satellite mod1 final

8/6/2019 Satellite Mod1 Final

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EC04 705(D)Satellite Communication Systems

MODULE I


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BASIC

SATELLI

TE

SYSTEM


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A communication satellite can be looked upon as abig microwave repeater.

It contains several transponder which listens to someportion of the spectrum, amplifies the incomingsignal, and broadcasts it in another frequency toavoid interference with incoming signals.Can relay signals over long distances.


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S atellite Frequency B ands


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C B and

4GHz-downlink and 6Ghz uplink thefirst to be designatedMore channels on uplinkCapacity is low and terrestrialinterference is a problem


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K u band

12 Ghz (11.7 to 12.2) downlink

14 Ghz (14.0 to 14.5 ) -uplinkHigher capacity and less crowdedRain interference is main problem

A ntenna size is smaller compared tothose in C band.


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Basic

Satellite

Systems

Ground S tations (Earth S tations)-transmit RF signals to satellite

Received signals signal conditioned-retransmitted to other E S

Communication b/w all E S withincoverage area


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Basic

Satellite

System

S pace S egment1 or more satellites and suitable orbitsT elemetry T racking and Command ( TT& C) stationsRedundant systems

Ground S egmentFSS (Fixed S atellite S ervice)

- Fixed E S- A ntennas vary from 11-30m diameter - interface b/w user and E S Design consideration


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Ground segment contd

MSSS everal mobile terminals linked to fixed T elecommn n/w via

the satellite3 types-maritime, aeronautical and land based.T x /Rx are signals are affected by environment around themobile n/w.

DBS (Direct B roadband S ystem)Live/rec programms tx through large gateway E S -via a highpower satellite- to small terminals dispersed over the servicearea.T erminal antennas 50-100 cm diameter


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SS -Design ConsiderationT ype of data

Voice ,data, video etc.

T ype of serviceFSS ,MSS ,DBS

S election of Radio FrequencyDepends on type of application, propagation characteristics, stateof technology, availability of BW , RF regulations etc

S election of Optimal Modulation and coding S cheme.Depends on type of message, radio link characteristics,complexity

in E S


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SS -Design Consideration contdES size and complexityS ize and shape of the service areaT echnology related to satellite and G S .S atellite BW and power.


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A pplications and A dvantages

Coverage over large geographical area, remote areas etcHigh bandwidth.Cheaper over long distancesMobile commn, T V and sound broadcastsVideo distribution, internet

DisadvantagesHigh cost at introduction

Failure in satellite circuitry during launch or after deploymentaffects a large area.T ransmission delay caused by the long propagation path-ingeosat.S usceptible to noise and interferenceS ecurity can be an issue


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Geostationary satellites

A circular orbit at altitude 35,786 km abovethe equator geosynchronous orbit.S atellites rotate in unison with earth.A ppear stationary.Minimum operational requirements of E S .

Provides commn. to larger areas.( 1/3 rd of earth)3 GEO SAT at 120 degree can cover entirepart of earth.


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Geosat contd.


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Geostationary and Geosynchronous orbits

A geostationary orbit is a special orbit for which any satellitein that orbit will appear to hover stationary over a point onthe earth's surface.

For any orbit to be geostationary, it must first begeosynchronous.

A geosynchronous orbit is any orbit which has a period

equal to the earth's rotational period. This requirement is notsufficient to ensure a fixed position relative to the earth.

While all geostationary orbits must be geosynchronous, notall geosynchronous orbits are geostationary.


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Figure below shows the difference between a geostationary orbit (GSO) and ageosynchronous orbit (GEO) with an inclination of 20 degrees.Both are circular orbits.While each satellite will complete its orbit in the same time it takes the earth to rotate

once, the geosynchronous satellite will move north and south of the equator during itsorbit while the geostationary satellite will not.


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Geosat - A dvantages

Stable signal strength due to constant groundsatellite range.

Simple tracking systems - simple design of groundterminalsMinimum Doppler frequency shift on RF signals.

Coverage available to most populated areas of theworld.Time between launch and deployment is relativelysmall (few weeks).


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GEO SAT -DisadvantagesDue to the high orbit, the spatial resolution of the data isnot as great as for the polar orbiting satellites . Since they

are always positioned above the equator they can't see thenorth or south poles and are of limited use for latitudesgreater than 60-70 degrees north or south.Large propgn delay effects voice and time sensitive data

protocols.Larger path loss, higher latencyHigh transmitter power


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S OL A R D AY

A solar day is the length of time between twosuccessive passes of the sun across the same spot

in the sky (e.g. crossing the meridian, overhead).Because the Earth moves in its orbit around theSun, the Earth must rotate more than 360 degreesin one solar day .That time period is, on average, 24:00:00, hours,or one mean solar day.


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S IDERE A L D AY

A sidereal day is the length of time betweentwo successive passes of the fixed stars

across the sky.S idereal time is time kept with respect to thedistant stars.A sidereal day lasts from when a distant star

is on the meridian at a point on Earth until it isnext on the meridian.A sidereal day lasts 23 hours and 56 minutes(of solar time), about 4 minutes less than a

solar day.


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S olar vs. S idereal Day

(Source: M.Richaria, Satellite Communication Systems, Fig.2.7)


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S olar vs. S idereal Day

A sidereal day is the time between consecutivecrossings of any particular longitude on the earthby any star other than the sun.A solar day is the time between consecutivecrossings of any particular longitude of the earthby the sun-earth axis.

S olar day = EX A CT LY 24 hrsS idereal day = 23 h 56 min. 4.091 s


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ComparisionB ecause the stars are so distant from us, themotion of the Earth in its orbit makes

negligible difference in the direction to thestars. Hence, the Earth rotates 360 degreesin one sidereal day.

T he Earth must rotate an extra 0.986 degreesbetween solar crossings of the meridian.T herefore in 24 hours of solar time, the Earthrotates 360.986 degree.


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OR B ITA L P A R A MET ER S

Orbital elements are the parameters required to uniquelyidentify a specific orbit.

Exactly six parameters are necessary to unambiguously definean arbitrary and unperturbed orbit.

A scending node is the point where the satellite crosses theequatorial plane moving in the direction from south to north.

Descending node is the point where the satellite crosses theequatorial plane moving in the direction from north to south.

T he line joining these 2 nodes - Line of nodes.


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OR B ITA L P A R A MET ER S Contd


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Orbital parameters contd


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OR B ITA L ELEMEN TS; Right A scension of the A scendingNode

i Inclination of the orbit[ A rgument of Perigeetp T ime of Perigeee Eccentricity of the elliptical orbita S emi-major axis of the orbit ellipse


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Orbital parameters contd

T he semi-major axis, a describes the size of a conic orbit(conic/elliptical).T he eccentricity, e , shows the ellipticity of the orbit.

T he inclination, i ,-angle between the plane of then orbit and theequatorial plane measured at the ascending node in thenorthward direction.T he right ascension of an ascending node, ,is the anglebetween the x axis and the ascending node.T

he argument of perigee, is the angle in the orbital plane b/wthe line of nodes and the perigee of the orbit.T ime tp is the time elapsed since the satellite passed theperigee.


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OR B IT CH A R A CT ERI ST ICS

Semi-Axis Lengths of the Orbit

21 e p

a !where

Q

2h p !

and h is the magnitude of the angular momentum

2/121 eab ! wheree is the eccentricity of the orbit


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OR B IT ECCEN T RICI TY

If a = semi-major axis,b = semi-minor axis, ande = eccentricity of the orbit ellipse,

then

baba

e !N OTE: For a circular orbit, a = b and e = 0


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Low Earth Orbit Concepts

Equator

South Pole

Ground track

Ascendingnode

Inclinationangle

Descending node

Orbit

Perigee

Apogee

Orbit


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K EPLER S T HREE L AWS

L AWS OF PL A NE TA RY MOT ION

The orbit of each planet follows an elliptical path in space, thesun being the focus.The satellite sweeps out equal arcs (area) in equal time( NOTE : for an ellipse, this means that the orbital velocity

varies around the orbit)The square of the period of revolution equals a constant v thecube of semi-major axis of the ellipse


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K EPLER 1: E lliptical Orbits

Law 1

T he orbit is anellipse

e = ellipses eccentricity

O = center of the earth (onefocus of the ellipse)C = center of the ellipse


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K EPLER 1: E lliptical Orbits (cont.)

(describes a conicsection, which is anellipse if e < 1)

)cos(*1 00 J e

pr !

e = eccentricitye


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K EPLER 2: Equ al Arc-Sweeps

Law 2

If t2

- t1

= t4

- t3

then A 12 = A 34

Velocity of satellite isSLOW E ST at

APOG EE; F AST E ST at P ERI G EE


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K EPLER 3: Orbital Period Orbital period and the Ellipse are related by

T 2

= (4 T 2

a3

) / QT hat is the square of the period of revolution is equal to aconstant v the cube of the semi-major axis.

Q = K eplers Constant =GM E


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EFFEC T OF E A RT HN on uniform distribution of earths mass.Ellipsoid with slight bulge at equator

Variation in gravitational pull


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Main effects of perturbation on a satellite :

1.Perigee of the elliptical orbit rotates in theorbital plane.

The rate of change of argument of perigee (in degrees/day) isgiven by

tyeccentricie

ninclinatioi

axismajor -semia

km)6378radius(~equatorialmeanR where

deg/day)1(

1)(cos597.4 22

25.3

!!!

!

!

y

ei

a R

[

W hen i=63.4 , the tends to be zero. hence perigee is constant.y

[


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Orbits with i=63.4:

Molniya OrbitApogee=~40,000km,Perigee=~1000km

Tundra OrbitApogee=~46,300km,Perigee=~25,250km

AdvantagesCoverage for higher latitude locationsLand mobile communication to higher latitudes


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Main effects of perturbation on a satellite contd..

2.The orbital plane rotates around the earthsnorth-south axis.

The rate of precession of the ascending node indegrees/day is given by

tyeccentricie

ninclinatioi

axismajor -semia kmsindistancegeocentre-satelliter where

deg/day)1()(cos

95.9 22

5.3

!!!!

!;

y

ei

ar


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Rotation is opposite to satellite motion~ 4.9 /year for geostationary orbits

The ascending node rotates around the earth in ~73years.


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Nodal Regression (nodal precession)N odal regression refers to the shift of the plane of an orbitunder the gravitational force of Earth's (or any planet's)equatorial bulge.

For low orbit satellites, it can be as much as 6 to 8 degrees per day westward (for example, at inclinations of 52degrees and 28 degrees respectively).

The regression rate depends on altitude (the higher, thelower the rate) and inclination (the higher, the lower therate).


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Main effects of perturbation on a satellite

contd3.The component of perturbating force along the orbital plane

imparts a force vector on a satellite.

Mainly effects the geostationary satellites.The gravitational force on the satellite is directed towards the nearestequatorial bulge instead of earths centre-producing a component of forceon the orbital plane.Since geosynchronous orbit is constant w.r.t earth, the perturbations addsup to cause a drift of satellite to one of the stable points.Stable points on minor axis.Max amplitude of drift acceleration=~+0.0018/ day 2.Drifts cancelled by firing regular thrusters.


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Gravitational effects from other heavenly bodies

For LEOGravitational pull from earth> sun,moon

For GeoSATGravitaional pull from sun,moon is high.Gravity gradient -- higher force when SAT nears theheavenly body.

i (inclination) of the orbit changes.The orbit normal moves towards the vernal equinoxSun+moon effects =i~ 0.75 to 0.94 .


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Gravitational effects from other heavenly bodies on GeoSAT

contd

Variation in i due to moon =0.48 (min) to 0.67 (max)This cyclic variation is due to effect of sun on moons

orbit.Max variation when orbit normal of moon and Satelliteare maximum apart.

Variation of i due to sun=~0.27 /year almost constant.


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Gravitational effects from other heavenly bodies on GeoSAT

contd

3 main forces effects i of Geostationary satellitesG.F of Sun and moon acting along the same direction

Force due to non spherical nature of earth (resulting in)acting opposite to the initial 2 forces.

When i=7.5 ,the forces cancel out.

If not corrected, i of geoSAT oscillates between 0 to 15 with a period of 53 years.

y;


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Solar Radiation PressureI ncreases as surface area of satellite projected towardssun increases.Increases as size of solar arrays increases.

The SRP on geostationary satellite results in adisturbing torque along the north-south axis of satellite.The SRP increases the orbital eccentricity.

Corrected periodically


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S un T ransit Outage

A sun outage , sun transit or sun fade is an interruption in or distortionof geostationary satellite signals caused by interference from solar radiation.

The effect is due to the sun's radiation overwhelming the satellite signal.Generally, sun outages occur in February, March, September and October

At these times, the apparent path of the sun across the sky takes itdirectly behind the line of sight between an earth station and a satellite.

As the sun radiates strongly at the microwave frequencies used tocommunicate with satellites (C-band, Ka band and Ku band) the sunswamps the signal from the satellite.

The effects of a sun outage can include partial degradation, that is, anincrease in the error rate, or total destruction of the signal.


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Other effects

Earths magnetic FieldMeteoritesS elf generated torque and pressuresdue to RF radiation from the antenna.


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A tmospheric DragEffects the LEO mostly.Friction due to collision with atoms and ions.Reduces the ellipticity of the orbitA t lower orbital altitudes. S atellites faceextensive heat due to friction and burns out

Lower limit for LEO satellites=180kms


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S atellite lifeS atellite (operational) life depends onequipment life,fuel capacity of satelliteetcOrbital life-LEO have lesser life thanMEO and GEO.


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Doppler effectFreq of RF signals appears to be increasing as satapproaches the ground observer and appears to bedecreasing as SAT moves away .

Compensations for DSFrequency budget for RX

Using Rx filter BW ,uncertainties /drifts in ES and SAT Localoscillators ,DS variations are analyzed .

Monitoring stationsMaintains correct RF frequency by continuously monitoring

parameters of each carrier in the n/w. The Doppler componentis removed from The ES and spacecraft signals by comparingwith values measured in the monitoring stations.


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Doppler effectT he Doppler shift f d at frequency f1

vt =relative radial velocity b/w the observer and the transmitterC=velocity of light f1=transmission frequency


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SAT ELLIT E P AT HS atellite T rajectory

A ssumptionsT he bodies of earth and satellite are symmetricspherically point masses.No other forces act on the system other thangravitational forces.T he mass of earth is much greater than that of thesatellite.


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Related expressions


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Related expressions.S ubstituting eqn1 in 2 we get


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T he trajectory eqn.Solving for r from the above eqn, the trajectory eqn isgiven by


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T he trajectory eqn.

( T he fig.describes aconic section, which isan ellipse if e < 1)

)cos(*10 Ue p

r !

e = eccentricitye


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T he satellite periodFrom K eplers third law, we know that the period of a satellite dependsonly on the semi major axis, a

where =GM E

For circular orbits, T he satellite period is given by..

W here R=radius of earthh= satellite altitude


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S atellite velocity Velocity of a satellite is given by

W here V=velocityr=distance from eartha-semi major axis

For circular orbits

a=r


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S atellite positionO is centre of earthC is centre of ellipseAt any time tp satellite is at position S.tp is measured with respect to perigee

A circle is drawn with C as centre and a (semi major axis) as the radius.

Draw BM Perpendicular ,passing through S.E (angle BCM) = eccentric anomaly(angle SOM) = True Anomaly

Mean Anomaly (M) = angle travelled by satellite from perigee,in the same time tp,movingat the average angular orbital velocity


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S atellite position contd..

M is the mean anomaly,E is eccentric anomaly

E in radians and is true anomaly

When e=0, E=M=


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S atellite position contd..

The distance between satellite and Geocentre is given by

Position of satellite relative to an earth Station is also given by satellite azimuth and elevation.

Elevation ( )=angular distance along the vertical circle,from the horizon to the satellite location.

Azimuth ( ) = angle between direction of true north anddirection of satellite measured in clockwise direction.


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Azimuth and elevation from a point T on earth


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Azimuth and elevation -Figure


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The ElevationThe elevation angle, also called the altitude, of anobserved object is determined by first finding the

compass bearing on the horizon relative to truenorth, and then measuring the angle between that point and the object, from the reference frame of the observer .


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AzimuthThe azimuth angle is the compass bearing, relativeto true (geographic) north, of a point on the

horizon directly beneath an observed object. Thehorizon is defined as a huge, imaginary circlecentered on the observer, equidistant from thezenith (point straight overhead) and the nadir

(point exactly opposite the zenith).


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Elevation of a satellite is also defined as the angle which the satellitemakes with the tangent at a specific point on the earth .The elevation isgiven by


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Selection of orbital slot

The service area should be served as high as possible.

Satellite eclipses should occur max. at night to minimize usage of storage batteries.

Maintain sufficient orbital distance from nearby satellites sharingsame frequency to minimize interference.

Considering the prevailing radio regulations, applications et.c


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Non-Geostationary satellitesConstellation designbased on:

InclinationA ltitudeEccentricity


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1.Based on inclination:

Type 1: Orbital planes have common intersecting point* Coverage towards common intersection point

Eg:Polar constellation

Type 2: Orbital planes are distributed and hence satellites areuniformly distributed.

*coverage uniformaly distributed

In both type the satellites have same time period.


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2.Based on eccentricity

Hybrid constellations: combination of circular andelliptical orbits of different altitudes

3.Based on AltitudeLEOMEO

GEO


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Orbits of DifferentS

atellites

Earth

1000 km

35,768 km

10,000 km

LEO (Iridium) GEO

HEO

MEO

Not drawn to scale


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L A UNCHING OF A GEO STAT ION A RY SAT

ELLIT

ET otal energy of a satellite is given by

satellitet h eof velocityv

satellitetoeart hof geocentre fromcedi sr

eart hof ma ssM

satelliteof ma ssmr

M mGvmU

!

!

!

!

!

tan

21 2


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LAU N CH IN G OF A GEOSTAT I O N ARY

SATELLITE

N eeds to attain a velocity of 3070m/s at the geostationary orbitheight of 42,165km from earths centre (~36,000km fromearths surface).

Max increment velocity v that a launch vehicle of Mass mocan impart is given by

fuel anded t h eof ma ssmf

de s ignnozzlerocket t h eand fuel of

typeondepend s ga st h eof velocityexh au s t effectivevg mo

mf vg v

exp

(

1

1ln

!

!

!(


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SATELLITE

To increase v ,mf/mo must increasei.E mo must decrease

Multiple stage rockets used

Each stage is jettisoned after imparting a trust to attain the final v

As mo decreases, the final stages needto give lower thrusts to attain v.

v final =sum of velocity incrementsof all stages


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HOHM A

NNT

R A

NS

FERA transfer b/w 2 coplanar circular orbits via an ellipticaltransfer orbit requires minimum velocity increment andfuel.

The orbital inclination depends on the latitude of the launchingstation and is given byCos(i)=sin( 1)cos( 1)

1-azimuth of launch1- latitude of launch site


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SATELLITE

i=minimum, for easterly launch ( 1=90 ).

Launch is closer to the equator to make max. advantage of earths rotational velocity and minimize fuel to bring downi to 0.

Vertical launch minimizes atmospheric drag.

The guidance system gradually tilts the vehicle to 90 east.

I nitially launched to a low earth parking orbit'. I gnition iscut off, the system drifts in the parking orbit.


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SATELLITE contd

N ext, just before reaching the equator, second stage rocketis ignited. The satellite is injected to elliptical orbit withapogee of a geostationary orbit and its line of nodes is inthe equatorial plane.

I n the elliptical orbit, the payload is separated from thelaunch vehicle.

After several revolutions, accurate set of satellite orbital parameters are attained.

The apogee kick motor (part of payload) is fired to convertelliptical to circular orbit.


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SATELLITE contd

At the apogee of elliptical orbit, the remainingincremental velocity is imparted normal to theorbital plane and the satellite is injected togeostationary orbit.

Satellite drifts slowly w.r.t to earth-drift phase.

Final corrections on i and other orbital parametersare done by final thrusters.


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EXPE N DABLE LAU N CHER -AQU I R IN G

PARK IN

G ORBIT

0 - Vertical lift off

1-Rocket tilting eastward

2-First stage drop off

3-Second stage ignition

4-Insertion to parking orbit (18 0 -25 0 kms) from earth

5-Second and third stageignitions at equator crossings.


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TRA N SFER ORB IT A N D F IN AL SATELL ITE POS ITIO NIN G


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TRA N SFER ORB I T A N D F IN AL SATELL I TE

POSIT

IO

NING

1. Velocity increment to attain transfer incrementSatellite spun for stabilization

2 Apogee kick motor fired for velocity incrementOrbital circularizedi~=0

3 Satellite is despun4 3 axis stabilization acquired

5 Minor orbital correctionsMinimize residual orbital errorsPositioning of satellite


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Final corrections on satellite parametersTransition of stabilization from spin mode to 3 axisstabilizations.Solar array deployment.Sun-earth acquisition.I n orbit tests checking the satellite performance beforereplacing the existing operational satellite.Drifts due to various perturbations are monitored

periodically


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LAU N

CH FROM SPACE SHUTTLE

Expendable launch-lose of most of the h/w duringlaunch.

Space shuttle-reusable launch vehicleRetrieves and repairs satellite in low orbitsThe shuttle contains a reusable orbital-injectssatellite to LO, re-enters the atmosphere asaircraft .


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LAU N

CH FROM SPACE SHUTTLEThe orbiter-launched vertically-using 2 recoverable solidrocket boosters.Liquid hydrogen/oxygen tank is the propellant for 3 mainengines-expendableI nitial inclination of i=28 ,in the parking orbit.The shuttle carries payloads of masses ~29500 kg (LEO)and 14500kg (polar orbits)Launch only upto LEO-additional propulsion to inject togeostationary orbits.


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LAU N

CH WIN

DOWLimitations on launch time to certainspecified intervals of time:

Position of Satellite is favorable w.r.t sun: to ensureadequate power supply, and thermal control throughoutthe missionSatellite should be visible to the control stations during

all critical maneuvers.

satellite mod1 final

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