satellite communication systems - edutalks.orgedutalks.org/downloads/satellite mod 1.pdf · what...
TRANSCRIPT
SATELLITE COMMUNICATION SYSTEMS
Girish K.P.
www.edutalks.org
Communication satellites bring the world to you anywhere and any time…..
www.edutalks.org
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 around another (usually much larger) one in a mathematically moves around another (usually much larger) one in a mathematically predictable 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 for communication.
www.edutalks.org
How do satellite work?
www.edutalks.org
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 linkfrequency is the frequency at which Ground Station is frequency is the frequency at which Ground Station is communicating with Satellite.
* The satellite Transponder converts the signal and sends it down to the second earth station. This frequency is called a Downlink.
www.edutalks.org
Consider the light bulb example:
www.edutalks.org
Components of a satellite
www.edutalks.org
Advantages of satellite over terrestrial communication :
* The coverage area of a satellite greatly exceeds that of aterrestrial system.
* Transmission cost of a satellite is independent of the distancefrom 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 communicationthan in terrestrial communication.
www.edutalks.org
How does a satellite stay in it’s orbit?
www.edutalks.org
How do we escape gravity & place an object in orbit?
• If an object is fired fast enough it should escape the earths pull.escape the earths pull.
• This is done through the use of Rocket Launchers
www.edutalks.org
Multi-stage Rockets
• Stage 1: Raises the payload e.g. a satellite to an elevation of about 50 miles.
• Stage 2: Satellite 100 miles and the third stage
• Stage 2: Satellite 100 miles and the third stage places it into the transfer orbit.
• Stage 3: The satellite is placed in its final geo-synchronous orbital slot by the AKM, a type of rocket used to move the satellite.
www.edutalks.org
Applications
www.edutalks.org
Major problems for satellites
• Positioning in orbit
• Stability
• Power
• Communications
• Harsh environment
www.edutalks.org
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
• Often it is the fuel availability which • Often it is the fuel availability which determines the lifetime of a satellite
• Commercial life of a satellite typically 10-15 years
www.edutalks.org
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 in order to provide stable communications.
* Modern satellites use reaction wheel stabilisation - a form of gyroscopic stabilisation.
www.edutalks.org
Power
• Modern satellites use a variety of power means
• Solar panels are now quite efficient, so solar power is used to generate electricitypower is used to generate electricity
• 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
www.edutalks.org
Satellite - satellite communication• It is also possible for
satellites to communicate with other satellites
• Communication can be by microwave or by optical laserby optical laser
1.
2.
1.
2.
1.
2.
Point-Point System Crosslink System Hybrid System
www.edutalks.org
Harsh Environment
• Satellite components need to be specially “hardened”
• Circuits which work on the ground will fail very rapidly in space
• Temperature is also a problem - so satellites use • Temperature is also a problem - so satellites use electric heaters to keep circuits and other vital parts warmed up - they also need to control the temperature carefully
www.edutalks.org
Early satellites• Telstar
– Allowed live transmission across the Atlantic• Syncom 2
– First Geosynchronous satellite
TELSTAR SYNCOM 2
www.edutalks.org
Satellite orbits
Classification of orbits:
www.edutalks.org
* Circular orbits are simplest
* Inclined orbits are useful for coverage of equatorial regions
* Elliptical orbits can be used to give quasi stationary behavior viewed from earth using 3 or stationary behavior viewed from earth using 3 or 4 satellites
* Orbit changes can be used to extend the life of satellites
www.edutalks.org
Classification of orbits:
Satellite orbits are also classified based on their heights above the earth:
– GEO– LEO– MEO– Molniya Orbit– HAPs
www.edutalks.org
Satellite orbit altitudes
www.edutalks.org
Geostationary Earth Orbit (GEO)• These satellites are in orbit 35,786 km above the earth’s
surface along the equator.• Objects in Geostationary orbit revolve around the earth
at the same speed as the earth rotates. This means GEO satellites remain in the same position relative to the surface of earth.
www.edutalks.org
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 doppler shift– Minimal doppler 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.
www.edutalks.org
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 be
useful
www.edutalks.org
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
www.edutalks.org
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 point communication.
A LEO satellite’s smaller area of coverage is less of a waste of bandwidth.of bandwidth.
• Disadvantages A network of LEO satellites is needed, which can be costly
LEO satellites have to compensate for Doppler shifts cause by their relative movement.
Atmospheric drag effects LEO satellites, causing gradual orbital deterioration.
www.edutalks.org
Medium Earth Orbit (MEO)• A MEO satellite is in orbit somewhere between 8,000
km 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 LEO • MEO satellites have a larger coverage area than LEO
satellites.
www.edutalks.org
MEO contd.
• Advantage A MEO satellite’s longer duration of visibility and
wider footprint means fewer satellites are needed in a MEO 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.
www.edutalks.org
MEO satellites
Glonass (Russian)
The GPS constellation calls for 24 satellites to be distributed equally among six circular orbital planes
www.edutalks.org
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.
www.edutalks.org
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 a comparatively strong signal.
Cheaper to put in position, but would require a lot of them in a network.
www.edutalks.org
HAP
www.edutalks.org
Satellite frequency band
Band Downlink,
GHzUplink, GHz
Bandwidth, MHz
L 1.5 1.6 15
S 1.9 2.2 70S 1.9 2.2 70
C 4 6 500
Ku 11 14 500
Ka 20 30 3500
www.edutalks.org
Solar day and Sidereal day
• A day is defined as the time that it takes the Earth to rotate on its axis.
• However, there is more than one way to define a day:
• 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 with
respect to the Sun.
www.edutalks.org
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.
• Astronomers use sidereal time because we are • Astronomers use sidereal time because we are mostly interested in distant celestial objects.
www.edutalks.org
Solar day and Sidereal day• A solar day is measured using the passage of the
Sun across the sky—it lasts 24 hours• A sidereal day (from the Latin word meaning
star) is measured with respect to fixed stars—it lasts a little less than 24 hours.
• Each solar day the Earth rotates 360 degrees with • Each solar day the Earth rotates 360 degrees with respect 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 stars
www.edutalks.org
• The actual length of a sidereal day on Earth is 23 hours 56 minutes 4 seconds
• This means that the Earth has to rotate slightly more than one turn with respect to a fixed star to reach the same Earth-Sun orientation (solar day)
www.edutalks.org
Solar day and Sidereal day
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 same position in the sky 4 minutes earlier on each consecutive day.consecutive day.
www.edutalks.org
Apparent Solar Time
• Apparent solar time is the time measured with respect to the actual position of the Sun.– At noon, the Sun would be exactly on the meridian.– 1 P.M. would be exactly one hour after the Sun was on – 1 P.M. would be exactly one hour after the Sun was on
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.
www.edutalks.org
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
www.edutalks.org
Kepler’s 1st Law: Law of Ellipses
The orbits of the planets are ellipses with the sun at one focus
www.edutalks.org
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
T2
T6
T2T1A2A3A4A5
A6
A1
www.edutalks.org
Kepler’s 3rd Law: Law of Harmonics
The squares of the periods of two planets’ orbits are
proportional to each other as the cubes of their semi-
major axes:major axes:T12/T22 = a13/a23
In English:
Orbits with the same semi-major axis will have the
same period
www.edutalks.org
Newton’s Laws• Kepler’s laws only describe the planetary motion
without attempting to suggest any explanation as to why the motion takes place in that manner.
• Derived three laws of motion.
Sir .Issac Newton
• Derived three laws of motion.
• Derived the Law of Universal Gravitation.
• Explained why Kepler’s laws worked.
www.edutalks.org
Newton’s 1st Law: Law of Inertia
• Every body continues in a state of uniform motion unless it is compelled to change that state by a force imposed upon itstate by a force imposed upon it
www.edutalks.org
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
• Change in momentum gives: F = ma• Change in momentum gives: F = ma
F
Fwww.edutalks.org
Newton’s 3rd Law: Action - Reaction
• For every action, there is an equal and opposite reaction
• Hints at conservation of momentum• Hints at conservation of momentum
www.edutalks.org
Newton’s Law of Universal Gravitation
Between any two objects there exists a force of attraction that is proportional to the product of their masses and inversely proportional to the square of the distance between themthe square of the distance between them
Fg = G( )M1m2
r2
www.edutalks.org
Classical orbital elements
www.edutalks.org
Apogee and Perigee• In astronomy, an apsis is the point of greatest or least distance
of 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 the apoapsis (apogee)
• A straight line drawn through the perigee and apogee is the line • 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.
www.edutalks.org
Classical orbital elements• Six independent quantities are sufficient to
describe the size, shape and orientation of an orbit.
These are– a, the semi-major axis– a, the semi-major axis– εεεε, the eccentricity– i, the inclination– ΩΩΩΩ, the right ascension of the ascending node– ωωωω, the argument of perigee– tp, mean anamoly
www.edutalks.org
• The semi-major axis describes the size of the orbit. It connects the geometric center of the orbital ellipse with the periapsis, passing through the focal point where the center of mass resides.
• The eccentricity shows the ellipticity of the orbit. • The inclination is the angle between the plane of the
orbit and the equatorial plane measured at the ascending node in the northward direction.node in the northward direction.
• The right ascension of an ascending node is the angle between 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 satellite passed the perigee.
www.edutalks.org
Major parameters of an elliptical orbit
• Satellite trajectory • Satellite period • Satellite velocity • Satellite position
www.edutalks.org
Satellite TrajectoryThe path of a satellite in space may be obtained
under 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.besides the gravitational forces.
3.The mass of the earth is much greater than satellite.
These assumptions lead to the two body problem.
www.edutalks.org
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 = 0
r3 r3
..Where r = vector acceleration in the given coordinate system
r = vector from M (mass of earth) to m (mass of satellite)r = vector from M (mass of earth) to m (mass of satellite)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 focusф = flight elevation angle , v = satellite velocitya = semi-major axis = (ra+rb)/2
www.edutalks.org
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
www.edutalks.org
Satellite velocityTotal specific mechanical energy ε of a satellite is the sum
of 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 gains speed as it loses height.
The velocity of a satellite in an elliptic orbit is :The velocity of a satellite in an elliptic orbit is :
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
www.edutalks.org
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 through the point S.Angle E is called eccentric anomaly and angle θ is the true anomaly.
www.edutalks.org
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
E = arccos[ (e + cosθ)/(1+ e cosθ)] E = arccos[ (e + cosθ)/(1+ e cosθ)]
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 is
r = a(1-e2)/(1-ecosθ)
www.edutalks.org
GEOSYNCHRONOUSAND
GEOSTATIONARY ORBITS
www.edutalks.org
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 minutes and 4 seconds in length.minutes and 4 seconds in length.
• A geosynchronous earth orbit is sometimes referred to as the Clarke orbit or Clarke belt, after Arthur C. Clarke, who first suggested its‘ existence in 1945 and proposed its use for communications satellites
www.edutalks.org
Clarke Orbit• The Clarke orbit meets the
concise set of specificationsfor geosynchronous satelliteorbits:– (1) be located directly above
the equatorthe equator
– (2) travel in the samedirection as Earth's rotation at6840 mph
– (3) have an altitude of 22,300miles above Earth
– (4) complete one revolution in24 hours
www.edutalks.org
Clarke Orbit
www.edutalks.org
Geo synchronous Satellites• There is only one geosynchronous earth orbit.
• It is occupied by a large number of satellites. In fact, thegeosynchronous orbit is the most widely used earth orbit for theobvious reason.
• An international agreement initially mandated that all satellitesplaced in the Clarke orbit must be separated by at least 1833 miles.placed in the Clarke orbit must be separated by at least 1833 miles.
• 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.
www.edutalks.org
Geo stationary orbit• A geostationary orbit is a special case of a
geosynchronous orbit.• A satellite is in a geostationary orbit when it appears
stationary from the point of view of an observer on the Earth's surface.
This can only occur when: This can only 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 geostationary
only with the additional restrictions of it being in a circular orbit situated over the equator.
www.edutalks.org
Geostationary Vs. Polar
Orbiting
http://cimss.ssec.wisc.edu/satmet/modules/sat_basics/images/orbits.jpg
www.edutalks.org
Geostationary Satellites
The satellite velocity in this orbit is 3075 m/s.
• Operate in the 2.0 GHz to 18 GHz range
• When the inclination and • When the inclination and eccentricity of the orbit is zero, the satellite appears to be stationary to an observer from ground.
www.edutalks.org
Geostationary Satellites in Orbit
www.edutalks.org
www.edutalks.org
Geostationary Satellite Coverage
http://www.ssec.wisc.edu/mcidas
www.edutalks.org
Geostationary Satellite Coverage
http://www.ssec.wisc.edu/mcidas
www.edutalks.org
Geostationary Satellite Coverage
http://www.ssec.wisc.edu/mcidas
www.edutalks.org
Geo-stationary satellites
Applications:Telecommunication systemsRadioData Transmission systems
• The geometric considerations like satellite elevation/look angle etc are very vital for reliable communication satellite system design.
www.edutalks.org
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]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 equatorTilt angle γ = arc tan [sin ψ / (6.6235-cos ψ)
www.edutalks.org
Azimuth:The azimuth ξ is the angle which the satellite direction makeswith 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 stationξ =1800 - A0;when the satellite is to the east of earth stationin southern hemisphere:ξ =3600 - A0;when the satellite is to the west of earth stationξ =3600 - A0;when the satellite is to the west of earth stationξ =A0;when the satellite is to the east of earth station
www.edutalks.org
Range:The range d of a geostationary satellite is given by,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 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).
www.edutalks.org
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 isreignited and the spacecraft is launched into ageostationary transfer orbit(GTO).
• Perigee of GTO is that of LEO altitude andapogee that of GEO altitude.apogee that of GEO altitude.
• After a few orbits in GTO, while the orbital parameters are measured, a rocket motor (AKM)is ignited at apogee and GTO is raised until it iscircular geostationary orbit.
• AKM (Apogee Kick Motor) is used to circularizethe orbit at GEO and to remove any inclinationerror so that the final orbit is very close togeostationary.
www.edutalks.org
Launching of geostationary satellite
www.edutalks.org
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 nothing 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".
• If we do nothing else, it will stay in this elliptical orbit, going from apogee to perigee and back again.
www.edutalks.org
Phase I and II of launching spacecraft
Phase II
Few Geostationary satellites: EDUSAT, INTELSAT , INSAT , PAKSAT, AMERICOM …….
Phase I ↑
Phase II
www.edutalks.org
ORBITAL MANEUVERS
Hohmann Transfer– Can be used to raise or lower
altitude– Most efficient method– At minimum, requires
completion of half revolution of transfer orbit
www.edutalks.org
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 equatorial plane.• Finally satellite is injected into GEO by imparting a • Finally satellite is injected into GEO by imparting a velocity increment at the apogee equal to the difference between 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.
www.edutalks.org
• A Hohmann transfer is a fuel efficient way to transfer from 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 the direction the vehicle is traveling.
• This will add velocity to the vehicle causing its trajectory to become an elliptic orbit (B). This elliptic orbit is carefully designed to reach the desired final altitude of the higher orbit (C).carefully designed to reach the desired final altitude of the higher orbit (C).
• In this way the elliptic orbit or transfer orbit is tangent to both the original orbit (A) and the final orbit (C). This is why a Hohmann transfer is fuel efficient.
• When the target altitude is reached the engine is fired in the same manner as before but this time the added velocity is planned such that the elliptic transfer orbit is circularized at the new altitude of orbit (C).
www.edutalks.org
Hohmann TransferTarget Orbit
Initial Orbit
Transfer Orbit
The orbital inclination is given by,cos i= sinξ1 cos θ1
where i=inclinationξ1 =azimuth of launchθ1 =latitude of launching site
www.edutalks.org
PERTURBATIONS
• Perturbation is a term used in astronomy to describe alterations to an object's orbit caused by gravitational interactions with other bodies.
Major sources are: Major sources are: • Effect of earth
• Third Body Effects
• Atmospheric Drag
• Solar radiation pressure
• Electro-Magnetic effect
www.edutalks.org
Space Weather EffectsSpace Weather Effects
www.edutalks.org
Effect of earth on satellites
• The effect of gravitational force is non uniform because of the non uniform distribution of earth’s mass -a slight bulge at the equator, with a difference of 21 km between polar and the equator radius.
• This deviation from spherical shape causes additional forces on the causes additional forces on the satellite.
• The effect of earth’s gravitational pull may be expressed as the harmonic series of the field. The first term represents the principal gravitational law and the higher order terms in the series as the perturbations.
www.edutalks.org
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.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.
www.edutalks.org
2. The orbital plane rotates around the earth’s north-south axis.The rate of change of rotation of ascending node isΩ = 9.95[r/a]3.5 cos i /(1-e2)2 deg/dayWhere 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.
3. The perturbating force along the orbital plane imparts a 3. The perturbating force along the orbital plane imparts 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 nearest stable points on the orbit.Stable points are approximately on the minor axis, showing that the elliptical approx. of earth is not precisely accurate.
www.edutalks.org
Third Body Effects (heavenly bodies)
• Gravitational pull of other massive bodies, i.e. Sun, moon
• Mainly noticeable in deep space orbits
www.edutalks.org
Gravitational effects from heavenly bodies:• In LEO satellites, the influence of gravitational forces
from sun and moon are small when compared to the gravitational force of earth.
• The order of magnitude of gravitational force of moon and sun are main sources of perturbations in GEO satellites.
• When nearer to heavenly bodies, the gravitational pull is • When nearer to heavenly bodies, the gravitational pull is stronger and hence causes a gravity gradient. main effect of such gradient is to change the inclination of the orbit.
• The combined effect of sun and moon is to cause a change in inclination of GEO satellites between 0.750
and 0.940
www.edutalks.org
• 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 1987 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 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 a maximum of 150 and a minimum of 00
www.edutalks.org
Atmospheric Drag
• Satellites below 2000 kilometres, are actually travelling through the Earth’s atmosphere. Collisions with air particles, even at these high altitudes slowly act to circularise the orbit and slow down the spacecraft causing it to drop to lower altitudes , this effect is known as atmospheric dragknown as atmospheric drag
• Emissions from the Sun cause the upper atmosphere to heat and expand.
• These energetic solar outputs increase dramatically during periods of high solar activity, and may result in Earth-orbiting satellites experiencing an increase in atmospheric drag
www.edutalks.org
A satellite orbiting the Earth would continue to orbit forever if gravity were the only force acting on it.
Perigee remains same, Apogee decreases
www.edutalks.org
• Reduces satellite’s energy• Changes the size (semi-major axis) and shape
(eccentricity)• The effect of drag is more severe at about 180km and
causes excess heat on satellite .Unless such LEO satellites are routinely boosted to higher orbits, they slowly fall, and eventually burn up
• Orbital life time of satellite at 400km circular earth orbit is typically few months, where as the life time is several typically few months, where as the life time is several decades if they are at 800km altitude
• In the former case, functional life time depends on orbital life time and for latter the life time of satellite equipments is the deciding factor.
• However, for GEO satellites the governing factors are equipment life time and fuel capacity of the satellite (typically 10-15 years).
www.edutalks.org
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
surface area of the satellite projected in thedirection of sun increases.
• The net effect is the increase in the orbital eccentricity and also introduces disturbing torque that effects the north-south axis of the satellite.north-south axis of the satellite.
www.edutalks.org
• Solar wind causes radiation pressure on the satellite• The solar wind is a stream of charged particles (a plasma) that are
ejected from the upper atmosphere of the sun. It consists mostly of electrons and protons with energies of about 1 keV. These particles are able to escape the sun's gravity because of the high temperature of the corona, and also because of high kinetic energy
• These perturbations are corrected periodicallywww.edutalks.org
Effects of Solar radiation pressure
1. HUMAN HEALTH• Intense solar flares produce very high energy particles
that can be as harmful to people as low-energy radiation from nuclear blasts. Earth’s atmosphere and magnetosphere provide protection for people on the ground, but astronauts in space are subject to potentially lethal doses of radiation. The penetration of potentially lethal doses of radiation. The penetration of high energy particles into living cells leads to chromosome damage and, potentially, cancer.
• Airline pilots and flight crews, as well as frequent fliers, also receive increased doses of radiation from solar flares. If you were travelling in an aircraft at high altitudes during a major solar flare, the amount of radiation you would be exposed to can be equivalent to getting a chest x-ray.
www.edutalks.org
2. COMMUNICATIONS
• Stormy space weather can damage Earth-orbitingsatellites such as those carrying TV and mobile phone signals.
• During high levels of solar activity, satellites are bombarded with high energy particles. If the deeply penetrating electrons build up faster than deeply penetrating electrons build up faster than the charges are able to dissipate out of the satellite material, a discharge can result that is capable of damaging the satellite electronics.
• These processes can result loss of control and even satellite failure.
www.edutalks.org
3. NAVIGATION• A Global Positioning System (GPS) receiver uses radio signals from
several orbiting satellites to determine the range, or distance, from each satellite, and determines from these ranges the actual position of the receiver.
• The radio signals must pass through the ionosphere, the uppermost part of the Earth’s atmosphere, and in doing so are subjected to variations in the electron density structure of the ionosphere.
• Changes in the electron density due to space weather activity can • Changes in the electron density due to space weather activity can change the speed at which the radio waves travel introducing a “propagation delay” in the GPS signal. Changing propagation delays cause errors in the determination of the range.
• An increase in space weather activity may cause widespread disruption to aircraft and ship navigation and emergency location systems that rely heavily on satellite navigation data.
www.edutalks.org
Electro-Magnetic effect
• Interaction between the Earth’s magnetic field and the satellite’s electro-magnetic field results in magnetic drag
www.edutalks.org
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 decrease depending 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 days.
www.edutalks.org
Non geostationary constellations
• The design of constellations can be categorized according to inclination, altitude and eccentricity.
On the basis of inclination, two types of constellations are designed.
Type I constellations are those having their orbital planes with a common intersection point.
Eg: Polar constellationsEg: 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 a
system (A LEO satellite can be used together with a geostationary orbit satellite)
www.edutalks.org
LEO Satellite coverage
www.edutalks.org
Demo of satellite coverage
www.edutalks.org
Advantages of Non-geostationary constellations
1. Since these orbits are closer to earth, the free space loss is lower and hence it is possible to use hand held terminals. Path loss at 1.5 GHz for LEO=152.87 dB, MEO=175.96 dB and GEO=187.10dB.
2. LEO and MEO reduces the propagation delay which reduces or eliminates delay related problems.
3. These orbits offer a higher frequency reuse. 3. These orbits offer a higher frequency reuse. Maximum distance between two points which view a
satellite at an elevation angle 100 isLEO (at altitude 700km) = 3885 kmMEO (10000km) = 12790 kmGEO (36000km) = 15914 km
Hence note that 4 LEO satellites would cover the same geographical distance as a single GEO. Thus LEO has 4 times frequency reuse than GEO.
www.edutalks.org
4 . Distributed architecture of LEO and MEO orbits make them more resistant to satellite failures and hence more reliable.
5. Competitions between operators has triggered a feverish technical, regulatory and financial activity in the industry.
Current non geostationary proposals are-
• MEO system -Offers a real time services. Medium/high • MEO system -Offers a real time services. Medium/high bit rates communication facility. eg: ICO system
• Little LEO -Offers a low bit rate non-real time services such 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 as internet high speed file download (bit rate=16kbps to 1 Gbps) eg: Teledesic.
www.edutalks.org
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 such as the earth) being orbited on each revolution. It therefore has an inclination of (or very close to) 900 to the equator.
• Polar orbits are often used for earth-mapping, earth observation, as well as some weather satellites
• The disadvantage to this orbit is that no one spot on the Earth's surface can be sensed continuously from a satellite Earth's surface can be sensed continuously from a satellite in a polar orbit.
• Polar satellites include: Defense meteorological satellite program (DMSP), Landsat, SPOT and NOAA. Landsat and SPOT 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
www.edutalks.org
Polar orbits
www.edutalks.org
Eg. of the positions of a sun-synchronous satellite in 12 hour intervals
Sun synchronous satellites pass over any given latitude at almost the same local time during each orbital pass
www.edutalks.org
Polar constellations
Here : Ψ = coverage circle
m = 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
www.edutalks.org
• 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 NΩ/4ПWhere NΩ = total solid angle
Ω = solid angle bounded by a single satellite=2П(1-cos Ψ)
Triple coverage:Triple coverage:
• The constellation geometry is similar to single coverage case, with at least 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
www.edutalks.org
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 coverageby deploying satellites in circular orbits of same period and 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.
www.edutalks.org
Adjacent orbital planes are separated equally around a reference plane (equatorial).Within each orbit ,neighboring satellites have equal angular separation.
www.edutalks.org
Inclined constellations
αi = right ascension angle of ith orbital plane =2Пi / Pβi = inclination angle of ith orbitγi = initial phase angle of ith satellite = m αi
m = (0 to N-1)/Q ; N = PQ (P,Q are integers)
Q = number of satellites per planewww.edutalks.org
Hybrid constellations
• This combines the various types orbits for full earth coverage
• These orbits have different orbital period
• Eg: using circular orbits for covering equatorial regions and elliptical orbits for higher altitude regions
Eg: using GEO for covering equatorial regions and Eg: using GEO for covering equatorial regions and inclined orbits for polar regions
www.edutalks.org
Regional coverage:In some cases it is necessary to cover only a part of the world. This is made possible by number of spot beams. Here it is necessary to ensure that all the satellites pass over the same service area.Eg: equatorial regions may be covered by deploying satellites in equatorial planesUsing elliptical orbits inclined at 63.40 can be used for covering high altitude because satellites in these orbits dwell over high altitudes over a considerable time.over a considerable time.
Thuraya allows to create more than 200 spot beams and handle 13,750 simultaneous phone calls. Telecommunications Services offered are: • Voice • Fax at 9.6 Kbps • Data at 9.6 Kbps
www.edutalks.org
Footprint and spot beams
www.edutalks.org
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 to end delay , td=delay in message transfer
• Eg :A single satellite in polar orbit can cover every regions of earth within a time dependent on the orbital period
ORBCOMM provides low cost, ORBCOMM provides low cost, reliable, two-way data communications services around the world through a global 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
www.edutalks.org
Random and phased constellations• Constellations can be categorized on the basis of phase relationship
between satellites with respect of each other.• In random constellation all or some constellation parameters such
as 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 simpler to maintain but are inefficient in Random constellations are simpler to maintain but are inefficient in terms of coverage property and tend to randomly crowd the celestial sphere around the chosen altitude.
www.edutalks.org
Design considerations of a non GEO satellite systems
1.Traffic distribution and coverage:
• A constellation design depends on service area and geographical distribution of traffic within that area.
• A worldwide coverage is essential for an operator interested in global operations, but regional operator is interested in only a specific region. Hence the constellations are completely different.
• Good RF visibility ensures adequate signal strength before a connection is established and the increase in spectrum reuse.connection is established and the increase in spectrum reuse.
• Complexity in coverage design is the dynamic variation in position of 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 and prevailing economic /political condition of the target market.
www.edutalks.org
2. Satellite capacity:• The capacity required per satellite increases as the orbit altitude
increases because a satellites field of view and captured trafficincreases with the altitude
• Total constellation capacity is the sum of capacities of satellites• At higher altitudes satellite capacities are better shared
• Hence as the altitude increases, total constellation capacity reduces and more efficient constellation capacity is utilized
3. State of spacecraft technology:3. State of spacecraft technology:
• Antenna size and complexity -as the altitude increases , larger antennas are required to meet link quality objective and maintain frequency reusability
• Spacecraft DC power -DC power determines the capacity of the satellite
• Inter satellite link -satellites with inter links influence the network routing scheme
www.edutalks.org
4.Terminal characteristic and communication requirement-• The size of terminals and their communication capability
influence a satellites power and sensitivity requirements• RF power of a handset is limited by radiation safety considerations,
battery size /capacity and the target terminal cost• If the satellites are brought closer , power required can be reduced
but number of satellites in the constellations increases.5.Quality of service:• Quality of service refers to RF link reliability, propagation delay
and signal quality measured as bit error rateand signal quality measured as bit error rate• Higher link reliability requires higher elevation angle• Propagation conditions improve as the elevation angle increases
because number of obstructions reduces• Propagation delay is related to the altitude of orbit• Hence for interactive applications lower orbits are best and non real
time applications are insensitive to altitude• Signal quality is related to link conditions and issues such as carrier
to noise power density/modulation/coding schemes
www.edutalks.org
6. Spectrum availability:• Frequency reusability can be increased by spatial/polarization
diversity• This is achieved by using spot and shaped beams• For a given spot beam size lower altitude constellations can give
increased reusability• Additional measures like modulation, coding and multiple access
schemes can maximize radio resource• 7. Orbital considerations:• 7. Orbital considerations:• Space environment affects the orbit selection• Atmospheric drag, eclipses, ionization8. 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.
www.edutalks.org
Assignment 01
Explain about the effect of
a) eclipse due to earth
b) eclipse due to moonb) eclipse due to moon
c) solar interference
on geo-stationary satellites.
Submit before:18.08.2008
www.edutalks.org
Problems
Q1. Find out the radius of a geostationary satellite orbit.
Given: T = 23Hr 56Min 4.1SecT = 23Hr 56Min 4.1SecG = 6.672 X10-11 m3/kg/s2
M = 5974 X1024 kgr = 6378.1414 km
take µ = √GM
www.edutalks.org
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= 42164.17 km
Altitude, h = R-r= 42164.17-6378.1414
= 35786.02 km
www.edutalks.org
Q2. A satellite orbiting in equatorial plane has a period from perigee-perigee of 12 Hrs. Given that the eccentricity is 0.002. Calculate semi-major axis.
Given:G = 6.672 X10-11 m3/kg/s2
M = 5974 X1024 kgr = 6378.1414 km
www.edutalks.org
Answer:
Eccentricity is 0.002 (0<e<1), hence orbit is elliptical.
For an elliptical orbit,
T2 = 4 П2 .(a)3
µµ
12x60x60 =2 П √ a3/6.672x10-11 x 5974x1024
a3 = 1.886x1025
a = 266183.1516 km
www.edutalks.org
Q3. Calculate the apogee and perigee heights for the given orbital parameters.
e=0.0011501 and
a= 7192.3 km
Given:
r = 6378.1414 km
www.edutalks.org
Answer:
ra=a(1+e) = 7200.57 km
rp= a(1-e) = 7184.03 km
Apogee height, ha = ra – r = 822.14 kmApogee height, ha = ra – r = 822.14 km
Perigee height, hp = rp – r = 805.89 km
www.edutalks.org
Q4. A satellite is in elliptical orbit with a perigee of 1000 km and an apogee of 4000 km. Using a mean earth radius of 6378.14 km , find the period of the orbit in hours ,minutes and seconds. and seconds.
Also find the ‘e’ of the orbit.
www.edutalks.org
Answer:
ra = ha+R =4000+6378.14=10378.14 km
rp =hp +R=1000+6378.14=7378.14 km
a= (ra + rp)T=√4П2a3/GM
=8320.94 sec = 2hr 19 mts 8 sec=8320.94 sec = 2hr 19 mts 8 sec
ra = a(1+e)
e= (ra/a) -1
= 0.169
www.edutalks.org
Communication Satellites
www.edutalks.org
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 areaiii) coverage areaiv) 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.
www.edutalks.org
First TV image of weather (1960)
www.edutalks.org
First complete view of world’s weather, photographed by TIROS 9 (13/2/1965).
Image assembled from 450 individual photographs
www.edutalks.org
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
provided for liquid flow.• The absence of gravity facilitates operation of the deployment mechanisms
used for stowing antennas and solar panels during launch.b) Atmospheric pressure and temperature :-b) Atmospheric pressure and temperature :-
• At high altitudes, atmospheric pressure is extremely low (10-7 torr).• This makes thermal conduction negligible and increase friction between
surfaces.• 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 during
sunlight and 95-120K during eclipses.www.edutalks.org
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.• While Satellites passing through Van Allen belt ,deflected charged particles • While Satellites passing through Van Allen belt ,deflected charged particles
that are trapped in this region affect electronic components• Hence special manufacturing mechanisms are used to harden the
components 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 in gradual 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.
www.edutalks.org
Life time and reliability• Lifetime of a geostationary satellite is determined by the maximum
acceptable deviation in inclination and orbital location.• Satellite is maintained in its orbital location by firing thrusters
regularly, using stored fuel • Hence the operational lifetime of a satellite is determined by-
a) Increasing fuel capacityb) Saving fuel by accepting orbital deviation to the maximum extent possible.possible.
• However there is a practical limit to a satellites fuel storage capacity. Hence satellite lifetime is between 12-15 years.
www.edutalks.org
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 component t
0dt
e− λ∫• R= where λ= failure rate of a component
• Unit of λ is specified as FIT, the number of failures in 109 Hr.
0e ∫
www.edutalks.org
• Three regions can be identified• Three regions can be identified
An early high failure rate region attributed to manufacturing faults, defects in materials etc
A region of low failure attributed to random component failuresA 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 the operational 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.
www.edutalks.org
• The reliability can be expressed as
R = e-λt = 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 :
Rs= e-(λ1+ λ2+… λn)t
• Parallel redundancy is useful when the reliability of an individual subsystem is high. 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- Qi =1- (1- R)i =1- (1- e-λt)iwww.edutalks.org
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
reliability of the communication system is obtained asreliability of the communication system is obtained asRs =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 employedR= reliability without employing redundancyM= increase in mass due to added redundancy
• The addition of redundant equipment increases the cost of the transponder
www.edutalks.org
www.edutalks.org
www.edutalks.org
Back up slides
www.edutalks.org
www.edutalks.org
Transponder
www.edutalks.org
Solar eclipses
www.edutalks.org
www.edutalks.org
----Introduction to Solar System Dynamics----
rp ra
Basic orbital elements (ellipse)
e=0: circle
e<1: ellipse
e=1: parabola
e>1: hyperbola
v
r
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
−−=
www.edutalks.org
----Introduction to Solar System Dynamics----
Useful orbital parameters (elliptical orbit)
1) Velocity:
2) Period:
−=ar
GMu12
GM
aT
3
2π=
M: mass of central body
m: mass of orbiting body
r: distance of m from M
(M>>m)
3) Energy:
4) Angular momentum:
GM
a
GMmE
2−=
)1(
,2eaMGmL
urmL
−⋅⋅⋅=
×⋅= rr
(Constant!)
(Constant!)
www.edutalks.org
www.edutalks.org
Spin stabilization
• With spin stabilization, the entire spacecraft rotates around its ownvertical axis, spinning like a top. This keeps the spacecraft's orientationin space under control.
• The spinning spacecraft resists perturbing forces.
• Designers of early satellites used spin-stabilization for their satellites,• Designers of early satellites used spin-stabilization for their satellites,which most often have a cylinder shape and rotate at one revolutionevery second.
• Spin stabilization was used for NASA's Pioneer 10 and 11 spacecraft,the Lunar Prospector, and the Galileo Jupiter orbiter.
www.edutalks.org
• 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 uselarge solar arrays to obtain power from the Sun. Thus, it requireslarge amounts of battery power.
• Another disadvantage of spin stabilization is that the instruments or • Another disadvantage of spin stabilization is that the instruments or antennas also must perform “despin” maneuvers so that antennas oroptical instruments point at their desired targets.
www.edutalks.org
Reaction wheel stabilisation
• With three-axis stabilization, satellites have small spinning wheels, called reaction wheels or momentum wheels, that rotate so as to keep the satellite in the desired orientation in relation to the Earth and the Sun.
• If satellite sensors detect that the satellite is moving away from the proper orientation, the spinning wheels speed up or slow down to return the satellite to its correct position.
• Some spacecraft may also use small propulsion-system thrusters to • Some spacecraft may also use small propulsion-system thrusters to 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
www.edutalks.org
Alignment
• There are a number of components which need alignment– Solar panels
– Antennae
• These have to point at different parts of the sky at • These have to point at different parts of the sky at different times, so the problem is not trivial
www.edutalks.org
Solar and sidereal day
www.edutalks.org
Elliptical Orbit Geometry & Nomenclature
Periapsis
Line of Apsides
ϑR
a c
V
RpApoapsis
Line of Apsides Rpb
• Line of Apsides connects Apoapsis, central body & Periapsis• Apogee~ Apoapsis; Perigee~ Periapsis (earth nomenclature)
S/C position defined by R & ϑ,ϑ is called true anomalyR = [Rp (1+e)]/[1+ e cos(ϑ)]
www.edutalks.org
ORBITAL ELEMENTSKeplerian Elements: True Anomaly
νννν
ωωωω
i
ΩΩΩΩ
ωωωω
www.edutalks.org
Prepared by:Girish K.P.Girish K.P.Assistant Professor in ECEJyothi Engineering College, Cheruthuruthy
www.edutalks.org