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

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