noise tempera-

ture of a component, the noise figure(F) is often used. This is defined asthe signal-to-noise ratio at the inputdivided by the signal-to-noise ratio atthe output:

(7.15)

If the source noise temperature is atthe standard temperatureTo =290 K,then

(7.16)

whereg=gain of the network.

Using the definition of effective inputnoise temperature:

(7.17)

or

T =(F - 1) 290 (7.18)

A satellite system consists of both anup-link and a down-link. Noise on the

up-link will also contribute to the over-all noise power received at the earthstation. The basic satellite link isshown in figure 34.

The total received noise power at thereceiving earth station is given by

Nt =Nu g gd +Nd

where

Nu = noise power on the up-link

Nd = noise power on the down-link

g= total transponder gain

gd = section gain on the down-link (free space loss andantenna gains).

The total noise-to-carrier ratio is then

since

and C =Cdown

C

g

g

d

= Cup

N

C=

Nupggd + NdownC

=Nupggd

C+

Ndown

C

=Nup

C/g gd+

Ndown

C

F=k To + T( )B g

kToBg= 1+

T

To

F=Si

So

No

Ni=

1

g

No

kToB

FS N

S N

i i

o o

=/

/

then

(7.19)

Using equation (7.19) we can calculatethe total C/ N ratio, knowing C/ Nratios on the up- and down-link.

If the (C/N)up is more than 10 dBabove(C/N)down, the noise contribu-tion from the up-link to the down-linkis only a few tenths of a dB. Often it isdisregarded in link calculations wherethere is enough power on the up-link.

Knowing the

- location of satellite and earth station

- EIRP and G/T of the satellite

- EIRP of the earth station

- the required quality (C/N)

- the carrier frequencies

- the receiver overall system tempera-tureT (or noise figureF)

- the equivalent noise bandwidth ofthe receiver (B)

we are now able to calculate the recei-ving stations figure-of-merit (G/T) orreceiving stations antenna diameter.

C

N= 1

1C/N( )up+ 1

C/N( )down

In figure 35 is shown the EIRP con-

tours for INTELSAT-V A satellite in 1west.

Using the above developed equations,the corresponding receiving antennadiameters have been calculated, figure35. For the calculations the followingparameters have been assumed:

C/N =12 dB (TV reception withgood quality)

B = 27 MHz (receiver noise band-width)

F = 1 dB (receiver noise figure)

19

WS IS

receiving andtransmittingpatterns

wanted

signal

receivingpattern

down-linkinterference

up-linkinterference

transmittingpattern

WE IE

Figure 36 Simplified interference geometryWs, Is= wanted and interfering satelliteWE, IE = wanted and interfering earth station = angular separation between satellites

Figure 35 EIRP contours for INTELSAT-V-A at 1westand the corresponding receiving TV antennadiamters. The different parameters are given inthe text

EIRP:

antennadiameter

70 cm 90 cm 210 cm

48 dBW

39 dBW

46 dBW


17/18

TA = 30 K (antenna temperature,clear sky condition)

L = 0.5 dB (feeder loss)

f= 11.7 GHz (carrier frequency).

Corresponding calculations have beendone for the diameters given in figure31. The noise from the up-link hasbeen disregarded.

In the calculations above we have onlyconsidered additive thermal noise.

There are also other sources of degra-dation such as interference from othersatellites and radio links, intermodula-tion from the unlinear transponder,etc. The effect of these sources areoften also treated as additive thermalnoise and theC/I terms should beadded in equation (7.19):

(7.20)

C

N =1

1C/N( )up+ 1

C/N( )down+ 1C/I + etc.

Keeping the interference from othersatellites below a certain value will putsevere limits on the minimum diame-ter of the receiving earth stationsantenna. Knowing the radiation pat-terns for the satellites and earth sta-

tions involved, one can calculate theinterference level. The geometry isshown in figure 36.

In the near future, the interferencefrom other satellites may be the criti-cal factor for the design of earth sta-tion antennas. This is already the casefor certain positions in the geostatio-nary arc over Europe.

8 Satellite networks

The simplest form of satellite net-works is the point-to-point or point-to-multipoint configuration shown infigure 37a) and b). So far, we haveonly considered these cases. However,there is often a need for several earth

stations to be interconnected through

the same transponder, multipoint-to-multipoint, shown in figure 37c). Themethods for allowing several users toutilise the same transponder are calledmultiple access techniques. The trans-ponder is a resource which can becharacterised by its available powerand bandwidth. Efficient use of thiscommon resource is a very importantproblem in satellite communication.

The channels are designed to theusers either fixed (pre-assigned) or ondemand (demand assignments).

There are basically three methods ofmultiple access:

- Frequency Division Multiple Access(FDMA)

- Time Division Multiple Access(TDMA)

- Code Division Multiple Access(CDMA).

Random Access (RA) may also be uti-lised.

In FDMA, each user is permanentlyallocated a certain frequency band, outof the total bandwidth of the transpon-

der, figure 38a). To reduce the adja-cent channel interference, it is neces-sary to have guard bands between thesub-bands. Frequency drifts of thesatellites and earth stations frequencyconverters have also to be taken intoconsideration. FDMA is the traditionaltechnique due to its simple implemen-tation. However, due to the non-linearcharacteristics of the transponder,figure 29, a certain back-off of the

TWT is necessary for multicarrier ope-ration, in order to control the intermo-dulation products. This reduces thetotal transponder capacity.

In TDMA all stations use the samecarrier frequency, but they are onlyallowed to transmit in short non-over-lapping time slots, figure 38b). Thus,the intermodulation products due tothe non-linear transponder are avoi-ded. This means that the high poweramplifier in the transponder can beoperated near saturation. Accordingly,both the total transponder power andbandwidth are available. However, the

TDMA technique obviously requirestime synchronisation and buffer stor-age. This leads to rather complex

earth stations.In CDMA all users occupy the totaltransponder bandwidth all the time.

20

point-to-point point-to-multipoint(broadcast mode)

multipoint-to-multipoint

Figure 37 a), b) and c) Satelli te network configurations

(a)(b)

(c)


18/18

The users can be separated because

each channel is multiplied by a uniquespreading code. The composite signalis then modulated onto a carrier. Theinformation is recovered by multiply-ing the demodulated signal with thesame spreading code. The receivingearth station is accordingly able torecover the transmitted message by aspecific user. No frequency or time co-ordination is needed before accessingthe transponder. New users can easilybe included. CDMA signals are resis-tant to interfering signals. This pro-perty can be utilised in systems withvery small aperture terminals (VSAT),

where interference may be receivedfrom adjacent satellites. The main dis-advantages are cost and complexity ofthe receivers.

9 Literature

1 Berlin, P.The geostationary appli-cations satelli te, Cambridge Uni-versity Press, 1988.

2 Maral, G, Bousquet, M. Satellitecommunications systems, JohnWiley & Sons, 1986.

3 Ha, T T. Digital satellite communi-cations, Macmillan Publishing Co.,1986.

4 Evans, B G. Satellite communica-tion systems, Peter Peregrinus Ltd.,2nd edition, 1991.

5 Pratt, T, Bostian, C W.Satellitecommunications, John Wiley &Sons, 1986.

6 Dalgleish, D I. An introduction tosatellite communications, PeterPeregrinus Ltd., 1989.

7 Feher, K. Digital communications.Satellite/earth station engineering,Prentice-Hall Inc., 1983.

8 ITU/CCIR. Handbook of satellitecommunications, ITU, Geneva,1988.

9 Stette, G. Forelesninger i satellitt-kommunikasjon, NTH.

10 Gutteberg, O, Lundstrm, L-I,Andersson, L. Satellittfjernsyn, Uni-versitetsforlaget, 1986.

11 Gagliardi, R G.Satelli te communi-cations, Van Nostrand Reinhold,2nd edition, 1991.

21

transponder bandwidth

frame length

transponder bandwidth

guard band

frequency

time

frequency

FDMA

TDMA

CDMA

guard time

synchonizing

burst

burst from

different earth stations

coded carrier spectra

Figure 38 a), b) and c) Basic multiple access methods

Elements of satellite technology and communication

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