MBS COLLEGE OF ENGG AND TECH. MANGALYAAN AND GROUND SYS.
INTRODUCTION TO MANGALYAAN CHAPTER 1
1.1 INTRODUCTIONThe Mars Orbiter Mission (MOM) also called
Mangalyaan (Mars-craft from Sanskrit mangala, Mars and Yana, craft,
vehicle), is a spacecraft orbitingMarssince 24 September 2014. It
was launched on 5 November 2013 by theIndian Space Research
Organization(ISRO) under the guidance of the Project Director
Mylswamy Annadurai. The mission is a "technology demonstrator
project to develop the technologies for design, planning,
management, and operations of an interplanetary mission.It carries
five instruments that will help advance knowledge about Mars to
achieve its secondary, scientific, objective. The Mars Orbiter
Mission probe lifted-off from theFirst Launch PadatSatish Dhawan
Space Centre(Sriharikota Range SHAR),Andhra Pradesh, using aPolar
Satellite Launch Vehicle(PSLV) rocket C25 at 09:08 UTC (14:38 IST)
on 5 November 2013.Thelaunch windowwas approximately 20 days long
and started on 28 October 2013.The MOM probe spent about a month
ingeocentric,low-Earth orbit, where it made a series of seven
altitude-raisingorbital maneuversbeforetrans-Mars injectionon 30
November 2013 (UTC). After a 298-day transit to Mars, it was
successfully inserted into Mars orbit on 24 September 2014.It
isIndia's first interplanetary missionandISROhas become the
fourthspace agencyto reach Mars, after the Soviet space
program,NASA, and theEuropean Space Agency.It is also the first
nation to reachMars orbit on its first attempt, and the first Asian
nation to do so. The spacecraft is currently being monitored from
the Spacecraft Control Centre atISRO Telemetry, Tracking and
Command Network(ISTRAC) inBangalorewith support fromIndian Deep
Space Network(IDSN) antennae at Byalalu.
Fig 1.1 Artists Rendering of MOM orbiting MARS1.2HISTORYThe MOM
mission concept began with a feasibility study in 2010, after the
launch of lunar satelliteChandrayaan-1in 2008. Thegovernment of
Indiaapproved the project on 3 August 2012,after theIndian Space
Research Organizationcompleted125crore(US$20million) of required
studies for the orbiter. The total project cost may
be454crore(US$74million).The satellite costs153crore(US$25million)
and the rest of the budget has been attributed to ground stations
and relay upgrades that will be used for other ISRO projects. The
space agency had planned the launch on 28 October 2013 but was
postponed to 5 November 2013 following the delay in ISRO's
spacecraft tracking ships to take up pre-determined positions due
to poor weather in thePacific Ocean.Launch opportunities for a
fuel-savingHohmann transfer orbitoccur every 26 months, in this
case, 2016 and 2018.The Mars Orbiter's on-orbit mission life is
six-to-ten months.Assembly of the PSLV-XL launch vehicle,
designated C25, started on 5 August 2013.The mounting of the five
scientific instruments was completed atISRO Satellite
Centre,Bangalore, and the finished spacecraft was shipped to
Sriharikota on 2 October 2013 for integration to the PSLV-XL launch
vehicle.The satellite's development was fast-tracked and completed
in a record 15 months.Despite theUS federal government shutdown,
NASA reaffirmed on 5 October 2013 it would provide communications
and navigation support to the mission.During a meeting in 30
September 2014, NASA and ISRO officials signed an agreement to
establish a pathway for future joint missions to explore Mars. One
of the working group's objectives will be to explore potential
coordinated observations and science analysis betweenMAVENorbiter
and MOM, as well as other current and future Mars missions.
TheISROplans to send a follow-up mission with a greater
scientificpayloadto Mars in the 20172020 timeframe; it would
include an orbiter and a stationary lander. 1.3 COSTThe total cost
of the mission was approximately450Crore(US$73 million),making it
the least-expensive Mars mission to date.The low cost of the
mission was ascribed by Kopillil Radhakrishnan, the chairman of
ISRO, to various factors, including a "modular approach", a small
number of ground tests and long (18-20 hour) working days for
scientists.BBC's Jonathan Amos mentioned lower worker costs,
home-grown technologies, simpler design, and significantly less
complicated payload than NASA'sMAVEN.An opinion piece inThe
Hindupointed out that the cost was equivalent to less than a single
bus ride for each of India's population of 1.2 billion.
1.4 OBJECTIVESThe primary objective of the Mars Orbiter Mission
is to showcase India's rocket launch systems, spacecraft-building
and operations capabilities.Specifically, the primary objective is
to develop the technologies required for design, planning,
management and operations of aninterplanetary mission, comprising
the following major tasks: design and realization of a Mars orbiter
with a capability to perform Earth-bound maneuvers, cruise phase of
300 days, Mars orbit insertion / capture, and on-orbit phase around
Mars; deep-space communication, navigation, mission planning and
management; Incorporate autonomous features to handle contingency
situations.The secondary objective is to explore Mars Surface
features, morphology, mineralogy and Martian atmosphere using
indigenous scientific instruments.1.5 SPAECRAFT SPECIFICATION
Mass:The lift-off mass was 1,350kg (2,980lb), including 852kg
(1,878lb) of propellant. Bus:The spacecraft'sbusis a modifiedI-1
Kstructure and propulsion hardware configuration, similar
toChandrayaan 1, India's lunar orbiter that operated from 2008 to
2009, with specific improvements and upgrades needed for a Mars
mission.The satellite structure is constructed of aluminium and
composite fibre reinforced plastic (CFRP) sandwich construction.
Power:Electric power is generated by threesolar arraypanels of 1.8m
1.4m (5ft 11in 4ft 7in) each (7.56m2(81.4sqft) total), for a
maximum of 840 watts of power generation in Mars orbit. Electricity
is stored in a 36AhLi-ion battery. Propulsion:A liquid fuel engine
with a thrust of 440Newtonis used for orbit raising and insertion
into Mars orbit. The orbiter also has eight 22-newton thrusters
forattitude control.Its propellant mass is 852kg.
1.6 PAYLOADSThe 15kg (33lb) scientific payload consists of five
instruments: Mars Orbiter Mission carries five scientific payloads
to observe Martian surface, atmosphere and exosphere extending up
to 80,000 km for a detailed understanding of the evolution of that
planet, especially the related geologic and the possible biogenic
processes on that interesting planet. These payloads consist of a
camera, two spectrometers, a radiometer and a photometer. Together,
they have a weight of about 15 kg.
PayloadPrimary ObjectiveWeight (Kg)
Mars Colour Camera (MCC)Optical imaging1.27
Thermal Infrared Imaging Spectrometer(TIS)Map surface
composition and mineralogy3.2
Methane Sensor for Mars (MSM)Detection of Methane
presence2.94
Mars Enospheric Neutral Composition Analyzer (MENCA)Study of the
neutral composition of Martian upper atmosphere3.56
Lyman Alpha Photometer (LAP)Study of Escape processes of Martian
upper atmosphere through Deuterium/Hydrogen1.97
Table 1.1: Different types of payload
Figure 1.6: Design of MOM Spacecraft showing payloads at their
respective mounting locations
1.7 TELEMETRY AND COMMANDTheIndian Space Research Organization
Telemetry, Tracking and Command Networkperformed navigation and
tracking operations for the launch with ground stations
atSriharikota,Port Blair,BruneiandBiakinIndonesia,and after the
spacecraft'sapogeebecame more than 100,000km, an 18-metre (59ft)
and an 32m (105ft) diameter antenna of theIndian Deep Space
Networkwere utilized.The 18-metre (59ft) dish-antenna was used for
communication with the craft until April 2014, after which the
larger 32m (105ft) antenna was used.NASA's Deep Space Networkis
providing position data through its three stations located
inCanberra, MadridandGoldstoneon the US West Coast during the
non-visible period of ISRO's network. TheSouth African National
Space Agency's (SANSA) Hartebeesthoek(HBK) ground station is also
providing satellite tracking, telemetry and command services.
1.8 COMMUNICATIONCommunications are handled by two
230-wattTWTAsand twocoherent transponders. The antenna array
consists of alow-gain antenna, a medium-gain antenna and ahigh-gain
antenna. The high-gain antenna system is based on a single
2.2-metre (7ft 3in) reflector illuminated by a feed atS-band. It is
used to transmit and receive the telemetry, tracking, commanding
and data to and from theIndian Deep Space Network.
PhaseDateEventDetailResult
Geocentric phase5 November 2013 09:08 UTCLaunchBurn time: 15:35
min in 5 stagesApogee: 23,550km
6 November 2013 19:47 UTCOrbit raising maneuversBurn time: 416
secApogee: 23,550km to 28,825km
7 November 2013 20:48 UTCOrbit raising maneuverBurn time: 570.6
secApogee: 28,825km to 40,186km
8 November 2013 20:40 UTCOrbit raising maneuversBurn time: 707
secApogee: 40,186km to 71,636km
10 November 2013 20:36 UTCOrbit raising maneuverIncomplete
burnApogee: 71,636km to 78,276km
11 November 2013 23:33 UTCOrbit raising
maneuvers(supplementary)Burn time: 303.8 secApogee: 78,276km to
118,642km
15 November 2013 19:57 UTCOrbit raising maneuverBurn time: 243.5
secApogee: 118,642km to 192,874km
30 November 2013, 19:19 UTCTrans-Mars injectionBurn time:
1328.89 secSuccessfulheliocentric insertion
Heliocentric phaseDecember 2013 September 2014En routeto Mars
The probe travelled a distance of 780,000,000 kilometers
(480,000,000mi) in a parabolic trajectory around the Sun to reach
Mars.This phase plan included up to four trajectory corrections if
needed.
11 December 2013 01:00 UTC1st Trajectory correctionBurn time:
40.5 secSuccess
9 April 20142nd Trajectory correction (planned)Not
requiredRescheduled for 11 June 2014
11 June 2014 11:00 UTC2nd Trajectory correctionBurn time: 16
secSuccess
August 20143rd Trajectory correction (planned)Not required
22 September 20143rd Trajectory correctionBurn time: 4
secSuccess
Aero centic phase24 September 2014Mars orbit insertionBurn time:
1388.67 secSuccess
Table 1.2 Different phases of a satellite mission1.9 LAUNCHAs
originally conceived, ISRO would have launched MOM on
itsGeosynchronous Satellite Launch Vehicle(GSLV),but as the GSLV
failed twice in 2010 and ISRO was continuing to sort out issues
with itscryogenic engine,it was not advisable to wait for the new
batch of rockets as that would have delayed the MOM project for at
least three years.ISRO opted to switch to the less-powerfulPolar
Satellite Launch Vehicle(PSLV). There is no way to launch on a
direct-to-Mars trajectory with the PSLV as it does not have the
thrust required. Instead, ISRO would first launch it into Earth
orbit and slowly boost toward an interplanetary trajectory using
multiple perigee burns to maximize theOberth effect. On 19 October
2013, ISRO chairmanK. Radhakrishnanannounced that the launch had to
be postponed by a week as a result of a delay of a crucial
telemetry ship reachingFiji. The launch was rescheduled for 5
November 2013.ISRO's PSLV-XL placed the satellite into Earth orbit
at 09:50 UTC on 5 November 2013,with a perigee of 264.1km
(164.1mi), an apogee of 23,903.6km (14,853.0mi), and inclination of
19.20 degrees,with both the antenna and all three sections of the
solar panel arrays deployed.During the first three orbit raising
operations, ISRO progressively tested the spacecraft systems. The
orbiter's dry mass is 500kg (1,100lb), and it carries 852kg
(1,878lb) of fuel and oxidizer. Its main engine, which is a
derivative of the system used on India's communications satellites,
uses the bipropellant combinationmonomethyl hydrazineanddinitrogen
tetroxideto achieve the thrust necessary forescape velocityfrom
Earth. It was also used to slow down the probe for Mars orbit
insertion and, subsequently, for orbit corrections.1.10 OBJECT
RAISING MANOEUVRESSeveral orbit raising operations were conducted
from theSpacecraft Control Centre(SCC) at ISRO Telemetry, Tracking
and Command Network (ISTRAC) at Peenya, Bangalore on 6, 7, 8, 10,
12 and 16 November by using the spacecraft's on-board propulsion
system and a series of perigee burns. The aim was to gradually
build up the necessaryescape velocity(11.2km/s) to break free from
Earth's gravitational pull while minimizing propellant use. The
first three of the five planned orbit raising maneuvers were
completed with nominal results, while the fourth was only partially
successful. However, a subsequent supplementary maneuvers raised
the orbit to the intended altitude aimed for in the original fourth
maneuver. A total of six burns were completed while the spacecraft
remained in Earth orbit, with a seventh burn conducted on 30
November to insert MOM into a heliocentricorbitfor its transit to
Mars.The first orbit-raising maneuver was performed on 6 November
2013 at 19:47 UTC when the 440 newtons (99lbf)liquidengine of the
spacecraft was fired for 416 seconds. With this engine firing, the
spacecraft'sapogee was raised to 28,825km, with aperigeeof 252km.
The second orbit raising maneuver was performed on 7 November 2013
at 20:48 UTC, with a burn time of 570.6 seconds resulting in an
apogee of 40,186km.The third orbit raising manoeuvre was performed
on 8 November 2013 at 20:40 UTC, with a burn time of 707 seconds
resulting in an apogee of 71,636km. The fourth orbit raising
maneuvers, starting at 20:36 UTC on 10 November 2013, imparted an
incrementalvelocityof 35m/s to the spacecraft instead of the
planned 135m/s as a result of under burn by the motor.Because of
this, the apogee was boosted to 78,276km instead of the planned
100,000km.When testing the redundancies built-in for the propulsion
system, the flow to the liquid engine stopped, with consequent
reduction in incremental velocity. During the fourth orbit burn,
the primary and redundant coils of the solenoid flow control valve
of 440 newton liquid engine and logic for thrust augmentation by
the attitude control thrusters were being tested. When both primary
and redundant coils were energized together during the planned
modes, the flow to the liquid engine stopped. Operating both the
coils simultaneously is not possible for future operations, however
they could be operated independently of each other, in sequence.As
a result of the fourth planned burn coming up short, an additional
unscheduled burn was performed on 12 November 2013 that increased
the apogee to 118,642km,a slightly higher altitude than originally
intended in the fourth maneuver.The apogee was raised to 192,874km
on 15 November 2013, 19:57 UTC in the final orbit raising
maneuver.
Figure 1.8: Orbit Trajectory Diagram (not to scale)
1.11 TRANS MARS INJECTIONOn 30 November 2013 at 19:19 UTC, a
23-minute engine firing initiated thetransferof MOM away from Earth
orbit and onheliocentrictrajectory toward Mars.The probe travelled
a distance of 780,000,000 kilometers (480,000,000mi) to reach
Mars.
1.12 TRAJECTORY CORRECTION MANEUVERSFour trajectory corrections
were originally planned, but only three were carried out.The first
trajectory correction maneuver (TCM) was carried out on 11 December
2013, 01:00 UTC, by firing the 22 newtons (4.9lbf) thrusters for a
duration of 40.5 seconds.As observed in April 2014, MOM is
following the designed trajectory so closely that the trajectory
correction maneuver planned in April 2014 was not required. The
second trajectory correction maneuver was performed on 11 June
2014, at 16:30 hrs IST by firing the spacecraft's 22 newton
thrusters for a duration of 16 seconds.The third planned trajectory
correction maneuver was postponed, due to the orbiter's trajectory
closely matching the planned trajectory.The third trajectory
correction was also a deceleration test 3.9 seconds long on 22
September 2014.
1.13 MARS ORBIT INSERTIONThe plan was for an insertion intoMars
orbiton 24 September 2014,approximately 2 days after the arrival of
NASA'sMAVENorbiter.The 440N liquid apogee motor was successfully
test fired at 09:00 UTC (14:30 IST) on 22 September for 3.968
seconds, about 41 hours before actual orbit insertion.On 24
September 2014, at IST 04:17:32 satellite communication changed
over to the medium gain antenna. At IST 06:56:32 forward rotation
started and locked the position to fire, at IST 07:14:32 anattitude
controlmaneuver took place with the help of thrusters after eclipse
started at IST 07:12:19 and LAM (Liquid Apogee Motor) started
burning at IST 07:17:32 and ended at IST 07:41:46. After that
reverse maneuver took place, the spacecraft successfully entered
Martian orbit
Fig 1.9: Simulated view of MARS Orbiter along with Mars, Earth,
Mercury and sun on 3rd October 2014 at 17:00 UTC. The MARS Orbiter
Mission satellite is an altitude of about 1300 miles from Mars
ISRO (INDIAN SPACE RESEARCH ORGANISATION) CHAPTER 2 2.1
INTRODUCTIONTheIndian Space Research Organisation(ISRO,/sro/;Hindi:
Bhratya Antarikha Anusandhn Sangahan) is the primaryspace
agencyofIndia. ISRO is among the largestgovernment space agencies
in the world. Its primary objective is to advancespace
technologyand use its applications for national benefit.Established
in 1969, ISRO superseded the erstwhileIndian National Committee for
Space Research (INCOSPAR). Headquartered inBangalore, ISRO is under
the administrative control of theDepartment of Spaceof
theGovernment of India.ISRO built India's firstsatellite,Aryabhata,
which was launched by theSoviet Unionon 19 April in 1975. In
1980,Rohini became the first satellite to be placed in orbit by an
Indian-made launch vehicle,SLV-3. ISRO subsequently developed two
other rockets: thePolar Satellite Launch Vehicle (PSLV)for
launching satellites intopolar orbitsand the Geosynchronous
Satellite Launch Vehicle (GSLV)for placing satellites
intogeostationary orbits. These rockets have launched
numerouscommunications satellitesandearth observation satellites.
Satellite navigation systems likeGAGAN andIRNSShave been deployed.
In January 2014, ISRO successfully used anindigenous cryogenic
enginein a GSLV-D5 launch of the GSAT-14.On 22 October 2008, ISRO
sent its first mission to theMoon,Chandrayaan-1. On 5 November
2013, ISRO launched its Mars Orbiter Mission, which successfully
entered theMarsorbit on 24 September 2014, making India the first
nation to succeed on its maiden attempt, and ISRO thefirst Asian
space agencyto reach Mars orbit.[6]Future plans include development
ofGSLV Mk III(for launch of heavier satellites), development of
areusable launch vehicle,human spaceflight,further lunar
exploration, interplanetary probes,a satellite to study the Sun,
etc.Over the years, ISRO has also conducted a variety of operations
for both Indian and foreign clients. ISRO has several field
installations as assets, and cooperates with the international
community as a part of several bilateral and multilateral
agreements. In June 2014, it launched five foreign satellites by
the PSLV. There are plans for the development and launch of a
satellite which will be collectively used by the
eightSAARCnations.
2.2 LAUNCH VEHICLE FLEET During the 1960s and 1970s, India
initiated its own launch vehicle programme owing to geopolitical
and economic considerations. In the 1960s1970s, the country
successfully developed a sounding rockets programme, and by the
1980s, research had yielded the Satellite Launch Vehicle-3 and the
more advanced Augmented Satellite Launch Vehicle (ASLV), complete
with operational supporting infrastructure.ISRO further applied its
energies to the advancement of launch vehicle technology resulting
in the creation of PSLV and GSLV technologies.2.3 SATELLITE LAUNCH
VEHICLE (SLV)The Satellite Launch Vehicle, usually known by its
abbreviation SLV or SLV-3 was a 4-stage solid-propellant light
launcher. It was intended to reach a height of 500km and carry a
payload of 40kg.[18]Its first launch took place in 1979 with 2 more
in each subsequent year, and the final launch in 1983. Only two of
its four test flights were successful.
2.3.1 AUGMENTED SATELLITE LAUNCH VEHICLE (ASLV)The Polar
Satellite Launch Vehicle, usually known by its abbreviation PSLV,
is anexpendable launch systemdeveloped to allow India to launch its
Indian Remote Sensing (IRS) satellites intoSun synchronous orbits,
a service that was, until the advent of the PSLV, commercially
viable only from Russia. PSLV can also launch small satellites
intogeostationary transfer orbit(GTO). The reliability and
versatility of the PSLV is proven by the fact that it has launched
70 satellites / spacecraft ( 30 Indian and 40 Foreign Satellites)
into a variety of orbits so far.In April 2008, it successfully
launched 10 satellites at once, breaking a world record held by
Russia.On 30 June 2014, the PSLV flew its 25th consecutive
successful launch mission,delivering a payload of five foreign
satellites into orbit. Its only failure in 26 flights was its
maiden voyage in September 1993, providing the rocket with a 96
percent success rate.
2.3.2 GEOSYNCHRONOUS SATELLITE LAUNCH VEHICLE (GSLV)The
Geosynchronous Satellite Launch Vehicle, usually known by its
abbreviation GSLV, is an expendable launch system developed to
enable India to launch itsINSAT-type satellites into geostationary
orbit and to make India less dependent on foreign rockets. At
present, it is ISRO's heaviest satellite launch vehicle and is
capable of putting a total payload of up to 5 tons to Low Earth
Orbit. The vehicle is built by India with the cryogenic engine
purchased from Russia while the ISRO develops its own engine
programme.In a setback for ISRO, the attempt to launch the GSLV,
GSLV-F07 carrying GSAT-5P, failed on 25 December 2010. The initial
evaluation implies that loss of control for the strap-on boosters
caused the rocket to veer from its intended flight path, forcing a
programmed detonation. Sixty-four seconds into the first stage of
flight, the rocket began to break up due to the acute angle of
attack. The body housing the 3rd stage, the cryogenic stage,
incurred structural damage, forcing the range safety team to
initiate a programmed detonation of the rocket.On 5 January 2014,
GSLV-D5 successfully launched GSAT-14 into intended orbit. This
also marked first successful flight using indigenous cryogenic
engine, making India sixth country in the world to have this
technology.2.3.3 GEOSYNCHRONOUS SATELLITE LAUNCH VEHICLE
MARK-IIIThe Geosynchronous Satellite Launch Vehicle Mark-III is a
launch vehicle currently under development by the Indian Space
Research Organization. It is intended to launch heavy satellites
intogeostationary orbit, and will allow India to become less
dependent on foreign rockets for heavy lifting. The rocket, though
the technological successor to theGSLV, however is not derived from
its predecessor.A GSLV III is planned to launch on a suborbital
test flight in the third quarter of 2014/15. This suborbital test
flight will demonstrate the performance of the GSLV Mk.3 in the
atmosphere. This launch has been delayed from May, June, July and
August of 2014.
2.4 EARTH OBSERVATION AND SATELLITEIndia's first satellite,
theAryabhata, was launched by theSoviet Unionon 19 April 1975
fromKapustin Yarusing aCosmos-3Mlaunch vehicle. This was followed
by the Rohini series of experimental satellites which were built
and launched indigenously. At present, ISRO operates a large number
of earth observation satellites.
2.4.1 THE INSAT SERIESINSAT (Indian National Satellite System)
is a series of multipurpose geostationary satellites launched by
ISRO to satisfy the telecommunications, broadcasting, meteorology
and search-and-rescue needs of India. Commissioned in 1983, INSAT
is the largest domestic communication system in the Asia-Pacific
Region. It is a joint venture of the Department of Space,
Department of Telecommunications,India Meteorological
Department,All India RadioandDoordarshan. The overall coordination
and management of INSAT system rests with the Secretary-level INSAT
Coordination Committee
2.4.2 THE IRS SERIESIndian Remote Sensing satellites (IRS) are a
series of earth observation satellites, built, launched and
maintained by ISRO. The IRS series provides remote sensing services
to the country. The Indian Remote Sensing Satellite system is the
largest constellation of remote sensing satellites for civilian use
in operation today in the world. All the satellites are placed in
polarSun-synchronous orbitand provide data in a variety of spatial,
spectral and temporal resolutions to enable several programmes to
be undertaken relevant to national development. The initial
versions are composed of the 1 (A,B,C,D) nomenclature. The later
versions are named based on their area of application including
OceanSat, CartoSat, Resource
2.4.3 RADAR IMAGING SATELLITESISRO currently operates twoRadar
Imaging Satellites.RISAT-1was launched from Sriharikota Spaceport
on 26 April 2012 on board a PSLV.RISAT-1 carries a C-band Synthetic
Aperture Radar (SAR) payload, operating in a multi-polarisation and
multi-resolution mode and can provide images with coarse, fine and
high spatial resolutions.India also operatesRISAT-2which was
launched in 2009 and acquired from Israel at a cost $110
million
1.4.4 OTHER SATELLITESISRO has also launched a set of
experimental geostationary satellites known as
theGSATseries.Kalpana-1, ISRO's first dedicated meteorological
satellite,was launched by thePolar Satellite Launch Vehicleon 12
September 2002.[33]The satellite was originally known as
MetSat-1.In February 2003 it was renamed to Kalpana-1 by the Indian
Prime MinisterAtal Bihari Vajpayeein memory ofKalpana Chawla a NASA
astronaut of Indian origin who perished inSpace Shuttle
Columbia.
Figure 2.1 Saral satellite model
ISRO has also successfully launched the Indo-French
satelliteSARALon 25 February 2013, 12:31 UTC. SARAL (or "Satellite
with ARgos and ALtiKa") is a cooperative altimetry technology
mission. It is being used for monitoring the oceans surface and
sea-levels. AltiKa will measure ocean surface topography with an
accuracy of 8mm, against 2.5cm on average using current-generation
altimeters, and with a spatial resolution of 2km.In June 2014, ISRO
launched French Earth Observation Satellite SPOT-7 (mass 714kg)
along withSingapore's first nano satellite VELOX-I,Canada's
satellite CAN-X5,Germany's satellite AISAT, via the PSLV-C23 launch
vehicle. It was ISRO's 4th commercial launch
ATMOSPHERE OF MARS CHAPTER 3
3.1 INTRODUCTIONTheatmosphere ofMarsis, like that ofVenus,
composed mostly ofcarbon dioxidethough far thinner. There has been
renewed interest in its composition since the detectionof traces
ofmethanethat may indicatelifebut may also be produced by
ageochemicalprocess,volcanicorhydrothermal activity.
Figure 3.1: MARS this atmosphere, visible on the horizon in this
low-orbit imageTheatmospheric pressureon the Martian surface
averages 600pascals(0.087psi), about 0.6% of Earth's mean sea level
pressure of 101.3 kilopascals (14.69psi) and only 0.0065% that
ofVenus's9.2 mega pascals (1,330psi). It ranges from a low of 30
pascals (0.0044psi) onOlympus Mons's peak to over 1,155 pascals
(0.1675psi) in the depths ofHellas Planitia. This pressure is well
below theArmstrong limitfor the unprotected human body. Mars's
atmospheric mass of 25teratonnescompares to Earth's 5148 tera
tonnes with ascale heightof about 11 kilometers (6.8mi) versus
Earth's 7 kilometers (4.3mi).The Martian atmosphere consists of
approximately 96%carbon dioxide, 2.1%argon, 1.9%nitrogen, and
traces of freeoxygen,carbon monoxide,waterandmethane, among other
gases,for a meanmolar massof 43.34 g/mol.The atmosphere is quite
dusty, giving the Martian sky a light brown or orange-red color
when seen from the surface; data from theMars Exploration
Roversindicate that suspended dust particles within the atmosphere
are roughly 1.5micro-metersacross.
3.2 STRUCTUREPressurecomparison
WherePressure
Olympus Monssummit0.03kilopascals(0.0044psi)
Mars average0.6 kilopascals (0.087psi)
Hellas Planitiabottom1.16 kilopascals (0.168psi)
Armstrong limit6.25 kilopascals (0.906psi)
Mount Everest summit33.7 kilopascals (4.89psi)
Earth sea level101.3 kilopascals (14.69psi)
Mars's atmosphere is composed of the following layers:Lower
atmosphere: A warm region affected by heat from airbornedustand
from the ground.Middle atmosphere: The region in which Mars'sjet
streamflowsUpper atmosphere, or thermosphere: A region with very
high temperatures, caused by heating from the Sun. Atmospheric
gases start to separate from each other at these altitudes, rather
than forming the even mix found in the lower atmospheric
layers.Exosphere: Typically stated to start at 200km (120mi) and
higher, this region is where the last wisps of atmosphere merge
into the vacuum of space. There is no distinct boundary where the
atmosphere ends; it just tapers away. There is also a complicated
ionosphere,and a seasonal ozone layer over the south pole.
Observations and measurement from EarthIn 1864,William Rutter
Dawesobserved "that the ruddy tint of the planet does not arise
from any peculiarity of its atmosphere seems to be fully proved by
the fact that the redness is always deepest near the centre, where
the atmosphere is thinnest."Spectroscopic observations in the 1860s
and 1870sled many to think the atmosphere of Mars is similar to
Earth's. In 1894, though,spectral analysisand other qualitative
observations byWilliam Wallace Campbellsuggested Mars resembles
theMoon, which has no appreciable atmosphere, in many respects. In
1926, photographic observations byWilliam Hammond Wrightat theLick
ObservatoryallowedDonald Howard Menzelto discover quantitative
evidence of Mars's atmosphere.
3.2 COMPOSITIONThe composition of the abundant gases which are
present on the mars are shown in the figure 3.2
Figure 3.2: Planet MARS most abundant gases
PAYLOAD CHAPTER 4 4.1 CLASSIFICATION OF SCIENTIFIC PAYLOADThe
15kg (33lb) scientific payload consists of five instruments:
Atmospheric studies: Lyman-Alpha Photometer (LAP) aphotometerthat
measures the relative abundance
ofdeuteriumandhydrogenfromLyman-alpha emissionsin the upper
atmosphere. Measuring the deuterium/hydrogen ratio will allow an
estimation of the amount of water loss toouter space. Methane
Sensor for Mars (MSM) will measuremethane in the atmosphere of
Mars, if any, and map its sources. Particle environment studies:
Mars Exospheric Neutral Composition Analyser (MENCA) is aquadrupole
mass analysercapable of analysing the neutral composition of
particles in the exosphere.Surface imaging studies: Thermal
Infrared Imaging Spectrometer (TIS) will measure the temperature
and emissivity of the Martian surface, allowing for the mapping of
surface composition and mineralogy of Mars. Mars Colour Camera
(MCC) will provide images in the visual spectrum, providing context
for the other instruments
4.2 EXPLANATION OF VARIOUS INSTRUMENTS IN MARS ORBITER:4.2.1
MARS COLOUR CAMERA (MCC)Mangalyaan carries a camera payload that
acquires color images of planet Mars. MCC covers a spectral range
of 400 to 700 nanometers the visible spectrum. This tri-color Mars
color camera gives images & information about the surface
features and composition of Martian surface. They are useful to
monitor the dynamic events and weather of Mars. MCC will also be
used for probing the two satellites of Mars-Phobos & Deimos. It
also provides the context information for other science
payloads
Figure 4.1: Mars color camera on-board Mangalyaan
4.2.1.1 COMPONENTS OF MCC Multi element lens assembly Pixel
array detector with RBG Bayer filter4.2.1.2 MULTI-ELEMENT LENS
Multi element lenses are used when a singlet lens cannot fulfill
the needed optical function due to aberration or wave front
distortion, or when more complex optical transformation is
required
Figure 4.2: Multi element lens in flow chart diagram
4.2.1.3 PIXEL ARRAY DETECTOR WITH BAYER FILTER
The PAD detector is a 2-dimensional imager capable of storing
subsequent frames in less than 0.5 microsecond. It will be used for
time resolved experiments where speed is a critical factor. Figure
4.3.1: Color filter array 3D viewABayer filtermosaic is acolor
filter array(CFA) for arrangingRGBcolor filters on a square grid of
photo sensors. Its particular arrangement of color filters is used
in most single-chip digitalimage sensorsused in digital cameras,
camcorders, and scanners to create a color image. The filter
pattern is 50% green, 25% red and 25% blue.
Figure 4.3.2: Working of CFA
4.2.1.4 SCIENTIFIC OBJECTIVES
To image the surface feature of Mars (mountains, valleys,
sedimentary features, various volcanic features). The geological
setting of the area of interest around methane source would be
mapped. Expected results from the MOM sensor would be co-analyzed
with MCC for determining the nature of source. To study Martian
polar ice caps and its seasonal variations. Mapping dynamic events
like dust storms and dust devils. To image the natural satellite of
Mars (Phobos) and other asteroids encountering the orbit.MOM has
uniqueness in terms of its highly elliptical orbit. Earth orbit
imaging experiments using MCChas yielded good quality images and it
is expected that MCC will return very good quality images from Mars
as well.
On November 19, 2013, from a 70,000 kilometers above Earth, the
Mars Orbiter Mission took this photo of the Indian
subcontinent.
Figure 4.4: Indias First mars Mission
4.2.2 METHANE SENSOR FOR MARSMethane is an organic molecule
present in gaseous form in the Earths atmosphere. More than 90% of
Methane on our home planet is produced by living organisms. The
recent detection of plumes of Methane in the northern hemisphere of
Mars is of great interest because of its potential biological
origin.
Figure 4.5: Methane sensor for MARSMethane sensor for Mars is
one of the scientific instruments of the payload on MOM spacecraft,
MSM payload weighing 2.94 kg is designed to measure amount of
Methane of the order of parts per billion (ppbs) in martian
atmosphere. MSM is a differential radiometer (radiometer is a
device used to measure temperature of cosmic background) based on
Fabry Perot Etalon (FPE) filters. MSM maps the source and sinks of
Methane by scanning the full Martian disc from apogee position of
Mars Orbiter.4.2.2.1 DIFFERENTIAL MICROWAVE RADIOMETER Figure 4.6:
Differential microwave radiometer4.2.2.2 SCIENTIFIC
INVESTIGATIONSBy correlating the temporal and spatial variation of
methane with other geophysical parameters, it may be possible to
find out more about the processes, biotic or abiotic which
determine the dynamics of Methane cycle within the Martian
atmosphere and ultimately solve some of the interesting things
about the existence of life forms in Mars.
Figure 4.7: Methane variation in MARS
4.2.2.3SENSORCONFIGURATIONFabry-perot Etalon sensor consists of
two channels - Methane channel, reference channel. Fore-optics
collects radiance from the sense and focuses it onto a field-Stop.
Diverging beam from the field stop is collimated and then divided
into two parts by a beam filter. One part of the beam transmits
through FPE filter of methane channel whereas the other part
transmits through FPE filter of reference channel and then focused
onto respective focal planes. In GaAS photo divider are used as
photo detectors. In GaAs or indium gallium arsenide is an alloy of
gallium arsenide and indium arsenide. As gallium and indium belong
to Group III of the Periodic Table, and arsenic and phosphorous
belong to Group V, these binary materials and their alloys are all
III-V compound semiconductors (In GaAS Photo detectors are
sensitive to wavelength over a wide spectral range and are
available as image sensors, and has applications in optoelectronic
technology.)
Figure 4.8: Geological maps of MARS
4.2.2.3.1 FABRY-PEROT ETALON SENSOR OPTICAL CONFIGURATIONAn FPE
filter transmit optical radiation at regular intervals of
frequency. FPE filter used in methane channel and reference
channels are exactly similar. But FPE filter of reference channel
is tilted by about 1 degree with respect to the optical axis so
that its transmission peaks are slightly shifted. Transmission
bands of first Etalon exactly coincide with the absorption lines of
methane where as transmission peaks of reference Etalon are
positioned in between the gaseous absorption lines where absorption
is nil.
Figure 4.8.1: Functioning of fabry-perot etalon sensor
Figure 4.8.2: Working of FPE filter
TECHNIQUE USED TO DETERMINE CONCENTRATION OF METHANE:Radiance
measured in methane channel varies with Methane concentration in
the atmosphere where as that of reference level is insensitive to
it. So, the differential signal gives a Measure of methane in the
atmosphere. Based on this technique, Methane concentration on
Martian atmosphere is determined.
4.2.2.4:IN GAAS PHOTO DETECTOR
Figure 4.9: In GAAS Photo detector4.2.2.5CONCLUSIONThe previous
rover missions to Mars reported that the Red Planets atmosphere
contained Methane and that its concentrations depend on seasonal
fluctuations. NASAs rover has come up empty-handed in its search
for Methane in the atmosphere of Mars, during 8 months of data
collection, the rover detected average Methane concentrations of
0.18 parts per billion. The researches say that, because of the
measurements margin of error, the finding translates to essentially
no methane in Martian atmosphere.Let us hope for the success of
Mangalyaan , MSM, through which we can ultimately determine the
dynamics of Methane cycle within the Martian atmosphere and
ultimately solve some of the interesting things about the existence
of life forms in Mars.4.2.3 Lyman Alpha PhotometerLyman Alpha
Photometer (LAP) is one of the scientific instruments of the
payload on MOM spacecraft, which is Indias maiden mission to the
red planet, Mars. Figure 4.10: Lyman alpha photometer
Why is it called Lyman Alpha Photometer?When electron in a
hydrogen atom makes transition from n=2 energy level to n=1 energy
level, a photon is released and this type of emission of photon is
known as Lyman Alpha emission. Photometer is an instrument for
measuring intensity of light. Lyman Alpha Photometer is an
absorption cell photometer.4.2.3.1 LYMAN ALPHA EMISSION
Figure 4.11: Lyman Alpha emissionWhat is an absorption cell
photometer?An absorption photometer for measuring the absorption by
conducting the light to a thin flow cell in which a liquid sample
flows, wherein the sample light for measuring the absorption peak
is superimposed on the reference light selected from the
transparent(window) range of the liquid and the absorbance is
detected by separating the sample light and reference light after
transmission of the flow cell changes in the light path conditions
can be mentioned accurately and therefore high accuracy measurement
immune to noises is made possible even using an elongated flow
cell.
4.2.3.1 ABSORPTION CELL PHOTOMETERLAP measures the relative
abundance of deuterium and hydrogen from Lyman-alpha emission in
the Martian upper atmosphere .Measurement of D/H (Deuterium to
Hydrogen abundance ratio) will improve our understanding of the
process involved in the loss of water from the planet. The
estimated D/H ratio will be used in MG CM (Mars General Circulation
Model) algorithms to the present Water escape rate from Martian
Exosphere.
Figure 4.12: Absorption cell photometer a) atomic absorption
meter b) mass spectrometer.4.2.3.2 FUNDAMENTALS AND PRINCIPLE OF
WORKING OF THE INSTRUMENTWhen the planet hasno or little intrinsic
magnetic field, direct interaction between the solar wind and the
atmosphere occurs and causes the escape of atmosphere through
thermal and non-thermal heating process.
4.2.3.3 ESCAPE OF ATMOSPHERE ON MARSIn upper atmosphere hydrogen
and deuterium atoms are produced by photo dissociation from H2O and
HDO molecules. In the escape of these atoms, the D/H ratio in the
atmosphere increases with time because escaping ratio of H atoms is
expected to be greater than that of D atoms because of the mass
difference.
Figure 4.13: Escape of atmosphere on marsLAP operates on the
principle of resonant scattering and absorption at Lyman alpha
wavelengths of H and D i.e., 121.56 nm, 121.53 nm respectively.
Thermally dissociated H2 and D2 molecules in the cells absorb the
incoming H2/D2 Lyman alpha incident on the cell.
4.2.3.4 TECHNICAL SPECIFICATIONS OF LAPThe fore-optics
comprising of a plano-convex lens collects the input radiation and
transmits to the gas cells. Gas filled cells of the instrument
works as an effective narrow band-pass rejection filter at hydrogen
and deuterium alpha wavelengths. Tungstun filament is used to
thermally dissociate the gases in to atoms. There atoms will
resonantly absorb the incoming hydrogen/deuterium lyman alpha
radiation at their wavelengths. A 15 nm bandwidth lyman alpha
filter placed in the front of the detector cuts-off the undesirable
radiation that lies outside the wavelength range of interest and a
solar-blind side-on type photo multiplier tube(PMT) is selected for
photon detection.
Figure 4.14: Photo multiplier tube (pmt)
Range of operation : 3000 km-periapsis-3000 kmSize (cubic meter)
: 27.6 X 138 X 100.5Mass (kg) : 1.97Total power (watt) : 8
4.2.4 MARTIAN EXOSPHERIC NEUTRAL COMPOSITION ANALYSERMENCA
payload weighing 3.56 kg, is a quadrupole mass spectrometer based
scientific payload on MOM, capable of measuring relative abundances
of neutral constituents, in the mass range of
Figure 4.15: Martian exospheric neutral compositionMENCA payload
weighing 3.56 kg, is a quadrupole mass spectrometer based
scientific payload on MOM, capable of measuring relative abundances
of neutral constituents, in the mass range of 1-300 amu .The core
objective of MENCA is to study the exospheric neutral density and
composition at altitudes as low as 372 kilometers above the Martian
surface. The instrument examines radial, diurnal and seasonal
variations in the Martian exosphere with Mangalyaan in its
operational orbit, MENCA is to estimate the upper limits of the
neutral density distribution and composition around mars. Studying
Martian exosphere will provide valuable data on the present
conditions.4.2.4.1 SCIENCE GOAL MENCA would provide the first ever
institute measurement of the neutral composition and density
distribution of the Martian exosphere (atmosphere ~ 500 km and
beyond from the Martian surface).
Explanation on what happens in a mass spectrometerAtoms can be
deflected by magnetic fields-provided the atom is first turned into
an ion. Electrically charged particles are affected by a magnetic
field although electrically neutral ones arent. The atom is ionized
by knocking one or more electrons off to give a positive ion. This
is true for things which you would normally expect to form negative
ions(chlorine for example) or never form ions at all( ex: argon).
The ions are accelerated so that they all have the same kinetic
energy. The ions are then deflected by a magnetic field according
to their masses. The lighter they are, the more they are deflected.
The more the ion is charged, the more it gets deflected. The beam
of ions passing through the machine is detected electrically. .
Figure 4.16: Mass spectrometer
4.2.4.2 Principle of working
4.2.4.2.1 Quadrupole rodsIt consists of four parallel metal rods
with opposing rods being connected electrically. A radio frequency
voltage is applied between the two pairs of rods and a direct
current voltage is applied between the two pairs of rods and a
direct current voltage is then superimposed on the RF voltage. Ions
entering the instrument travel down the quadrupole between the
rods. Depending on their mass-to-charge ratio, ions either enter
unstable trajectories and collide with the rods or make it through
to the detector (detectors being used in MENCA are channel electron
multiplier (CEM) and Faraday Cup (FC) Figure 4.17: Quadrupole mass
spectrometer
4.2.4.2.2 Electron multiplierThe m/z of ions reaching the
detector is a function of the voltage setting which allows the
operator to select an ion with a particular mass-to-charge ratio to
measure its abundance or run the instrument through a range of
voltages to scan for a number of species that might be present.Ions
are generated via electron ionization Figure 4.18: Electron
multiplierElectrons are produced through thermionic emission. The
electrons are accelerated in an electric field and focused into a
beam by a trap electrode. The atoms and molecules enter the ion
source perpendicular to the electron beam. As high-energy electrons
pass by and collide with the particles, large fluctuations in the
electric field around the neutral molecules are caused leading to
ionization and fragmentation. Figure 4.19: Working of electron
multiplierThe MENCA instrument operates at an m/z range of 1 to 300
amu (atomic mass unit) with a mass resolution of 0.5u which allows
detailed detection of species. The instrument can operate at the
low partial pressure found in the upper Martian atmosphere.
4.2.4.3 Additional instruments in MENCA payloadMENCA has an
in-built pressure gauge for the measurement of total pressure. The
instrument has a provision to study the time-evolution of a set of
selectable species in the mode of operation. The primary science
goal of MENCA is the in-situ measurement of neutral composition and
distribution of the martian upper atmosphere and exosphere and to
examine its radial, diurnal and possibly seasonal variations. The
instrument has tele command, telemetry and data interface to the
space craft optical combination of operating parameters which can
be chosen through tele commands will be used to control the
instrument at different observation phases so that best possible
scientific data could be derived.4.2.4.4 Tele-command architecture
Figure 4.20.1: Satellite telemetry structure 4.2.4.5 CONCLUSIONDue
to various thermal and non-thermal processes, Mars lost its
atmosphere deserting it in its present form. Study of the
composition and the distribution of the Martian Exosphere by MENCA
may help in understanding the thermal escape of the Martian
atmosphere.4.2.5 THERMAL INFRARED IMAGING SPECTROMETERMars is a
terrestrial planet which means that its bulk composition, like
Earth consists of silicates, is metals and other elements that
typically make up rock. Also like Earth, Mars is a differentiated
planet, meaning that it has a central core made up of metallic iron
and nickel surrounded by a less dense silicate mantle and crust.
The planets distinctive red colour is due to oxidation of iron on
its surface.The knowledge on type of minerals present in any
planetary system provides the information on the conditions under
which minerals are formed and process by which they are weathered.
Much of what we know about the elemental composition of Mars comes
from orbiting spacecraft and landers. Most of these spacecraft
carry spectrometers (A spectrometer is an instrument used to
measure properties of light over a specific portion of the
electromagnetic spectrum, typically used in spectroscopic analysis
to identify materials) and other instruments to measure the surface
composition of Mars. Figure 4.21: Thermal Infrared Spectrometer
payload on MOM
Thermal Infrared Spectrometer is one of the five instruments on
MOM. TIS weighing 3.2 kg can measure the thermal emissions and can
be operated during both day and night. Temperature and emissivity
are the two basic physical parameters estimated from thermal
emission measurement. The TIS instrument measures thermal emissions
from the Martian surface to deduce surface composition and
mineralogy.
4.2.5.1 Science goals of TIS are To estimate ground temperature
of Mars surface. To map surface composition and mineralogy of Mars.
To detect and study the variability of aerosol/dust in Martian
atmosphere. To detect hot spots, which indicate underground
hydrothermal systems.TIS will be useful in mapping mineral
compositions and surface temperature during perigee imaging (The
perigee is the point in a satellite's elliptical path around the
earth at which it is closest to the center of the earth)and it will
be used for assessment of global temperature distribution and
aerosol turbidity in Martian atmosphere during apogee
viewing(apogee is the point in the orbit of an artificial satellite
most distant from the center of the earth).
Figure 4.22:3D Image of TISThe TIS instrument consists of a
spectrometer that features a typical infrared grating spectrometer
design. TIS consists of fore-optics, slit, collimating optics,
grating and re-imaging optics. A 120X160 element bolometer array is
placed at the focal plane of the re-imaging optics
4.2.5.2 Fiber-port lens positions for collimating Figure 4.23.1:
Sketch of multi wavelength re-imaging optics
Figure 4.23.2 Working of TIS
In the common design, radiation is directed through an entrance
slit (available light energy depends on light intensity of the
source as well as the dimensions of the slit and acceptance angle(
acceptance angle refer to the angle in an optical fiber below which
rays are guided rays) of the system. The slit is placed at the
effective focus of a collimator (A collimator is a device that
narrows a beam of particles or waves, which means either to cause
the directions of motion to become more aligned in a specific
direction (i.e., collimated or parallel) or to cause the spatial
cross section of the beam to become smaller.) that directs
collimated radiation (focused at infinity) to a diffraction grating
that acts as dispersive element. Another mirror refocuses the
dispersed radiation onto a detector.TIS uses a 120 by 160 element
bolometer array detector. A bolometer is a device for measuring the
power of incident electromagnetic radiation via the heating of a
material with a temperature-dependent electrical resistance
4.2.5.3 Principle of operation of a bolometerPower P from an
incident signal is absorbed by the bolometer and heats up a thermal
mass with heat capacity C and temperature T. The thermal mass is
connected to a reservoir of constant temperature through a link
with thermal conductance G. The temperature increase is T = P/G.
The change in temperature is read out with a resistive thermometer.
The intrinsic thermal time is T=c/g. Figure 4.24: Bolometer
A bolometer consists of an absorptive element, such as a thin
layer of metal, connected to a thermal reservoir (a body of
constant temperature) through a thermal link. The result is that
any radiation impinging on the absorptive element raises its
temperature above that of the reservoir the greater the absorbed
power, the higher the temperature. The intrinsic thermal time
constant, which sets the speed of the detector, is equal to the
ratio of the heat capacity of the absorptive element to the thermal
conductance between the absorptive element and the reservoir. The
temperature change can be measured directly with an attached
resistive thermometer, or the resistance of the absorptive element
itself can be used as a thermometer. Bolometer receivers measure
the energy of incoming photons. TIS is sensitive for an infrared
wavelength range of 7 to 13 microns.
4.2.5.3.1 Micro bolometerThe micro bolometer array does not
require cooling. Each pixel on the array consists of several layers
including an infrared absorbing material and a reflector underneath
it that directs IR radiation that passes through the absorber back
to the absorbing layer to ensure a near complete absorption. As IR
radiation strikes the detector, the absorbing material is heated
and changes its electrical resistance which can be measured via
electrodes connected to each micro bolometer and processed into an
intensity read-out in order to create an IR spectrum.
Figure 4.25: Internal structure of bolometer a) side view b) top
view
4.2.5.3.2 IR spectrum image Figure 4.26: IR Spectrum table
4.2.5.4ConclusionThe analysis of TIS data would involve
estimation of brightness temperature from observed and calibrated
thermal radiance data. The retrieval of surface temperature and
emissivity spectra for different regions would be carried out. The
estimated emissivity spectra would be compared with Mars analog
mineral emissivity spectra. It is proposed to generate the
emissivity spectra between 7-13 microns for minerals reported to
exist in Martian surface. In this way, spectral library will be
used to know the mineral composition on Mars using TIS data.GROUND
SEGMENT CHAPTER 5
5.1 TELEMETRY AND TELECOMMAND
The Indian Space Research Organisation Telemetry, Tracking and
Command Network performed navigation and tracking operations for
the launch with ground stations at Sriharikota, Port Blair, Brunei
and Biak in Indonesia, and after the spacecraft's apogee became
more than 100,000 km, two large 18-metre and 32-metre diameter
antennas of the Indian Deep Space Network started to be utilised.
The 18-metre diameter dish-antenna will be used for communication
with craft till April 2014, after which the larger 32-metre antenna
will be used.NASA's Deep Space Network is providing position data
through its three stations located in Canberra, Madrid and
Goldstone on the US West Coast during the non-visible period of
ISRO's network. The South African National Space Agency's (SANSA)
Hartebeesthoek (HBK) ground station is also providing satellite
tracking, telemetry and command services. Additional monitoring is
provided by technicians on board two leased ships from the Shipping
Corporation of India, SCI Nalanda and SCI Yamuna which are
currently in position in the South Pacific near Fiji.
5.1.1 ISRO Telemetry Tracking and Command Network (ISTRAC) will
be providing support of the TTC ground stations, communications
network between ground stations and control center, Control center
including computers, storage, data network and control room
facilities, and the support of Indian Space Science Data Center
(ISSDC) for the mission. The ground segment systems form an
integrated system supporting both launch phase, and orbital phase
of the mission.
Table 5.1: Ground segment features and specification
Figure 5.1: Stations for tracking and command for ISRO in the
world
5.2 TRACKING AT DIFFERENT PHASES 5.2.1 LAUNCH PHASE The launch
vehicle is tracked during its flight from lift-off till spacecraft
separation by a network of ground stations, which receive the
telemetry data from the launch vehicle and transmit it in real time
to the mission computer systems at Sriharikota, where it is
processed. The ground stations at Sriharikota, Port Blair, Brunei
provide continuous tracking of the PSLV-C25 from liftoff till
burnout of third stage of PSLV-C25. Two ships carrying Ship Borne
Terminals (SBT) are being deployed at suitable locations in the
South Pacific Ocean, to support the tracking of the launch vehicle
from PS4 ignition till spacecraft separation.5.2.2 ORBITAL PHASE
After satellite separation from the launch vehicle, the Spacecraft
operations are controlled from the Spacecraft Control Centre in
Bangalore. To ensure the required coverage for carrying out the
mission operations, the ground stations of ISTRAC at Bangalore,
Mauritius, Brunei, and Biak are being supplemented by Al cantara
and Cuiaba TTC stations of INPE.5.3 MAIN FRAME ELEMENTSThe
spacecraft configuration is a balanced mix of design from flight
proven IRS/INSAT/ Chandrayaan 1 bus. Improvisations required for
Mars mission are in the areas of communication, power, propulsion
(for liquid engine restart after nearly a year) Systems and Onboard
Autonomy.The 390 liters capacity propellant tanks accommodate a
maximum of 852 kg of propellant is adequate with sufficient
margins. A liquid engine of 440 N thrust is used for orbit raising
and Martian Orbit Insertion (MOI).Additional flow lines and valves
have been incorporated to ensure LE 440 N engine restart after 300
days of Martian Transfer Trajectory cruise and to take care of fuel
migration tissues.
Figure 5.2: Different views of mainframe elements
Eight numbers of 22N thrusters are used for wheel de saturation
and attitude control during maneuvers. Accelerometers are used for
measuring the precise incremental velocity and for precise burn
termination. Star sensors and gyros provide the attitude control
signals inputs in all phases of mission.To compensate for the lower
solar irradiance (50% to 35% compared to earth), the mars orbiter
requires three solar panels of size 1800x 1400 mm. Single 36 AH
Li-Ion battery is sufficient to take care of eclipses encountered
during earth bound phase and in mars orbit. The communication dish
antenna is fixed to spacecraft body. The antenna diameter is 2.2 m
is arrived after the trade off study between antenna diameter and
accommodation within the PSLV-XL envelope. On-board autonomy
functions are incorporated as the large earth-mars distance does
not permit real time interventions. This will also take care of
on-board contingencies.
TABLE 5.2: Salient features of space segment
44