satellite mod 1

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SATELLITE

COMMUNICATION SYSTEMS

Girish K.P.

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Communication satellites bring the world

to you anywhere and any time…..

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What exactly is a satellite?

• The word satellite originated from the Latin word “Satellit”- meaning

an attendant, one who is constantly hovering around & attending to a

“master” or big man.

• For our own purposes however a satellite is simply any body that

moves aroun ano er usua y muc arger one n a ma ema ca ypredictable path called an orbit.

• A communication satellite is a microwave repeater staion in space that

is used for tele communcation , radio and television signals.

• The first man made satellite with radio transmitter was in 1957.

. There are about 750 satellite in the space, most of them are used forcommunication.

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How do satellite work?

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How do Satellites Work?

* Two Stations on Earth want to communicate through radio

broadcast but are too far away to use conventional means.

The two stations can use a satellite as a relay station for their

communication.

* One Earth Station transmits the signals to the satellite. Up link

communicating with Satellite.

* The satellite Transponder converts the signal and sends it down

to the second earth station. This frequency is called a Downlink.

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Consider the light bulb example:

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Components of a satellite

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Advantages of satellite over terrestrial communication :

* The coverage area of a satellite greatly exceeds that of a

terrestrial system.

* Transmission cost of a satellite is independent of the distance

from the center of the coverage area.* Satellite to Satellite communication is very precise.

* Higher Bandwidths are available for use.

Disadvantages of satellites:

* Launching satellites into orbit is costly.

* Satellite bandwidth is gradually becoming used up.* There is a larger propagation delay in satellite communication

than in terrestrial communication.



How does a satellite stay in it’s orbit?

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How do we escape gravity & place an

object in orbit?

• If an object is fired

fast enough it should

escape e ear s pu .

• This is done through

the use of RocketLaunchers

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Multi-stage Rockets

• Stage 1: Raises the

payload e.g. a satellite toan elevation of about 50miles.

• Sta e 2: Satellite 100

miles and the third stageplaces it into the transferorbit.

• Stage 3: The satellite is

placed in its final geo-synchronous orbital slotby the AKM, a type ofrocket used to move the

satellite.

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Applications

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Major problems for satellites

• Positioning in orbit

• Stability

• Power

• Communications

• Harsh environment

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Positioning

• This can be achieved by several methods

• One method is to use small rocket motors

• These use fuel - over half of the weight of most satellites is made up of fuel

determines the lifetime of a satellite

• Commercial life of a satellite typically 10-

15 years

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Stability

• It is vital that satellites are stabilised

- to ensure that solar panels are aligned properly,

communication antennae are aligned properly

• Early satellites used spin stabilisation

- either this requires an inefficient omni-directional

aerial Or antennae were precisely counter-rotated inorder to provide stable communications.

* Modern satellites use reaction wheel

stabilisation - a form of gyroscopic stabilisation.



Power

• Modern satellites use a variety of power means• Solar panels are now quite efficient, so solar

• Batteries are needed as sometimes the satellites

are behind the earth - this happens about half the

time for a LEO satellite• Nuclear power has been used - but not

recommended

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Satellite - satellite communication

• It is also possible forsatellites tocommunicate withother satellites

• Communication canbe by microwave orb o tical laser

1.

2.

1.

2.

1.

2.

Point-Point System Crosslink System Hybrid System

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Harsh Environment

• Satellite components need to be specially

“hardened”

• Circuits which work on the ground will fail veryrapidly in space

• empera ure s a so a pro em - so sa e es useelectric heaters to keep circuits and other vital

parts warmed up - they also need to control the

temperature carefully

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Early satellites• Telstar

– Allowed live transmission across the Atlantic

• Syncom 2– First Geosynchronous satellite

TELSTAR SYNCOM 2

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Satellite orbits

Classification of orbits:

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* Circular orbits are simplest

* Inclined orbits are useful for coverage of

equatorial regions* Elliptical orbits can be used to give quasi

4 satellites

* Orbit changes can be used to extend the life of

satellites

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Classification of orbits:

Satellite orbits are also classified based on theirheights above the earth:

– GEO– LEO

– MEO

– Molniya Orbit– HAPs

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Satellite orbit altitudes

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Geostationary Earth Orbit (GEO)

• These satellites are in orbit 35,786 km above the earth’ssurface along the equator.

• Objects in Geostationary orbit revolve around the earthat the same speed as the earth rotates. This means GEO

satellites remain in the same position relative to thesurface of earth.

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GEO contd.

• Advantages– A GEO satellite’s distance from earth gives it a large

coverage area, almost a fourth of the earth’s surface.

– GEO satellites have a 24 hour view of a particular area.

– These factors make it ideal for satellite broadcast and other

multipoint applications

– Minimal do ler shift

• Disadvantages

– A GEO satellite’s distance also cause it to have both a

comparatively weak signal and a time delay in the signal,

which is bad for point to point communication.– GEO satellites, centered above the equator, have difficulty for

broadcasting signals to near polar regions

– Launching of satellites to orbit are complex and expensive.

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Low Earth Orbit (LEO)• LEO satellites are much closer to the earth than GEO satellites,

ranging from 500 to 1,500 km above the surface.

• LEO satellites don’t stay in fixed position relative to the surface,and are only visible for 15 to 20 minutes each pass.

• A network of LEO satellites is necessary for LEO satellites to beuseful



The Iridium system has 66 satellites in six LEO orbits,

each at an altitude of 750 km.

Iridium is designed to provide direct worldwide voice and data

communication using handheld terminals, a service similar to

cellular telephony but on a global scale

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LEO Contd.

• Advantages

A LEO satellite’s proximity to earth compared to a GEO

satellite gives it a better signal strength and less of a time

delay, which makes it better for point to pointcommunication.

A LEO satellite’s smaller area of coverage is less of a waste

of bandwidth.

• Disadvantages

A network of LEO satellites is needed, which can be costly

LEO satellites have to compensate for Doppler shifts cause bytheir relative movement.

Atmospheric drag effects LEO satellites, causing gradual

orbital deterioration.

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Medium Earth Orbit (MEO)

• A MEO satellite is in orbit somewhere between 8,000km and 18,000 km above the earth’s surface.

• MEO satellites are similar to LEO satellites in

functionality.• MEO satellites are visible for much longer periods of

time than LEO satellites, usually between 2 to 8 hours.

• MEO satellites have a larger coverage area than LEOsatellites.

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MEO contd.

• Advantage

A MEO satellite’s longer duration of visibility and

wider footprint means fewer satellites are needed in aMEO network than a LEO network.

• Disadvantage

A MEO satellite’s distance gives it a longer time

delay and weaker signal than a LEO satellite, though

not as bad as a GEO satellite.

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MEO satellites

Glonass (Russian)

The GPS constellation

calls for 24 satellites to be

distributed equally amongsix circular orbital planes

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Molniya Orbit

Used by Russia for decades.

Molniya Orbit is an elliptical orbit. The satellite remains in a

nearly fixed position relative to earth for eight hours.

A series of three Molniya satellites can act like a GEO satellite. Useful in near polar regions.

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High Altitude Platform (HAP)

One of the newest ideas in satellite communication. A blimp or plane around 20 km above the earth’s surface is

used as a satellite.

HAPs would have very small coverage area, but would have acomparatively strong signal.

Cheaper to put in position, but would require a lot of them in anetwork.

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HAP

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Satellite frequency band

BandDownlink,

GHzUplink, GHz

Bandwidth,

MHz

L 1.5 1.6 15

C 4 6 500

Ku 11 14 500

Ka 20 30 3500



Solar day and Sidereal day

• A day is defined as the time that it takes theEarth to rotate on its axis.

• However, there is more than one way to define

a day:– A sidereal day is the time that it takes for the Earth

to rotate with respect to the distant stars.

– A solar day is the time that it takes to rotate withrespect to the Sun.

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The Length of the Day

• A solar day is slightly longer than a sidereal day.

– A sidereal day is 23h 56m 4.091s.

• We set our watches according to the solar day.

• s ronomers use s erea me ecause we are

mostly interested in distant celestial objects.

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• A solar day is measured using the passage of theSun across the sky—it lasts 24 hours

• A sidereal day (from the Latin word meaningstar) is measured with respect to fixed stars—itlasts a little less than 24 hours.

• Eac so ar ay t e Eart rotates egrees w trespect to the Sun

• Each sidereal day the Earth rotates 360 degrees

with respect to the background stars• During each solar day the motion of the Earth

around the Sun means the Earth rotates 361

degrees with respect to the background starswww.edutalks.org



• The actual length of a sidereal day on Earth is 23 hours 56minutes 4 seconds

• This means that the Earth has to rotate slightly more than oneturn with respect to a fixed star to reach the same Earth-Sunorientation (solar day)

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The difference between solar days and sidereal days

means that a given star will rise earlier each day

These 3 photos show how Orion reaches the sameposition in the sky 4 minutes earlier on each

.

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Apparent Solar Time

• Apparent solar time is the time measured withrespect to the actual position of the Sun.

– At noon, the Sun would be exactly on the meridian.

– . .

the meridian.– 9 A.M. would be exactly 3 hours before the Sun was on

the meridian.

– The apparent solar time depends on your longitude.

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O i i f l t l



Origin of planetary laws

Sir. Johannes Keppler

Derived 3 laws based

upon his observations

of planetary motion.

Sir.Tycho Brahe

• Introduced precision into

astronomical measurements.• Mentor to Johannes Keppler

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Kepler’s 1st Law: Law of Ellipses

The orbits of the planets are ellipses with

the sun at one focuswww.edutalks.org

’ f A



Kepler’s 2nd Law: Law of Equal Areas

The line joining the planet to the center of the sun

sweeps out equal areas in equal times

T5T4 T3

T6

T1A2A3A4A5

A6

A1

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Kepler’s 3rd Law: Law of Harmonics

The squares of the periods of

two planets’ orbits areproportional to each other as

the cubes of their semi-

major axes:

T12 /T2

2 = a13 /a2

3

In English:Orbits with the same semi-

major axis will have the

same periodwww.edutalks.org

Newton’s Laws



Newton s Laws

• Kepler’s laws only describe the planetary motionwithout attempting to suggest any explanation as

to why the motion takes place in that manner.

Sir .Issac Newton

.

• Derived the Law of Universal

Gravitation.

• Explained why Kepler’s laws

worked.

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Newton’s 1st Law: Law of Inertia

• Every body continues in a state of uniformmotion unless it is compelled to change that

state b a force im osed u on it

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Newton’s 2nd Law: Law of Momentum

• Change in momentum is proportional to and

in the direction of the force applied

• Momentum equals mass x velocity

• =

F

F

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Newton’s 3rd Law: Action - Reaction

• For every action, there is an equal andopposite reaction

•

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Newton’s Law of Universal Gravitation

Between any two objects there exists a force of

attraction that is proportional to the productof their masses and inversely proportional to

e square o e s ance e ween em

Fg = G( )M1m2

r2

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Classical orbital elements

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Apogee and Perigee



p g g

• In astronomy, an apsis is the point of greatest or least distanceof the elliptical orbit of an astronomical object from its center of attraction, which is generally the center of mass of the system.

• The point of closest approach is called the periapsis (Perigee)or pericentre and the point of farthest excursion is called theapoapsis (apogee)

• A straight line drawn through the perigee and apogee is the line

of apsides. This is the major axis of the ellipse.

Ascending & Descending nodes

• These are the 2 points at which the orbit of a satellite penetrates

the equatorial plane.

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Cl i l bi l l



Classical orbital elements

• Six independent quantities are sufficient to

describe the size, shape and orientation of an

orbit.These are

– a the semi-ma or axis

– εεεε, the eccentricity – i, the inclination

– ΩΩΩΩ, the right ascension of the ascending node

– ωωωω, the argument of perigee – tp, mean anamoly

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• The semi-major axis describes the size of the orbit. It



The semi major axis describes the size of the orbit. Itconnects the geometric center of the orbital ellipse withthe periapsis, passing through the focal point where thecenter of mass resides.

• The eccentricity shows the ellipticity of the orbit.

• The inclination is the angle between the plane of theorbit and the equatorial plane measured at the ascendingnode in the north ard direction

• The right ascension of an ascending node is the anglebetween the x axis and the ascending node.

• The argument of periapsis (perihelion) is the angle in the

orbital plane between the line of nodes and the perigee of the orbit.

• The mean anomaly is the time elapsed since the satellitepassed the perigee.

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Major parameters of an elliptical orbit



Major parameters of an elliptical orbit

• Satellite trajectory

• Satellite period

• Satellite velocity

• Satellite position

Satellite Trajectory



j y

The path of a satellite in space may be obtainedunder the following assumptions:

1.The satellite and earth are symmetric spherically

and may be treated as point masses.

2.There are no other forces acting on the system

besides the gravitational forces.3.The mass of the earth is much greater than

satellite.

These assumptions lead to the two body problem.

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Applying Newton's laws to such systems,



..∑F = m r (second law) …………(1)F = -GMm. r (third law) ……………..(2)

r2 r

Substituting (1) in (2) we get,.. ..

r + GM .r = 0 (or) r + µ .r = 0r3 r3

..

Where r = vector acceleration in the given coordinate system

r = vector rom mass o eart to m mass o sate te

r = distance between M and m , µ = GM (gravitational parameter)A partial system is easy to obtain and is adequate for illustrating the size and shape of

an orbit.

The resulting trajectory equation has a general form of conic section:

r = P ; p = a geometric constant called parameter of conic

1+e cos θ = (r v cos ф)2 / µe = the eccentricity which determines type of conic section

=√(1-P/a)

θ = angle between r and the point on the conic nearest the focus

ф = flight elevation angle , v = satellite velocity

a = semi-major axis = (ra+rb)/2

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Satellite period



Satellite period

The period T of a satellite is given as:

T2= 4 П2 .a3 (period depends only on semi major axis,a)

µFor a satellite in circular orbit around earth-

T2= 4 П2 .(R+h)3

µWhere , R= radius of earth,

h= satellite altitude

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Satellite velocity



Total specific mechanical energy ε of a satellite is the sumof kinetic energy/unit mass and potential energy/unit

mass, but there is an interchange between these energies.

Thus a satellite slows down when it moves up and gainsspeed as it loses height.

V2= µ(2/r -1/a)

also ε = V2 / 2 - µ /r and ε = –µ /2a

For circular orbit the equation reduces to:

V2 = µ /r

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Satellite position with time



The origin O is the geocentre.

The satellite at any instant tp is assumed to be at S.

The circle is drawn from centre C of the ellipse with a radius equal to

the semi major axis and a perpendicular BM is drawn passing throughthe point S.

Angle E is called eccentric anomaly and angle θ is the true anomaly.

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Satellite position



For an elliptic orbit, the time tp elapsed from a perigee pass is defined as-

tp = T/2Π (E-e sin E)

= (T/2Π)M ; where M = E-e sin E

Eccentric anomaly is defined as

= arccos e cos e cos

where θ = true anomaly

= 2tan-1 [( 1+e)/(1-e)]1/2 tanE/2

When θ=0 ,the mean and true anomalies are equal.

Hence distance between satellite and geocentre isr = a(1-e2)/(1-ecosθ)

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GEOSYNCHRONOUS

AND

GEOSTATIONARY ORBITS

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GEOSYNCHRONOUS ORBITS



• A geosynchronous orbit is the one with an orbital

period (the time needed to orbit once around the

Earth) that matches the rotation rate of the Earth.This is a sidereal day, which is 23 hours 56

.

• A geosynchronous earth orbit is sometimesreferred to as the Clarke orbit or Clarke belt , afterArthur C. Clarke, who first suggested its‘ existencein 1945 and proposed its use for communicationssatellites

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Clarke Orbit



• The Clarke orbit meets the

concise set of specifications

for geosynchronous satellite

orbits:– (1) be located directly above

– (2) travel in the samedirection as Earth's rotation at

6840 mph

– (3) have an altitude of 22,300

miles above Earth

– (4) complete one revolution in

24 hours

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Clarke Orbit



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Geo synchronous Satellites



• There is only one geosynchronous earth orbit.

• It is occupied by a large number of satellites. In fact, the

geosynchronous orbit is the most widely used earth orbit for theobvious reason.

• An international agreement initially mandated that all satellitesp ace n e ar e or mus e separa e y a eas m es.

• This stipulation equates to an angular separation of 4° or more,which limits the number of satellite vehicles in a geosynchronousearth orbit to less than 100.

• Today, however, international agreements allow satellites to beplaced much closer together.

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Geo stationary orbit



y

• A geostationary orbit is a special case of ageosynchronous orbit.

• A satellite is in a geostationary orbit when it appears

stationary from the point of view of an observer on theEarth's surface.

This can onl occur when:

• The orbit is geosynchronous• The orbit is a circle

• The orbit lies in the plane of the Earth's equator

• Thus, a geosynchronous satellite will be geostationaryonly with the additional restrictions of it being in acircular orbit situated over the equator.

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Geostationary Vs. Polar



y

Orbiting

http://cimss.ssec.wisc.edu/satmet/modules/sat_basics/images/orbits.jpg

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Geostationary Satellites



The satellite velocity in this

orbit is 3075 m/s.

• Operate in the 2.0 GHz to18 GHz range

eccentricity of the orbit iszero, the satellite appears to

be stationary to an observer

from ground.

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Geostationary Satellites in Orbit



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Geostationary Satellite Coverage

http://www.ssec.wisc.edu/mcidas

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Geo-stationary satellites



Applications:

Telecommunication systems

Radio

Data Transmission systems

• The geometric considerations like satellite elevation/look

angle etc are very vital for reliable communication

satellite system design.

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Satellite elevation:



The elevation of a satellite,η is the angle which a satellite makes with

the tangent at the specified point on the earth.

η = arc tan [(cosψ-σ)/ sin ψ]

Where, coverage angle ψ = arc cos (cosθc cosφcs )φcs = φc - φs and σ =R /(R+h) = 0.151

In terms of elevation angle:

ψ = 900 – η-sin-1(cos η / 6.63235)

In terms of tilt angle : ψ = sin -1(6.6235 sinγ- γ) where θc = latitude

of earth station, φc = the longitude, φs = longitude of sub satellite

point, R=radius of earth, h=satellite height above equator

Tilt angle γ = arc tan [sin ψ / (6.6235-cos ψ)www.edutalks.org

Azimuth:The azimuth ξ is the angle which the satellite direction makes

with the direction of true north measured in the clockwise direction



with the direction of true north measured in the clockwise direction.

The azimuth ξ = arc tan [tan φcs /sinθc]

in northern hemisphere:

ξ =1800 + A0;when the satellite is to the west of earth station

ξ =1800 - A0;when the satellite is to the east of earth station

in southern hemisphere:0 -

ξ =A0

;when the satellite is to the east of earth station

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Range:The range d of a geostationary satellite is given by,

1/2



d = 35786[1+0.41991-cos ψ1/2

In terms of radius of earth (ie, der = d/r)

der= [13.47(1-cosβ+31.624)1/2 also der = 6 .6235 sinψ /cos η

• The angle β, is the angle between the solar vector and the orbit

plane. If the solar vector is in the orbit plane, β = 0°. Beta can go

to ± 90°. The general convention is that β is positive when the sun

is on the same side of the orbit plane as the positive orbit normal(right hand rule).

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Launching of geostationary satellite:• Initially place spacecraft with the final rocket



stage into LEO.• After a couple of orbits, during which the orbital

parameters are measured, the final stage is

reignited and the spacecraft is launched into a

geostationary transfer orbit(GTO).

• Perigee of GTO is that of LEO altitude and

.

• After a few orbits in GTO, while the orbitalparameters are measured, a rocket motor (AKM)

is ignited at apogee and GTO is raised until it is

circular geostationary orbit.

• AKM (Apogee Kick Motor) is used to circularizethe orbit at GEO and to remove any inclination

error so that the final orbit is very close to

geostationary.

Launching of geostationary satellite



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Geostationary Transfer

Orbit



• If we speed the satellite up while it's in low

circular earth orbit it will

go into elliptical orbit,

heading up to apogee.

• If we do nothin else, it

• BUT, if we fire a rocket motor when the satellite's at apogee,

and speed it up to the required circular orbit speed, it will stay at

that altitude in circular orbit. Firing a rocket motor at apogee is

called "apogee kick", and the motor is called the "apogee kick

motor".

will stay in this elliptical

orbit, going from apogee

to perigee and back again.

Few Geostationary satellites: EDUSAT, INTELSAT ,



Phase I and II of launching spacecraft

INSAT , PAKSAT, AMERICOM …….

Phase I ↑

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ORBITAL MANEUVERS

Hohmann Transfer



– Can be used to raise or loweraltitude

– Most efficient method

– At minimum, requirescompletion of half revolutionof transfer orbit

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Hohmann transfer• Most satellites launched today are initially placed into an



low earth orbit.

In the next phase the satellite is injected into an elliptical

transfer orbit which has an apogee at the height of GEO and

its apsides (line joining perigee-apogee) in the equatorialplane.

velocity increment at the apogee equal to the differencebetween satellite velocity at GTO and velocity in GEO.

• A transfer between two coplanar circular orbits via

elliptical transfer orbit requires the least velocity increment(and hence fuel). This principle was recognized by Hohmann

in 1925 and is referred as Hohmann transfer.

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• A Hohmann transfer is a fuel efficient way to transferfrom one circular orbit to another circular orbit that is in



the same plane (same inclination), but a different altitude.• To change from a lower orbit (A) to a higher orbit (C),an engine is first fired in the opposite direction from thedirection the vehicle is traveling.

• This will add velocity to the vehicle causing its trajectoryto become an elliptic orbit (B). This elliptic orbit iscarefully designed to reach the desired final altitude of the higher orbit (C).

• In this way the elliptic orbit or transfer orbit is tangent toboth the original orbit (A) and the final orbit (C). This iswhy a Hohmann transfer is fuel efficient.

• When the target altitude is reached the engine is fired inthe same manner as before but this time the addedvelocity is planned such that the elliptic transfer orbit iscircularized at the new altitude of orbit (C).

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Hohmann Transfer

Target Orbit



Initial

Orbit

Transfer Orbit

The orbital inclination is given by,

cos i= sinξ1 cos θ1

where i=inclination

ξ1 =azimuth of launch

θ1 =latitude of launching site



S W h EffS W h Eff



Space Weather EffectsSpace Weather Effects

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Effect of earth on satellites

Th ff t f it ti l f i



• The effect of gravitational force isnon uniform because of the nonuniform distribution of earth’s mass -a slight bulge at the equator, with a

difference of 21 km between polarand the equator radius.

• This deviation from spherical shape

satellite.

• The effect of earth’s gravitationalpull may be expressed as theharmonic series of the field. Thefirst term represents the principal

gravitational law and the higherorder terms in the series as theperturbations.

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The main effects of perturbations are:

1. The component of perturbations in the orbital plane causes the

perigee to rotate in the orbital plane.



2. Another effect of perturbations is that the orbital plane rotates around

the earth’s north-south axis.

3. The perturbating force along the orbital plane imparts a force vector

on a satellite

1. The component of perturbations in the orbital plane

causes the perigee to rotate in the orbital plane.

The rate of change of argument of perigee is

ω = 4.97[R/a]3.5 (5cos2i-1)/(1-e2)2 deg/day

where R= mean equatorial radius , a=semi major axis

i = inclination, e=eccentricity

• when i=63.40 , ω reduces to zero, implying that perigee remains

fixed in space.

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2. The orbital plane rotates around the earth’s north-south axis.

The rate of change of rotation of ascending node is

= 9.95[r/a]3.5 cos i /(1-e2)2 deg/day



Where r = satellite-geo centre distance

The rotation is in a direction opposite to the satellite motion. For a

geostationary orbit magnitude is 4.90 /year ,implying the ascending

node rotates around the earth in 73 years.

. e per ur a ng orce a ong e or a p ane mpar s a

force vector on a satellite.For most orbits such components cancel out as the satellite position

changes continuously.

In the geostationary orbits, resultant perturbating component do not

cancel but cause a satellite to drift towards one of the two neareststable points on the orbit.

Stable points are approximately on the minor axis, showing that the

elliptical approx. of earth is not precisely accurate.

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Gravitational effects from heavenly bodies:

• In LEO satellites the influence of gravitational forces



• In LEO satellites, the influence of gravitational forcesfrom sun and moon are small when compared to thegravitational force of earth.

• The order of magnitude of gravitational force of moonand sun are main sources of perturbations in GEOsatellites.

• When nearer to heavenly bodies, the gravitational pull is

stronger and hence causes a gravity gradient. maineffect of such gradient is to change the inclination of theorbit.

• The combined effect of sun and moon is to cause achange in inclination of GEO satellites between 0.750

and 0.940

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• The inclination of orbital plane caused by moon changes

cyclically between 0.480 and 0.670 with a period of 18.6

years Maximum inclination change occurred in year



years. Maximum inclination change occurred in year1987 and minimum in Feb 1997

• The change in inclination due to sun is 0.270 /year.

Note: Among the three forces affecting the inclination (gravity pull,

sun and non spherical nature of earth) the later force has a component

in the direction opposite to the former two forces.

Hence these forces cancel out at an inclination angle of about 7.50

Thus the inclination of satellite when left uncorrected oscillates around

the stable inclination with the period of about 53 years reaching amaximum of 150 and a minimum of 00

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Atmospheric Drag

• Satellites below 2000 kilometres are actually travelling



• Satellites below 2000 kilometres, are actually travellingthrough the Earth’s atmosphere. Collisions with air

particles, even at these high altitudes slowly act to

circularise the orbit and slow down the spacecraftcausing it to drop to lower altitudes , this effect is

• Emissions from the Sun cause the upper atmosphere toheat and expand.

• These energetic solar outputs increase dramatically

during periods of high solar activity, and may result inEarth-orbiting satellites experiencing an increase in

atmospheric drag

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A satellite orbiting the Earth would continue to orbit

forever if gravity were the only force acting on it



forever if gravity were the only force acting on it.

Perigee remains same, Apogee decreases

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• Reduces satellite’s energy

• Changes the size (semi-major axis) and shape

(eccentricity)



(eccentricity)• The effect of drag is more severe at about 180km and

causes excess heat on satellite .Unless such LEO satellitesare routinely boosted to higher orbits, they slowly fall,

and eventually burn up• Orbital life time of satellite at 400km circular earth orbit is

,decades if they are at 800km altitude

• In the former case, functional life time depends on orbitallife time and for latter the life time of satellite equipmentsis the deciding factor.

• However, for GEO satellites the governing factors areequipment life time and fuel capacity of the satellite(typically 10-15 years).

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Solar radiation pressure• Solar radiation pressure is the force exerted by solar

radiation on objects within its reach

• The effect of solar radiation pressure increases as the



• The effect of solar radiation pressure increases as thesurface area of the satellite projected in the

direction of sun increases.

• The net effect is the increase in the orbital eccentricityand also introduces disturbing torque that effects thenorth-south axis of the satellite.

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Effects of Solar radiation pressure

1. HUMAN HEALTH



1. HUMAN HEALTH• Intense solar flares produce very high energy particles

that can be as harmful to people as low-energyradiation from nuclear blasts. Earth’s atmosphere andmagnetosphere provide protection for people on theground, but astronauts in space are subject to

. high energy particles into living cells leads tochromosome damage and, potentially, cancer.

• Airline pilots and flight crews, as well as frequentfliers, also receive increased doses of radiation from

solar flares. If you were travelling in an aircraft at highaltitudes during a major solar flare, the amount of radiation you would be exposed to can be equivalent togetting a chest x-ray.

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3. NAVIGATION

• A Global Positioning System (GPS) receiver uses radio signals fromseveral orbiting satellites to determine the range, or distance, from

each satellite, and determines from these ranges the actual position of h i



g pthe receiver.

• The radio signals must pass through the ionosphere, the uppermostpart of the Earth’s atmosphere, and in doing so are subjected to

variations in the electron density structure of the ionosphere.

• anges n e e ec ron ens y ue o space wea er ac v y canchange the speed at which the radio waves travel introducing a“propagation delay” in the GPS signal. Changing propagation delayscause errors in the determination of the range.

• An increase in space weather activity may cause widespreaddisruption to aircraft and ship navigation and emergency locationsystems that rely heavily on satellite navigation data.

Electro-Magnetic effect

I t ti b t th E th’ ti



• Interaction between the Earth’s magneticfield and the satellite’s electro-magnetic

field results in magnetic drag

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Magnetic storm



A geomagnetic storm is a temporary disturbance of the earth’s

magnetosphere caused by a disturbance in space weather. A geomagnetic

storm is caused by a solar wind shock wave. This only happens if the

shock wave travels in a direction toward Earth.

The solar wind pressure on the magnetosphere will increase or decreasedepending on the Sun's activity. These solar wind pressure changes

modify the electric currents in the ionosphere.

Magnetic storms usually last 24 to 48 hours, but some may last for many

da s.www.edutalks.org

Non geostationary constellations

• The design of constellations can be categorized according

to inclination, altitude and eccentricity.



, yOn the basis of inclination, two types of constellations are

designed.

Type I constellations are those having their orbital planeswith a common intersection point.

E : Polar constellations

• Type II constellations have optimized inclined orbit

constellations and distribute satellites uniformly.• Eg: inclined constellations

• Depending on altitudes, constellations may be

LEO,MEO etc.• A hybrid of orbital altitudes are also possible within asystem (A LEO satellite can be used together with ageostationary orbit satellite)

LEO Satellite coverage



LEO Satellite coverage

Demo of satellite coverage



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Advantages of Non-geostationary constellations

1. Since these orbits are closer to earth, the free space loss islower and hence it is possible to use hand held terminals.



Path loss at 1.5 GHz for LEO=152.87 dB, MEO=175.96dB and GEO=187.10dB.

2. LEO and MEO reduces the propagation delay whichreduces or eliminates delay related problems.

Maximum distance between two points which view asatellite at an elevation angle 100 is

LEO (at altitude 700km) = 3885 km

MEO (10000km) = 12790 km

GEO (36000km) = 15914 kmHence note that 4 LEO satellites would cover the samegeographical distance as a single GEO. Thus LEO has 4times frequency reuse than GEO.

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4 . Distributed architecture of LEO and MEO orbits makethem more resistant to satellite failures and hence morereliable.

5. Competitions between operators has triggered a feverishtechnical regulatory and financial activity in the



technical, regulatory and financial activity in theindustry.

Current non geostationary proposals are-

• - .bit rates communication facility. eg: ICO system

• Little LEO -Offers a low bit rate non-real time servicessuch as messaging (bit rate< 4kbps). eg. ORBCOM

• Big LEO -Offering medium bit rate interactive services

such as voice (bit rate 1-10 kbps) eg: Iridium.• Broad band LEO -Offers broadband services such asinternet high speed file download

(bit rate=16kbps to 1 Gbps) eg: Teledesic.

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Polar constellations

• A polar orbit is an orbit in which a satellite passes above or

nearly above both poles of the body (usually a planet suchas the earth) being orbited on each revolution It therefore



as the earth) being orbited on each revolution. It thereforehas an inclination of (or very close to) 900 to the equator.

• Polar orbits are often used for earth-mapping, earthobservation, as well as some weather satellites

• The disadvantage to this orbit is that no one spot on theart s sur ace can e sense cont nuous y rom a sate te

in a polar orbit.• Polar satellites include: Defense meteorological satelliteprogram (DMSP), Landsat, SPOT and NOAA. Landsat andSPOT are Commercial polar orbiters and are intended for

geophysical remote sensing• To achieve a polar orbit requires more energy, thus more

propellant is needed than an orbit of low inclination

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Polar constellations



Here : Ψ = coverage circlem = number of orbital planes

n = satellites /plane

∆ = cos

-1

[cos Ψ /cos Π /m]www.edutalks.org

Single coverage:

• Satellites in adjacent planes move in same direction, shifted with

respect to each other by half intra-orbit satellite separation (Π /m),where m = number of planes.



The separation between adjacent planes is (Ψ+∆) and the relative

geometry remains constant because they move in phase.

• Satellites are separated by 2∆,when the satellites move in opposite

directions and the relative geometry is not constant.

The total number of satellites,N = 4 /(1-cos Ψ) ; 1.3n < m < 2.2n

In the cases of non integer, next highest integer satisfying the

inequality can be taken.If the N is much large, then the condition (n-1) Ψ + (n+1) ∆ = П

• When the coverage is required beyond a latitude λ ,the equations are –

(n-1) Ψ+(n+1) ∆ = П cos λ and N = 4cos λ /(1-cos Ψ)

• The coverage efficiency of the constellations is given by N/4П



• The coverage efficiency of the constellations is given by N /4П

Where N = total solid angle

= solid angle bounded by a single satellite=2П(1-cos Ψ)

Triple coverage:

• The constellation geometry is similar to single coverage case, with atleast three satellites must be visible at all points.

• The coverage angle is adjusted such that at least 3 satellites lie within

angle Ψ of each point of set.

• The resulting relationship for providing triple coverage from pole up

to latitude λ is

N = 11cos λ /(1-cos λ ) ; 1.4n < mcos λ < 2.4n

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Inclined orbit

• A satellite is said to occupy an inclined orbit around the earth if the orbit exhibits an angle other than zero degrees with the



equatorial plane

• They have an inclination between 0 degrees (equatorial orbit)

and 90 degrees

• This family of satellites provides unbiased worldwide coverage

by deploying satellites in circular orbits of same period and

inclination, distributed uniformly on the sphere.• The orbital altitude of these satellites is generally on the order of a

few hundred km, so the orbital period is on the order of a few

hours. These satellites are not sun-synchronous, however, so they

will view a place on Earth at varying times.

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Hybrid constellations

• This combines the various types orbits for full earthcoverage



• These orbits have different orbital period

• Eg: using circular orbits for covering equatorial regionsand elliptical orbits for higher altitude regions

Eg: using GEO for covering equatorial regions and

inclined orbits for polar regions

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Footprint and spot beams



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Constellations for store and forward system:

• Systems which do not require real time coverage (messaging/paging)is less stringent because gaps in coverage are allowed, provided at

least one satellite is visible within (ta-td) where ta=specified end toend delay , td=delay in message transfer



• Eg :A single satellite in polar orbit can cover every regions of earthwithin a time dependent on the orbital period

,

reliable, two-way data communications

services around the world through aglobal network of 29 low-earth orbit

(LEO) satellites and accompanying

ground infrastructure. The system can

send and receive short messages,

between six bytes and several kilobytes

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Random and phased constellations

• Constellations can be categorized on the basis of phase relationshipbetween satellites with respect of each other.

• In random constellation all or some constellation parameters such



• In random constellation all or some constellation parameters suchas altitude inclination, inter orbital plane separation and inter satellite

phase are chosen at random.• In phased constellation all these parameters are well defined.

Random constellations are sim ler to maintain but are inefficient interms of coverage property and tend to randomly crowd the celestial

sphere around the chosen altitude.

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Design considerations of a non GEO satellite systems

1.Traffic distribution and coverage:

• A constellation design depends on service area and geographicaldistribution of traffic within that area.



• A worldwide coverage is essential for an operator interested inglobal operations, but regional operator is interested in only aspecific region. Hence the constellations are completely different.

• Good RF visibility ensures adequate signal strength before aconnection is established and the increase in spectrum reuse.

• Complexity in coverage design is the dynamic variation in positionof the footprint of each satellite and the constellation as a whole,making the geometric relationship time dependent.

• Hence an estimate can be made on the basis of known growth trends

from existing systems, population density, per capita income,existing infrastructure, market segmentation due to competition andprevailing economic /political condition of the target market.

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6. Spectrum availability:

• Frequency reusability can be increased by spatial/polarizationdiversity

• This is achieved by using spot and shaped beams

• For a given spot beam size lower altitude constellations can give



• For a given spot beam size lower altitude constellations can giveincreased reusability

• Additional measures like modulation, coding and multiple accessschemes can maximize radio resource

.

• Space environment affects the orbit selection

• Atmospheric drag, eclipses, ionization

8. Launch considerations

• Important practical consideration is the launch cost, feasibility of

launching the satellites in the acceptable time frame• Probability of launch failure and in-orbit satellite failure increases as

the number of satellites in the constellation increases.

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Problems

Q1. Find out the radius of a geostationarysatellite orbit



satellite orbit.

Given:

T = 23Hr 56Min 4.1Sec

G = 6.672 X10-11 m3 /kg/s2

M = 5974 X1024 kg

r = 6378.1414 km

take µ = √GM

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Answer:

T2 = 4 П2 .(R)3

µR = [T√GM]2/3



[ ]-------------

2П

= [23x60x60+56x60+4.1√ 6.672 X10-11x 5974 X1024 ]2П

= 42164.17 km

Altitude, h = R-r= 42164.17-6378.1414

= 35786.02 km

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Q3. Calculate the apogee and perigee heightsfor the given orbital parameters.



g p

e=0.0011501 and

a= 7192.3 km

Given:

r = 6378.1414 km

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Answer:

ra=a(1+e) = 7200.57 km



rp= a(1-e) = 7184.03 km

, a a – .

Perigee height, hp = rp – r = 805.89 km

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Q4. A satellite is in elliptical orbit with aperigee of 1000 km and an apogee of 4000

k U i th di f 6378 14 k



km. Using a mean earth radius of 6378.14 km

, find the period of the orbit in hours ,minutesand seconds.

Also find the ‘e’ of the orbit.

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Communication Satellites



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Communication satellites (comsat)

• Satellite is a RF repeater in orbit.

• The design of a satellite is governed by the communication capacity,physical environment in which it is operated and state of technology.



Main considerations of a comm. satellite are:-

i) Type of service to be provided (eg: mobile communication, DTH)ii) communication capacity (transponder BW and satellite EIRP)

iii) coverage area

iv) Technological limitations• Basic specifications are laid out for satellite depending on the

communication requirement.

• A domestic fixed satellite service it is the EIRP per carrier, number

of carriers and coverage area.• A direct broadcast satellite it is the number of television channels and

coverage area.

First TV image of weather (1960)



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First complete view of world’s weather, photographed by

TIROS 9 (13/2/1965).

Image assembled from 450 individual photographs



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Environmental Conditions:• A spacecraft must be reliable in all types of environments beginning from

launch to the in-orbit deployment and throughout its operation phase.

Most important stresses are-

a) Zero gravity :-

• At GEO, gravitational force is negligible giving rise to zero gravity effects.

• Major effect is on liquid fuel flow and hence external means are to be



• Major effect is on liquid fuel flow and hence external means are to beprovided for liquid flow.

• The absence of gravity facilitates operation of the deployment mechanismsused for stowing antennas and solar panels during launch.

mosp er c pressure an empera ure :-• At high altitudes, atmospheric pressure is extremely low (10-7 torr).

• This makes thermal conduction negligible and increase friction betweensurfaces.

• Hence special materials are used for lubrication of moving parts.

• However, pressure inside the spacecraft is higher because of out gassing of

electronic components.• The temperature of a spacecraft is mainly affected by heat from sun and

various spacecraft subsystems.

• The excursion in the external temperature varies from 330-350K duringsunlight and 95-120K during eclipses.

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c) Space Particles :-

• Various types of particles like cosmic rays, protons, electrons, meteoroids,manmade debris etc exist in space.

• Main effect of bombardment of particles on a satellite is the degradation of

solar cells and certain solid state components within the satellite.• Effect of meteoroids is negligible in GEO satellites.

d) Magnetic fields :-



• Magnitude of earth’s magnetic field is very weak at GEO (1/300 of earths

surface).• The effect of magnetic field can be compensated by the use of large coil.

• ,that are trapped in this region affect electronic components

• Hence special manufacturing mechanisms are used to harden thecomponents against radiations.

e) Other Considerations :-

• Due to the variation of distance of earth from sun, a variation in DC generationcapability must be taken into account in design of satellite power system.

• Also satellites must be prepared for loss of power during eclipses and may result ingradual degradation of solar cell efficiency.

• There are several perturbations affecting the satellites due to movement of mechanical parts and fuel within it.

• There may be a small drift in position of antennas.

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Life time and reliability

• Lifetime of a geostationary satellite is determined by the maximumacceptable deviation in inclination and orbital location.

• Satellite is maintained in its orbital location by firing thrustersregularly, using stored fuel

• Hence the operational lifetime of a satellite is determined by-



Hence the operational lifetime of a satellite is determined by

a) Increasing fuel capacityb) Saving fuel by accepting orbital deviation to the maximum extent

.

• However there is a practical limit to a satellites fuel storage

capacity. Hence satellite lifetime is between 12-15 years.

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Reliability

• The overall reliability of a satellite is governed by its critical

components.

• Reliability is improved by employing redundancy in the critical sub

systems and in components such as TWT amplifiers.



• Reliability is defined as the probability that a given component/system

performs its function within a specified time t.

• R= where λ = failure rate of a componentt

dt− λ

• Unit of λ is specified as FIT, the number of failures in 109 Hr.

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• ree reg ons can e ent e

An early high failure rate region attributed to manufacturing faults, defects in

materials etc

A region of low failure attributed to random component failures

A region of high failure rate attributed to component wear-out.

In a satellite system, early failures are eliminated to a large extent during testing and

burn-in. The main aim is to minimize the random failures which occur during theoperational phase of the satellite by using reliability engineering techniques. The

beginning of wear-out failure can best be delayed by improving the manufacturing

technique and the type of material used.

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• The reliability can be expressed as

R = e-λ t = e-t/m ; where m =1/ λ (mean time between failures)

• When several components or sub-systems are connected in series, the

overall reliability is

Rs=R1 R2…….Rn where Ri is the reliability of the ith component.

• In terms of the failure rate :



• In terms of the failure rate :

Rs= e-(λ 1+ λ 2+… λ n)t

• Parallel redundancy is useful when the reliability of an individual

subsystem is high.

• If Qi is the unreliability of the ith parallel element, the probability that

all units will fail is the product of the individual unreliabilities

Qs=Q1 Q2…Qi

• When the unreliabilities of all elements are equal, this expression

reduces to Qs = Qi ;Where Q is the unreliability of each element.Therefore the reliability is

R = 1-Qs

= 1-i

=1- 1- Ri

=1- 1-e-λ t

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A Typical reliability model of a Geostationary Satellite:

• All the major sub-systems are shown in series.



Simplified reliability model • Applying the equation for series and parallel combination, the

Rs =RRXRTX [1-(1-RT)2]

• When RT =0.9,reliability of transponder increases to 0.99• Figure of merit, Fγ = r/M ;where r = R’/R

R’= reliability with redundancy employed

R= reliability without employing redundancy

M= increase in mass due to added redundancy

• The addition of redundant equipment increases the cost of the

transponder

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Back up slides

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Transponder



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Solar eclipses



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----Introduction to Solar System Dynamics----

rp ra

Basic orbital elements (ellipse)

e=0: circle

e<1: ellipse



e e pse

e=1: parabola

e>1: hyperbola

vr

2.a

a: semimajor axis

e: eccentricity

v: true anomaly (0…360 deg)

rp: Radius of periapsis(perihelion)

ra: Radius of apoapsis(aphelion)

)1(

)1(

ear

ear

a

p

+=

−=ve

ear

cos1

)1(2

−

−=

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----Introduction to Solar System Dynamics----

Useful orbital parameters (elliptical orbit)

1) Velocity:

−=ar

GM u12

M: mass of central body

m: mass of orbiting body

r: distance of m from M



2) Period: aT

3

2π =

(M>>m)

3) Energy:

4) Angularmomentum:

a

GMm E

2−=

)1(

,2ea M Gm L

ur m L

−⋅⋅⋅=

×⋅=r r

(Constant!)

(Constant!)

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Spin stabilization

• With spin stabilization, the entire spacecraft rotates around its own

vertical axis, spinning like a top. This keeps the spacecraft's orientationin space under control.

• The spinning spacecraft resists perturbing forces



The spinning spacecraft resists perturbing forces.

• Desi ners of earl satellites used s in-stabilization for their satellites,

which most often have a cylinder shape and rotate at one revolution

every second.

• Spin stabilization was used for NASA's Pioneer 10 and 11 spacecraft,

the Lunar Prospector, and the Galileo Jupiter orbiter.

• The advantage of spin stabilization is that it is a very simple way to

keep the spacecraft pointed in a certain direction.

• A disadvantage of this stabilization is that the satellite cannot use

large solar arrays to obtain power from the Sun. Thus, it requires



large amounts of battery power.

antennas also must perform “despin” maneuvers so that antennas or

optical instruments point at their desired targets.

Reaction wheel stabilisation

• With three-axis stabilization, satellites have small spinning wheels,

called reaction wheels or momentum wheels, that rotate so as tokeep the satellite in the desired orientation in relation to the Earthand the Sun.

• If satellite sensors detect that the satellite is moving away from the



proper orientation, the spinning wheels speed up or slow down toreturn the satellite to its correct position.

-continually nudge the spacecraft back and forth to keep it within a

range of allowed positions.• Voyagers 1 and 2 stay in position using 3-axis stabilization.

• An advantage of 3-axis stabilization is that optical instruments

and antennas can point at desired targets without having to perform“despin” maneuvers

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Alignment

• There are a number of components which needalignment

– Solar panels



– Antennae

ese ve o o e e s o e s y

different times, so the problem is not trivial

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Solar and sidereal day



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ORBITAL ELEMENTSKeplerian Elements: True Anomaly

ν νν ν



i

ΩΩΩΩ

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Prepared by:

r s . .

Assistant Professor in ECE

Jyothi Engineering College, Cheruthuruthy

satellite mod 1

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