final report on exergy analysis of power plant

8/9/2019 Final Report on Exergy Analysis of Power Plant

1/51

1

[MARS ORBITER MISSION]

A SEMINAR REPORT

SUBMITTED BY

[BHARANIDHARAN K] CB.EN.U4MEE12010

[HARISUDHAN S] CB.EN.U4MEE12018

[KARTHICK R] CB.EN.U4MEE12020

[KARTHIK N] CB.EN.U4MEE12021

[NANDHAKUMAR M] CB.EN.U4MEE12030

Submitted as an Assignment for the Course

MEC301 HEAT POWER ENGINEERING

DEPARTMENT OF MECHANICAL ENGINEERING

AMRITASCHOOL OF ENGINEERINGAMRITA VISHWA VIDYAPEETHAM

COIMBATORE 641 112

November, 2014


2/51

2

ABSTRACT

The interest in the study of our neighboring planet have started in early 1600s, through telescope, whichmade to suspect that it is earth-like planet. Since 1960, there have been 45 missions to mars with just aboutthree of them being successful. The Indian mars mission MANGALYAAN designed and developed byIndian Space Research Organization (ISRO) is to study Martian atmosphere like climate, geology, originand evolution and sustainability of life on the planet. Mars orbiter mission MANGALYAAN is ISROs firstinterplanetary mission designed to orbit the planet Mars. It is launched on 5 thNovember 2013 by PSLV-XL

rocket. The spacecraft started its mars transfer trajectory on November 26

th

2013and placed in a highlyelliptical orbit after the crucial mars orbit insertion maneuver by firing the 440 N Liquid Apogee motor. Oneof the main objectives of MANGALYAAN is to develop the technologies required for the design,

planning, management and operations of an interplanetary mission.


3/51

3

TABLE OF CONTENTS

SL.NO. TITLE PAGE

LIST OF FIGURES 5LIST OF TABLES 6

NOMENCLATURE 7

1 INTRODUCTION TO SPACE1.11.21.3

1.41.5

1.61.7

Laws Of Planetary MotionUniversal Law of GravitationRequirement for Motion In Space1.3.1Motion In Rotating Frame of Reference1.3.2Orbital Velocity1.3.3Orbital PeriodGeo-Synchronous and Geo-Stationary OrbitsEccentricity and Inclination of Orbits1.5.1Inclination: Polar and Retrograde Orbit1.5.2Transfer Orbit

Energy and Velocity Requirement To Reach A Particular OrbitEscape Velocity

8899

101213141417

1718

2 OBJECTIVES2.1 Mission Objectives

2.1.1Researching Mars Orbit 202.1.2Searching for Methane 202.1.3 Detection of Moisture 202.1.4 Finding of Various Gases 21

2.2

33.13.23.33.43.53.63.73.8

2.1.5 Finding RadiationScientific Objectives

CHALLENGES IN SPACE MISSIONPower SystemCommunication SystemGround SegmentPropulsion SystemOn-board AutonomyWeatherComet StrikeOverall Mission Intricacies

2121

2223232324242425

44.1

LAUNCH VEHICLEIntroduction To PSLV

4.24.3

55.15.2

4.1.1 Stages of PSLV4.1.2 Payload FiringPSLV-XL C25Why PSLV-XLC25 Is Preferred Instead of GSLV

SPACECRAFTLiquid Apogee MotorPropulsive Attitude Control System

27333435

3738


4/51

4

5.3 Antenna

6 MISSION PAYLOADS6.1 Complications in Selection of Payloads396.2 Payloads 39

6.2.1 Lyman Alpha Photometer 40 6.2.2 Methane Sensor for Mars 41 6.2.3 Mars Exospheric Neutral Composition Analyzer 42

6.2.4 Mars Color Camera 42 6.2.5 Thermal Infrared Imaging Spectrometer 43

7 TRAJECTORY AND COMMAND

7.1 Three Phases of Mangalyaan 44 7.1.1 Geo-Centric Phase 45

7.1.2 Helio-Centric Phase 46 7.1.3 Martian Phase 467.2 Tracking and Command 46

8 CONCLUSION 47 REFERENCE 50


5/51

5

LIST OF FIGURES

FIG. NO. TITLE

Fig.1.1Attraction in Neighborhood of the Large Sphere9Fig.1.2 Centripetal Force 10Fig.1.3 Force on the Rotating body on the Rotating Frame 11Fig.1.4 Variation of Orbital Velocity with height 12Fig.1.5 Variation of Time Period with height13

Fig.1.6 Geo-Stationary Orbit14Fig.1.7 Eccentricity of an Orbit14Fig.1.8 Orbit Inclination15Fig.1.9 Polar Orbit16Fig.1.10 Geo-Synchronous Transfer Orbit17

Fig.1.11 Velocity of Orbiting bodies with Different Altitudes18Fig.1.12 Variation of Escape Velocity with height above the earth19Fig.4.1 Configuration of PSLV27Fig.4.2 First Stage of PSLV with six strap on motors surrounding it28Fig.4.3 PSLV with four strap of which ignited on ground and two are air lit30Fig.4.4 Second stage of PSLV30Fig.4.5 Combination of third and fourth stage32

Fig.4.6 Payload34Fig.4.7 Assembly of PSLV XLC2536Fig.5.1 Spacecraft38Fig.6.1 View of LAP42Fig.6.2 View of MSM 43Fig.6.3 Details of MENCA43Fig.6.4 MCC equipment44Fig.6.5 TIS assembly44Fig.7.1 Trajectory Design45Fig.7.2 MOM after fifth liquid burn47Fig.7.3 Martian Orbit47Fig.7.4 Tracking and Command48


6/51

6

LIST OF TABLES

TABLE NO TITLE PAGE NO

1.1 Mass and diameter of the planets 82.1 Composition of gases 214.1 Specifications of First Stage Launch Vehicle 294.2Specifications of Second Stage Launch Vehicle 314.3Specifications of Third Stage Launch Vehicle 32

4.4Specifications of Fourth Stage Launch Vehicle 334.5Specifications of Payload Firing 354.6 Differences in PSLV and GSLV 376.1 Payload summary 41


7/51

7

NOMENCLATURE

CH4 -Methane

Superscripts

D/ H -Deuterium to Hydrogen abundance Ratio

AbbreviationsHTPB -Hydroxyl-terminated polybutadieneMMH- MonomethylhydrazineMON3 -Mixed Oxides of NitrogenISRO -Indian Space Research OrganisationPSLV-Polar Satellite Launch VehicleGSLV- Geo-synchronous Satellite Launch VehicleGTO-Geo-synchronous TransferLEO- Low Earth OrbitSIVTC-Secondary Injection Thrust Vector Control

TTC -Telemetry Tracking and CommandingBWG -Beam Wave GuideUMDH -Unsymmetrical Dimethyl hydrazineLAP -Lyman Alpha Photometer

MSM -Methane Sensor for MarsMENCA-MENCA -Mars Exospheric Neutral Composition Analyser

MCC -Mars Color Camera

TIS -Thermal Imaging Spectrometer.TIR -Thermal InfraredLEOS -Laboratory for Electro Optics SystemsSAC -Space Application CentreVSSC -Vikram Sarabhai space Centre

FOV -Field of ViewSOI -sphere of influence


8/51

8

1. INDRODUCTION TO SPACE

1.1 LAWS OF PLANETRY MOTION

Johannes Kepler, a mathematician and astronomer, formulated basic laws of planetary motion based on his

observation

1. Law of Ellipses: The paths of planets in solar system about the sun are elliptical in shape with the

centre of the sun being located at the focus.

2. Law of Equal Areas: The imaginary line from the centre of the sun to the centre of the planets

sweeps out equal areas in equal intervals of time.

3. Laws of harmonies: The ratio of the square of the orbital period of any two planets is equal to the

ratio of the cubes of their average distance from the sun.

The mass and diameter of the planet are shown in this table 1.1

1.2 UNIVERSAL LAW OF GRAVITATION

The force F2-1 acting on the body of mass m2 due to mass m1 at a distance r away to be attractive and

directed towards the centre of the body m1 as shown in Fig1.1 the magnitude of the force is inversely

proportional to the square of the distance of separation between them. The force is given by

2-1= (Gm1m2 ) / 3

The negative sign indicate that force is attractive i.e. m2 is being attracted towards m1 .G is a gravitational

constant and g=6.6701011

Nm2

Kg2.


9/51

9

Fig.1.1 Attraction in the neighborhood of the large sphere

1.3 REQUIREMENT FOR MOTION IN SPACE

The trajectory or path of the motion is referred to orbit. The universal law of gravitation and suitable frame

of reference are used to find the desired orbits and requirement of putting up the spacecraft in orbit. The

orbit could either be elliptical or circular

1.3.1 MOTION IN ROTATING FRAME OF REFERENCE

Consider a body of mass m revolving at a radius R with a constant angular velocity ? around a point o

the x and y coordinates of the mass at any time t are

= --? 2R cos ? t

= --? 2R sin ? t

F=(? ? 2R cos ? t) 2 + (? ? 2R sin ? t) 2

F= m? 2 r

This force is directed in the negative R direction. The force is known as centripetal force and is shown

in Fig.1.2 and the acceleration is called centripetal acceleration = --? 2R. A force (pseudo force) acting

equal and opposite to the above centripetal force must be added in order to make the body stationary. The

pseudo force in the rotating frame of reference is called centrifugal force.


10/51

10

Fig.1.2 CENTRIPETAL FORCE

1.3.2 ORBITAL VELOCITY

The body in the rotating frame of reference will experience a force of attraction due to earth and also the

centrifugal reaction, the force of attraction towards the earth is universal gravitational force as shown in Fig

1.3

2-1= (GmME) / 3 (ME= MASS OF EARTH)

And centripetal reaction acting = m? 2R


11/51

11

Fig.1.3 FORCE ON THE ROTATING BODY ON THE ROTATING FRAME

Hence equating both the force

(GmME) / 3 = m? 2R

The angular velocity from the above equation is

? = (GmME/ R3)

The velocity along the circular path known as orbit velocity

V0=R? =(GmME/ R)

The space craft orbiting around the earth at a height of 200 km has orbital velocity

V0=7.76 km/s

It is clear from the Fig 1.4 that the orbital velocity decrease with increase in altitude. The

Orbital velocity is seen to decrease from a value of about 7.9 km/s at the earths surface to about 2.66km/s ata height of 50000 km.


12/51

12

Altitude (h)

Fig.1.4 VARIATION OF ORBITAL VELOCITY WITH HEIGHT

1.3.3 ORBITAL PERIOD

The time period of revolution for one orbit is

=2R / v0

=2(R3/GME)0.5

=0.304106R1.5

From the Fig1.5 it is clearly understood that the period of revolution increases with increase in altitude. The

time period increases about 1.4 hours at a height of 100km to 35.7 hours at height of 50000 km.

At 35,786 km above the earth, the time period for one orbit is 24 hours (one day)


13/51

13

Fig. 1.5 VARIATION OF TIME PERIOD WITH HEIGHT

1.4GEOSYNCHRONOUS AND GEOSTATIONARY ORBITS

The earth rotates from east to west in a counter clockwise direction and complete one rotation in a day.

The angular velocity if the rotation of the earth is ? E = 2/ 24 3600

= 7.273105 rad/s

If a body is in circular orbit with an angular velocity equal to that of earth and also rotate in the same direction as that

of earth from east to west in counter clockwise direction, its movement is synchronous with the rotation of the earth

and it is said to be in geostationary orbit. It will complete one rotation in a day .the radius of geosynchronous orbit RGis found by equating the angular velocity

2/ 24 3600(? E) = (GmME/ R3)

RG3

= GME/? E2

Substituting the values of G, ME,? E , from the table The value of RG=42,164 km

With the radius of earth being 6378 km, the geosynchronous altitude would therefore be 42164-6378 =35786 km above the earth. This can be well understood by the Fig.1.6 if the body is in geostationary orbit it

would appear stationary for the observer at a point.


14/51

14

Fig. 1.6 GEOSTATIONARY ORBIT

1.5 ECCENTRICITY AND INCLINATION OF ORBITS

Orbits are not always circular and equatorial plane. An orbit could be elliptical and departure from circular

orbit is known as eccentricity. It is defined as the ratio of distance between the foci to the length of the major

axis of the ellipse. This is shown in fig.1.7

Fig.1.7 ECCENTRICITY OF AN ORBIT

A body get into elliptical orbit if the orbital velocity produced is not equal to the orbital velocity required.

1.5.1 INCLINATION: POLAR AND RETROGRADE ORBIT

The inclination of an orbit is the angle between the orbital plane and the equatorial plane of the planet. An

orbit at inclination of is shown in fig.1.8


15/51

15

Fig.1.8 ORBIT INCLINATION

An inclination of zero denotes that orbit is in equatorial plane, i.e. orbital plane and equatorial plane

coincide. An inclination of 90o represent the orbit from pole to pole and is called polar orbit. A circular polar

orbit is shown in Fig.1.9


16/51

16

Fig.1.9 POLAR ORBIT

Polar orbit is useful for spacecraft that carry mapping of the planet. This is because as the planet rotates, the

spacecraft has to access to every point on the surface of the planet. In the case of polar orbit about the earth,


17/51

17

the sun would maintain the same inclination with line joining the earth and the spacecraft on the orbital

plane remain constant. And this orbit is called sun synchronous orbit

1.5.2 TRANSFER ORBIT

Instead of directly sending the spacecraft to geosynchronous orbit, the spacecraft is initially placed in an

elliptical orbit about the earth. This orbit is known as the geosynchronous transfer orbit about the earth and it

is shown in Fig 1.10. The point of the spacecraft closest to earth is called perigee and farthest is apogee. The

apogee of the elliptical orbit so chosen that it is near to the radius of geosynchronous orbit of 42,164 km.The perigee is 6,630km. The elliptical transfer orbit is circularized to give the geosynchronous orbit.

Fig.1.10 GEOSYNCHRONOUS TRANSFER ORBIT

1.6 ENERGY AND VELOCITY REQUIREMENT TO REACH A PARTICULAR ORBIT

In order to reach the given orbit from the surface of a planet, work has to be done to overcome attractive

force and centripetal force of the planet.

The total energy required to orbit a body of mass m at an altitude h

ET=GME(RE + 2h) / 2RE(RE + h)

The total velocity VT required for orbiting the body is obtained by

ET = mvT2 / 2

VT= GME(RE + 2h) / 2RE(RE + h)


18/51

18

The total energy and velocity requirement to reach particular orbit can be determined by the above equation.

And the variation in total velocity and orbital velocity is shown in Fig 1.11.

Altitude

Fig 1.11 VARIATION OF ORBITTING BODIES AT DIFFERENT ALTITUDE

1.6 ESCAPE VELOCITY

If a body is provided with sufficient velocity to overcome the gravitational attraction, it would reach infinite

radius and escape the attractive force of the planet.

VE= 2GME / RE (DERIVED FROM ENERGY EQUATION)

The value of escape velocity VE=2 V0

The variation of escape velocity with height above the earth is shown in Fig.1.12


19/51


20/51

20

2. OBJECTIVES

2.1 MISSION OBJECTIVE

There are various technological and scientific objectives of sending the Mangalyaan. First and importantobjective is to develop technologies for the operation of interplanetary missions. Then after know about thedesign of the mars orbital with the capability to survive and perform Earth bound manoeuvres, cruise phase

of 300 days, Mars orbital insertion, space communication, navigation, mission planning and exploration ofmars surface to know about the features of the surface and its atmosphere and also detect the presence oflife.

2.1.1 RESEARCHING MARS ORBIT

The main crucial part is to hit the destination and for this ISRO focuses and study the process ofsending the Mangalyaan into Mars orbit which the huge part ofMOM (Mars Orbital Mission).This launch isvery useful as it helps the ISRO to studymore about Interplanetary missions by developing technologieswhich are needed fordesigning, planning, operating and managing the interplanetary missions. That is whyit is so called the technological part of the mission Orbit manoeuvres to transfer the probe from Earth-cantered orbit to heliocentric trajectory and finally capture into Martian orbit.

2.1.2 SEARCHING FOR METHANE

One of the main objective of Mangalyaan is to look that whether the methane ispresent on the red planetor not. Methane is considered as the host chemical fordetermine the possibility of life on mars even in theform of microbe. The Methanesensor For Mars (MSM) is the main sensor which detects the presence ofmethane andconfirms its presence and give the proof of life on Red planet. Methane breaks up in the

presence of ultraviolet solar radiation. Based on photochemical models and on the current understanding ofthe composition of the Martian atmosphere, methane has a chemical lifetime of about 300-600 years, whichis very short on geological time scales. This implies that the methane that is observed today cannot have

been produced 4.5 billion years ago, when the planets formed. So what can explain the presence of this gason the Red Planet? If the methane on Mars is biotic, two scenarios could be considered: either long-extinct microbes,which disappeared millions of years ago, have left the methane frozen in the Martian upper subsurface, orthis gas is being released into the atmosphere today as temperatures and pressure near the surface change.Methane is also a geological origin. It could be produced, for example, by the oxidation of iron, similar to

what occurs in terrestrial hot springs, or in active volcanoes. This gas could have been trapped in solid formsof water, or cages that can preserve methane of ancient origin for a long time. These structures are known asclathrate hydratesBiological origin of methane would be to measure the isotope ratios of carbon and hydrogen, the two

elements in methane. Life on Earth tends to use lighter isotopes, for example, more Carbon-12 than Carbon-13, because this requires less energy for bonding.

2.1.3 DETECTION OF MOISTURE

Mangalayaan also have to detect the moisture content on its surface by clicking the pictures of the MartinSurface. As it is widely believed that there is the presence of large ice burgs but no water bodies TheMangalyaan objective is to prove it and if it is proven that the mars in known to be the only planet afterearth to have life-giving liquid.

On the protest of water it is concluded that water would be deposit on Martian surface would evaporatesbefore it could liquefy .As it warms towards 0C it will evaporates increase rate. The ice loss hear


21/51

21

proportionally equal to evaporation rate. The evaporation frost deposit at constant temperature in Martianatmosphere .so Temperature of Martian could be maintained at 0C .The melting point of ice and the frostwould probably disappears by the temperature reaches the melting point

2.1.4 FINDING OF VARIOUS GASES

Checks the presence of neutral gases in its atmosphere. It alsodetects the presence of Hydrogen anddeuterium which are considered as theessential elements which gives the proof about the presence of waterin the pastand may be also gives the detailed history of water presence.The atmosphere of Mars is 100 about

times thinner than earth and the composition of gases is shown in table 2.1Table 2.1

The composition of gases

Carbon-di-oxide 95.32%

Argon 1.6%

Oxygen 0.13%

Carbon-mono-oxide 0.08 %

And also small amount of water vapour, nitrogen oxide, Neon, Hydrogen, Krypton and Xeon

2.1.5 FINDING RADIATION

Instruments on Mangalyaan also study the effect of solar radiation on Martinatmosphere and its surface which may give the reason of decomposing and eroding ofits surface.Radiation on Mars comes from two sources: galactic cosmic rays and solar energetic particles. The sun hashad a muted peak to its solar cycle so that affect the expected amount of particles on Mars. But the MarsCuriosity rover, in its first 300 Earth days of roaming, has plenty of data on galactic cosmic rays.On the Martian surface, the average dose is about 0.67 mSv per day.

2.2 SCIENCE OBJECTIVES

The scientific objectives deal with the following major aspects: Exploration of Mars surface features by studying the morphology, topography and mineralogy using

specific scientific instruments. Study the constituents of Martian atmosphere like methane, CO2, etc. using remote sensing techniques

Study the dynamics of upper atmosphere of Mars, effects of solar winds and radiation and the escape ofvolatiles to space.


22/51

22

3. CHALLENGES IN SPACE MISSION

Since time immemorial undertaking developments in technologies have always been

challenging and same is the case with space technologies. Over a period of time, various space technologies

particularity in the fields of communication, navigation, remote sensing etc. have reached to a certain level

of maturity. However, in the fields like deep space much needs to be achieved in regards to technology. For

many years, humans have ambition to reach Mars and subsequently to establish human colonies over there.

Presently none of the space-faring states are even somewhat close to realise this dream. Mars offers the

greatest challenge for the rocket scientists, space scientists, aerospace engineers and astronomers. Hence, forany state developing the missions, the Mars mission would always have to face challenges at various planes.

The scientists basically have to confront situation in regards to the technological challenges and at times the

challenges posed by the nature.

The success rate in case of various Mars missions undertaken so far is not encouraging (approximately 1 in 3becomes successful); hence, the scientific community over a period of time has become well aware of thechallenges. They have become aware of the fact that strategic knowledge gaps do exit and are trying their

bests to overcome them.Over the years, it has been observed that the different categories of Mars missions has their own set ofchallenges. Flyby, orbiter, rover and lander missions have few common challenges while there are alsomany mission-specific challenges. For a spacecraft to navigate through weather, gravity and radiation for alonger distance and survive for minimum of 1 year and keep the communication alive with the earth stationsis extremely challenging. Historically, it has been witnessed that space agencies learn from the successesand misfortunes of previous missions but at times in spite of taking all precautions based on the previousexperiences they end up is experiencing some new set of challenges. Hence it is also important to remain

prepared to address surprises.The unmanned missions are expected to encounter various environmental and technological challengeswhile the human missions could face technical, physical, and psychological challenges. Human Marsmissions (when it happens) risks of radiation exposure, de-conditioning and the psychological impacts ofisolation. There are very many challenges in devising human missions to Mars. ISRO had their ownexperiences in the deep space area with their Moon mission. Following paragraphs indicates that how ISROovercame some of those challenges.

3.1 Power Systems

The power system required to support the mission during various phases of mission liketransfer orbit and on-orbit phase. The power system consists of power Generation, energy storage and powerconditioning elements. One of the major challenges in the design of power system is due to the largerdistance from Sun the dependence on solar energy reduces. The Sun is a very powerful, clean andconvenient source of power, particularly for satellites. The only thing needed is a means to convert theenergy contained in the suns radiation into electrical power. The most efficient way to achieve this today is

by using panels composed of semiconductor photovoltaic cells. These were first used in space in 1958 topower the Vanguards satellite.

Mars is the last planet in the solar system where solar power generation can be usedeffectively. The power generation in Mars orbit is reduced to nearly 50 % compared to Earths orbit. Due tothe eccentricity of Mars orbit around sun, the power generation variation is nearly 15 % . If one Watt is

power generation when Earth is at perihelion, 0.35 W will be the power generation when Mars is ataphelion. The ordeal for the scientific community is to cater for such power generation variations at differentdistances and in different orbits. Also, there are specific requirements for powering the spacecraft duringeclipse phase and payload data download phases. ISRO made specific provisions to cater for such challengesand they succeeded it. Nothing can change its sate or position without energy. Just like many othermachines, satellites also need electrical power to function.


23/51

23

What happens when the sun is hidden? Solar power generation is very convenient in space, especially because there are noclouds and the sun never sets. Or does it? Satellites orbiting planets through a shadow region on the oppositeside of the earth from the sun. Depending on the type of orbit, this can happen just a few times a year. Theseare so called eclipses. The solar panels cannot produce electrical energy and the satellite would not only

be unable to operate, but would also freeze to incredibly low temperature. Therefore, electrical energy has tobe stored. They can be stored in rechargeable batteries or electrical accumulators.

3.2 Communication Systems

The communication systems for the Mars mission are responsible for the challenging task ofcommunication management at a distance of nearly 200400 million km. For one way communication (fromearth station to Mars orbiter and back), it would take approximately 20-min time. Hence, ISRO would haveto cater for 40 min delay in the communications and make the planning accordingly to the receiver. To caterfor such requirements ISRO has invested into the Telemetry, Tracking and Commanding (TTC) systems andData transmission systems. For this purpose, ISRO has developed a major infrastructure for the antennasystem constituting of low, medium and high gain antennas. In fact, a major structure was erected during theMoon mission (2008) and that has been upgraded to cater for the requirements related to Mars.

3.3 Ground Segment

To support such a complex mission, it is essential to have a state-of-art ground

infrastructure available to realise the operational communications between a control centre and a spacecraft.This being Indias second deep space mission luckily; the ground station is in place but in regards todistance as compared to Moon, and Mars is a quantum jump. The Indian Deep Space Network station(IDSN32), a 32-m-diameter network antenna system located in Byalalu , Bangalore, was established with aview to meet, not only the requirements of Chandrayaan-1 mission, but also ISROs future missions toMercury, Venus and up to Mars . So, the specifications of IDSN-32 are arrived such that it will be able tocater TTC and science data reception functionalities for a mission to Mars. The hallmark of Indian DSNfacility is the state-of-the-art technology 32-m-diameter Beam Wave Guide (BWG) antenna that had beenindigenously designed, developed and installed to support all future deep space missions of ISRO. For Marsmission certain amount of mission-specific modifications have been carried out keeping in view therequirements of this mission. To manage a deep space mission for 24 hours, the geometry necessitates thesupport of minimum two ground stations: one located in eastern and other in western hemispheres. Ground

support from JPL-NASA stations is envisaged to support the mission in addition to IDSN support.The long coasting of PSLV PS4 stage for 1644s before PS4 operation requires two portable sea-borne S-

band terminals to be deployed in Pacific Ocean to monitor PS4 and satellite separation.

3.4 Propulsion Systems A spacecraft propulsion system is a method used to accelerate the spacecraft(satellite), and there are different ways to do this with each having some advantages as well as limitations.Reaching Mars could be viewed as one of the most challenging task for the rocket scientists. This mission

being first attempt by ISRO to reach such a distance, they are putting significant amount of efforts todevelop their systems. Their propulsion systems embody the truly enabling technology for departing earthand reaching the Mars. The Mars Orbiter Mission propulsion system has its heritage from GEO mission andconsists of a unified bipropellant system for orbit raising and attitude control. It consists of one 440-NLiquid Engine (LE440) and 8 numbers of 22-N thrusters. The propellants are stored in Titanium propellant

tanks each with a capacity of 390 L pressurised with Helium gas. The tanks have combined storage capacityup to 850 kg propellant. The 67 litre helium pressurant tank is used to pressurise the propellant. The 22-Nthrusters are used for attitude control during the various phases of the mission like orbit raising using liquidengine, attitude maintenance, Martian orbit maintenance (if any) and momentum dumping. As the last fewLE burns around Mars are to occur after 10 months of launch (this is because it would take 300 days, time toreach Mars), suitable isolation techniques are adopted to prevent fuel/oxidiser migration issues. For around300 days the craft has to experience the vagaries of space weather including the radiation. The biggestchallenge is that after undertaking the journey of 9/10 months the all systems of the orbiter need to restartand function properly.


24/51

24

3.5 On-board Autonomy: On-board autonomy refers to the capacity of the orbiter to make its own decisionsabout its actions. As the distance between the Mars Orbiter and earth increases, the need for autonomyincreases dramatically. Given an average round trip time (to and fro) from earth to Mars of approximately 40min, it would be impractical to micromanage a mission from Earth. Due to this communications delay,mission support personnel on Earth cannot easily monitor and control all the spacecraft systems in real-time

basis. Therefore, it is configured to use on-board autonomy to automatically manage the nominal and non-nominal scenarios on-board the spacecraft. Autonomy is in charge of the spacecraft when communication

interruptions occur when the spacecraft it occulted by planet Mars or sun. Autonomy also ensures therecovery from safe mode occurrences on-board the spacecraft.

3.6 Weather The weather on Mars is unpredictable. One week, the sky is pink and cloudless,filled with windblown dust raised from the rusty Martian surface. By Martian standards, it's warm, aboutminus 40 degrees Fahrenheit. Then, in a matter of days, the dust is swept from the atmosphere, temperatures

plummet 40 degrees, and brilliant water ice clouds appear against a dark blue sky.Dramatic weather changes like these may not seem very different from a batch of severe thunderstorms

passing through your home town, but for Mars these changes can sweep over the entire planet every week. Itappears that Mars' roller coaster-like weather is more chaotic and unpredictable than scientists first thought.Observations by the Hubble Space Telescope and the National Radio Astronomy Observatory (NRAO)

radio telescope at Kitt Peak, Ariz., show that the atmosphere of Mars is more complex and variable than thepicture revealed by the Viking and Mariner 9 orbiters. These spacecraft collected information from theplanet in the 1970's and painted a fairly one-dimensional picture of Mars' climate. Images snapped by theorbiters revealed huge dust storms spreading throughout the entire atmosphere when Mars was closest to thesun (perihelion). These dusty conditions continued to dominate the planet's climate when it was farthestfrom the sun (aphelion).

But information captured by Hubble and NRAO show that Mars is more often cloudy thandusty, experiencing abrupt planet-wide swings between dusty and hot and cloudy and cold. A state ofemergency would be declared on Earth if an ice or dust storm blanketed the entire planet. These shifts inclimate are driven by three important factors: Mars' thin atmosphere, its elliptical orbit around the sun, andstrong climatic interactions between dust and water ice clouds in the atmosphere. Mars' atmosphere is sothin that it weighs less than 1 percent of Earth's atmosphere. Because Mars' atmosphere is so paper-thin and

there are no oceans to store up heat from the sun, the planet's temperatures respond more quickly andintensely to surface changes and atmospheric heating by the sun. There are also much larger annual changesin sunlight falling on Mars than on Earth, because Mars' distance from the sun varies by 20 percent in itsorbit around the sun every two years.Mars' elliptical orbit leads to planet-wide changes in atmospheric and surface temperatures over the courseof a Mars year. During perihelion, when Mars is closest to the sun (summer in the southern hemisphere), the

planet receives 40 percent more sunlight than during aphelion, when it is farthest from the sun (summer inthe northern hemisphere). This annual variation in sunlight causes 35-degree Fahrenheit increases duringsouthern summer (perihelion), forcing continental-scale dust storms at the planet's surface. The dust is sweptaloft to altitudes of tens of miles, where it spreads globally, absorbs light from the sun, and heats the entireatmosphere by another 30 to 50 degrees Fahrenheit. This dusty perihelion climate was observed by Vikingand Mariner 9 and by NRAO in 1992, 1994, and 1996.

It has been emphasised that for Mars Orbiter Martian atmosphere, real-time climate conditions and radiationcould pose some challenges. ISRO prepared to handle the challenges of bad weather (if any) on ground andextending to stratosphere (50/60 km above the earths surface), during the launch phase. From the suitabilityof weather point of view, month of November is not the most appropriate time to plan the launch. However,

because of the astronomical compulsions ISRO had no options but to abide by the available launch window.

3.7 Comet Strike: When Indias mission is its final stages of preparations, suddenly a new challenge

has arisen and this is not a technological challenge, but a challenge of entirely different nature which


25/51

25

normally no one could have even imagined to occur.NASA has predicted that there is an outside chance thata newly discovered comet might be on a collision course with Mars. Astronomers determined the trajectoryof the comet, named C/2013 A1 (Siding Spring), but at the very least, it came fairly close to the Red Planetin October of 2014. The nucleus of the comet is probably 13 km in diameter, and it is approached fast,around 56 km/s (125,000 mph) speed. If it does hit Mars, then it is likely to deliver as much energy as 35million megatons of TNT (may be leave a crater about 500 km wide and 2 km deep). For comparison, theasteroid striked Earth 65 million years ago was about three times as powerful, 100 million megatons. TheMars comet is packing 80 million times more energy than that relatively puny meteor. Initially, the scientistswere putting the odds of likely impact at 1 in 2000 and subsequently to about 1 in 8,000. However, based on

Apr 7, 2013, data set it has been calculated that the chances of the comet impacting the Red Planet are about1 in 120,000. This comet which was only discovered on Jan 3, 2013 was posing a major challenge for theIndia decision makers. What was at stake was the Indian investments of about US$ 83 million. But now therecent assessment must have given a breather to Indias mission managers!. Most comets have presence of methane in their tails. One of the important scientific objectives of theIndias mission is assessment in regards to methane in Martian atmosphere. Amongst Indias five payloads,one is the methane sensor for Mars (MSN). There is a good chance that this MSN payload may confuse themethane it detects from the comet as that of Mars and transmit wrong data. Such data could be misleading.This knowledge of the likely presence of the comet close to Mars offers a challenge to the ISROs scientificcommunity and they would have to factor this possibility in their mission planning.

3.8 Overall Mission Intricacies:

Indias mission could be viewed as simple mission in comparison with the most of othermissions launched so far. However, it is important to note that since globally very few missions havesucceeded so far, hence many scientists all over the word are keenly monitored Indian investment in theMars project. They felt that there is much to learn from the way India has devised this project and are keenlylooking forward for the information which could be made available.For ISRO, Mars mission is a technology demonstrator until the craft successfully reached Mars. Hence,there could be different set of challenges when the mission reached Mars. There would be five different

payloads undertaking various observations. Every payload has a specific role. There exists a possibilitysome unknown technological challenges from sensor malfunction to communication breakdown couldemerge at any point in time.


26/51

26

4. LAUNCH VEHICLE

4.1 INTRODUCTION TO PSLV

PSLV, is commonly known as polar launch satellite vehicle and it is used for launching the remote sensing

satellites into orbits. The systems are developed by ISRO inertial system unit (IISU)at Thiruvanathapuram.

PSLV is the most successful operational launch vehicle of ISRO. There had been 22 continuously successful

flights of PSLV, till February 2013.As of 2014 the PSLV has launched 71 spacecraft (31 Indian and 40

foreign satellites) into a variety of orbits. PSLV has successfully undertaken Sun-synchronous, Geo-synchronous Transfer Orbit (GTO) and low inclination missions in the past. The height of PSLV launch

vehicle is 44.4 metres tall and a weight of 295 tonnes. The capacity of launch vehicle differs, if PSLV is to

launched in Low Earth Orbit (LEO) the capacity will be around 3200 kilograms, if it is to be launched in

Sun- synchronous Orbit (HCO) the capacity will be around 1600 kilograms, if it is to be launched in Geo-

synchronous Orbit (GTO) the capacity will be around 1400 kilograms. PSLV continues to be the workhouse

of INDIAN REMOTE SENSING satellites especially for Low Earth Orbits(LEO).PSLV has four stages

using solid and liquid propulsion systems alternately[Fig.4.1].The vehicle can fly in three different

configurations to adjust for mission requirements. The first stage is solid-fuel rocket booster,the booster

develops a maximum thrust.To perform PSLV launch in standard mode, cluster of six strap-on[Fig.4.2] is

attached to the first stage motor, four of which are ignited on the ground and two are air-lit[Fig.4.3].Thesecond stage is liquid fuelled while the third stage is a solid rocket motor. The Upper Stage of the PSLV

uses liquid Propellant.


27/51

27

Fig.4.1-Configurations of PSLV

4.1.1STAGES OF PSLV

FIRST STAGE

The first stage of the PSLV, is a solid fuelled rocket stage. It is 20.34 meters long and 2.8 meters in

diameter with an empty mass of 30,200 Kilograms. The stage contains 138,000kg of HTPB

(Hydroxyl-terminated polybutadiene) bound propellant at liftoff. PS1 has a vacuum thrust of 4,860

Kilonewton [Table 4.1]

First stage attitude control is provided by a Secondary Injection Thrust Vector Control (SITVC) foryaw and pitch. For SITVC, Strontium Perchlorate is used as a secondary fluid that is injected from

the side into the Hypersonic Flow in the nozzle, causing a lateral thrust element that is precisely

steered for attitude control. The SITVC fluid is stored in two cylindrical aluminium tanks strapped to

the core stage. The tanks are pressurized with Nitrogen. Roll control is provided by two Roll Control

Thrusters that are mounted radially on opposite sides of the core stage between the Solid Rocket

Boosters.

The first stage burns for 105 sec and is separated at an altitude of 76 KM by a flexible linear shaped

charge. Separation motors are used to ensure the first stage moves away from the second stage before

it ignites.


28/51

28

Fig.4.2-First stage of PSLV with six strap-on Motors surrounding it

Table4.1-Specifications of First Stage Launch Vehicle

Type PS1

Inert Mass 30,200kg

Launch Mass 168,200kg

Diameter 2.8m

Length 20.34m

Propellant Solid - HTPB Based

Fuel Mass 138,00kg

Propulsion PS1 Solid Rocket Motor

Thrust (Vacuum) 4,860Kn

Impulse 269s

Burn Time 105sec

Restart Capability No


29/51

29

Attitude Control SITVC (Pitch&Yaw), Roll RCS

Stage Separation Flexible Linear Shaped Charge

Separation Motors

Fig:4.3, PSLV with four strap of which are ignited on the ground and two strap are air-lit

SECOND STAGE

The second stage of the PSLV rocket is 2.8 meters in diameter, 12.8 meters long and has a liftoff mass of

46,000 Kilograms with an empty mass of 5,300 Kilograms. The stage uses Unsymmetrical Di-

Methylhydrazine / Hydrazine Hydrate as fuel and Nitrogen Tetroxide as oxidizer. It is powered by a single

799-Kilonewton Vikas engine. The engine has a dry weight of 900 Kilograms and operates at a chamber

pressure of 58.5 bar and provides a specific impulse of 294 sec. The second stage burns for 158 seconds

[Table4.2]. Vehicle control during second stage flight is accomplished by gimbaling the main engine up to 4

degrees, and roll control is provided by a Hot Gas Reaction Control Motor. Second stage separation is

accomplished with a merman band system and separation motors [Fig:4.4]


30/51

30

Fig:4.4 second stage of PSLV

Table 4.2-Specifications of Second Stage Launch Vehicle

Type PS2

Diameter 2.8m

Length 12.8m

Inert Mass 5,300kg

Launch Mass 46,000kg

Fuel Unsymmetrical Dimethylhydrazine

Oxidizer Nitrogen Tetroxide

Propellant Mass 40,700kg

Propulsion 1 Vikas

Thrust 799Kn

Impulse 293s

Engine Dry Weight 900kg

Burn Time 158sec

Chamber Pressure 58.5bar

Mixture Ratio 1.7 (Ox/Fuel)

Area Ratio 31

Prop Flow Rate 278.04kg/s

Attitude Control Main Engine Gimbaling, Roll RCS

Stage Separation Merman Band, Sep Motors


31/51

31

THIRD STAGE

The third stage in PSLV is a solid rocket motor that also uses an HTPB based propellant. The third stage of

the PSLV launcher has a reduced diameter of 2.02 meter and a length of 3.54 meters. It has an empty mass

of 1,100 Kilograms and a liftoff weight of up to 7,800 Kilograms. The stage has a burn time of 112 secondsduring which it generates a thrust of 244 Kilo Newton[Table-4.3].The stage has a Kevlar-polyamide fibre

case and a submerged nozzle equipped with a flex-bearing-seal gimballed nozzle that can be gimballed by

up to 2 degrees for Thrust Vector Control. Roll control is provided by the Reaction Control System of the

fourth stage. The third stage separates at an altitude of 580 Kilometres.[Fig:4.5]

Fig.4.5-combinations of third and fourth stages

Table4.3-Specifications of Third Stage Launch Vehicle

Type PS3

Diameter 2.02m

Length 3.54m

Inert Mass 1,100kg

Launch Mass 7,800kg

Propellant Solid - HTTP Based

Propellant Mass 6,700kg

Propulsion S-7

Thrust 244kN

Impulse 294s

Burn Time 112sec

Attitude Control 4th Stage RCS

Separation Ball Lock


32/51


33/51

33

4.1.2 PAYLOAD FIRINGThe Payload Firing, or "Heat Shield" is positioned on top of the stacked vehicle and its integrated Payload. It

protects the spacecraft against aerodynamic, thermal and acoustic environments that the vehicle experiencesduring atmospheric flight. When the launcher leave the atmosphere, the firing is integrated by pyrotechnicalinitiated systems. Separating the firing as early as possible increases launcher performance.The typicalPSLV Payload Firing is 3.2 meters in diameter, 8.3 meters long and weighs 1,150 Kilograms. It consists oftwo all-aluminium halves that consist of different section. On top, the firing has a spherical nose cone and aconical section at the forward end. A long cylindrical and a short conical boat-tail follow. The conicalsections are stiffened semi-monocoque structures and the cylindrical section is and integrally stiffened

isogrid structure made up of three 1.5-meter panels.[TABLE 4.5]

The payload firing features a controlled environment and Payload purge supply as well as acousticabsorption blankets. Additionally, the firing can be outfitted with RF windows.

The firing is separated about 165 seconds into the flight at an altitude of 130 Kilometres. Separation isaccomplished by a linear piston cylinder separation and jettisoning mechanism (zip cord) running along thefull length of the PLF and a clamp and joint at the base of the firing. Both systems are pyrotechnicallyinitiated. The gas pressure generated by the zip cord expands a rubber bellow that pushes that piston andcylinder apart, pushing the firing halves laterally away from the launcher.[Fig.4.6]

Fig-4.6-Payload

Table 4.5-Specifications of Payload Firing

Diameter 3.2m

Length 8.3m


34/51

34

Mass 1,150kg

Separation Piston Cylinder Mechanism

Clamp Band

ConstructionAluminum Alloy, Semi Monocoque

Spherical Nose Cap, Forward

Conical Section, Cylindrical Section,

Conical Boat-tail

4.2 PSLV-XLC25

In case of mangalyaan , extended model PSLV-XL is used. PSLV-XL version is boosted by more powerful,

stretched strap on boosters and allows putting additional 200 kg payload (total 1,800 kg) in space.. The firstversion of PSLV-XL had successfully launched Indias first Moon mission called Chandrayaan-1.The

success of Chandrayaan-1 has demonstrated that ISROs existing launch vehicle technology with some

additional modifications is capable of undertaking an unmanned mission to a celestial body. ISRO has the

capability of launching spacecraft to Mars with the existing Polar Satellite Launch Vehicle (PSLV-XL

c25),[Fig-4.7]which is a proven launcher technology. To launch a spacecraft to Mars, in general, two major

options could be considered: one a fly-by mission, and second, the orbiter missions. Comparatively, the

complicities are more with the orbiter missions. However, an orbiter, in an orbit of very long time period,

offers the opportunity to study Mars for a considerably longer period as compared to a spacecraft in a fly-by

mission. Hence, ISRO went in the favour of planning Mars orbiter capability. It needs to be noted that the

PSLV has few limitations to undertake missions like Mars mission particularly it has very limited weightweight carrying capability. For November 2013, Mars mission launch the ISRO would be using PSLV-XL, a

novel mission (C-25) design carried out, wherein a spacecraft mass of 1,350 kg would be placed in an

elliptical orbit of 370 km by 80,000 km. This means that the mission would go round the Mars in an

elliptical path closest at 370 km and farthest at 80,000 km. Total five different sensors with a combined

payload of 15 kg would be undertaking various observations.


35/51

35

Fig.4.7-The complete assembly of PSLV XL-C25 launch vehicle

4.3 WHY PSLV-XL IS PREFERED IN MANGALYAAN INSTEAD OF GSLV

*The main reason behind the advent of the GSLV is the capability to lift greater loads into space. While the

PSLV can only lift slightly over a ton of payload to GTO (Geostationary Transfer Orbit), the GSLV is

capable of lifting more than double that with a rated capacity of 2 to 2.5 tons. One of the main reasons whythe GSLV has such an increased load is its utilization of a cryogenic rocket engine for its last stage. The

cryogenic rocket engine provides more thrust than conventional liquid rocket engines but the fuel and

oxidizer needs to be super cooled in order to keep them in a liquid state.

*In order to reach geosynchronous orbit, the GSLV needs more power than the PSLV could supply.

*The GSLV uses cryogenic engine, but PSLV uses Vikas engine, GSLV is used only for launching large

satellites [Table-4.6].


36/51

Table4.6-Different Stages in PSLV an

*PSLV is a launch vehicle primarily use

version of the PSLV can be used to laun

satellites to the GTO (communication sa

more powerful vehicle than the PSLV b

*The Mars Orbiter Mission spacecraft

that the PSLV-XL could put into a GTO

provide direct injection capabilities. For

velocity (11.2 km/s). If the GSLV was pa more expensive mission to launch a si

* In the case of the Mars Orbiter Missio

on course into a Trans-Martian Trajecto

36

d GSLV

d for launching satellites in the low earth orbit

ch 1 ton plus satellites in the GTO. GSLV is pri

tellites) and can inject a 2 ton satellite into orbi

t had reliability issues which are being worked

hich weighs in at a little over 1300 kg was just

. While the GSLV could carry more payload an

a direct injection to reach Mars it requires mor

roven and employed in case of Mangalyaan, itilar payload because the PSLV has a lower un

, the trajectory design is frugal. Instead of put

y( called direct injection), it uses parking orbi

r the polar orbit. A

marily for launching

. It is a larger and

out.

he heaviest payload

d maybe even

Earths escape

ould still have beenit-payload cost.

ing the craft directly

.


37/51

37

5.SPACECRAFT

The 1,337-Kilogram spacecraft carried a suite of five instruments to study Mars, its atmosphere and toacquire photos of the Red Planet. The Mars Orbiter Mission spacecraft was largely based on theChandrayaan-1 Moon Orbiter featuring the same core structure and spacecraft systems. The Mangalyaanspacecraft bus is cuboid in shape featuring composite and metallic honeycomb sandwich panels and acentral composite cylinder that facilitates all spacecraft equipment that is mounted on the panels as well asthe cylinder. The spacecraft has a dry mass of 475 Kilograms including a payload mass of 15 Kilograms andit carried a fuel load of 852 Kilograms.

Fig 5.1 Spacecraft

The spacecraft was equipped with a single deployable solar array that consists of three panels each

being 1.4 by 1.8 meters in size. The assembly also included a yoke and drive mechanism. The solar array

will provide 840 Watts of electrical power at Mars that is fed to a power distribution unit that will provide

power to the various systems and payloads and controls the state of charge of a 36-Amp-hour battery for

night passes.

5.1 Liquid Apogee Motor:

TheMain Propulsion System is centered around the Liquid Apogee Motor which has become the Indian

workhorse on Geostationary Satellites. Mangalyaan ensured that the engine can still fire after a 300-day

coast to Mars for the orbit insertion maneuver. LAM would provide 440 Newtons of thrust which equates to


38/51

38

44.87 Kilograms. The engine operates and an mixture ratio (O/F) of 1.65 and has a nozzle ratio of 160

providing a specific impulse of 3,041N*sec/ kg . The engines injector is a co-axial swirl element made of

titanium while the thrust chamber is constructed of Columbium alloy that is radiatively cooled. Electron

welding technique is used to mate the injector to the combustion chamber. LAM is a robust engine that can

tolerate injection pressures of 0.9 to 2.0 MPa, propellant temperatures of 0 to 65C, mixture ratios of 1.2 to

2.0 and bus voltages of 28 to 42 Volts. The engine is certified for long firings of up to 3,000 seconds and a

cumulative firing time of >23,542 seconds.

5.2 Propulsive Attitude Control System:

MOM was equipped with a propulsive Attitude Control System consisting of eight 22 Newton thrusters that

also uses UMDH and MON-3 propellants. The thruster also uses a co-axial swirl type .Titanium alloy

injector and a Columbium combustion chamber. The thrusters will operate in blowdown mode at a chamber

pressure of 0.68 Mpa creating a specific impulse of 2,780 N*sec/kg. The 22N thrusters have an area ratio of

100. It can be operated in pulse mode with a minimum pulse duration of 8 milliseconds that supplies a

minimum impulse of 65mN*sec. Each 22N thruster assembly weighs 0.8 Kilograms. The 22N thruster is

qualified for 300,000 duty cycles as it is mostly operated in pulse mode, but it can also withstand a single

burn of up to 10,000 seconds and a cumulative burn time of 70,000 seconds. The engine tolerates a variety

of operating conditions: 0.9 to 1.9 MPa on injection pressure, 0.2 to 2.0 on mixture ratio, -5 to 65C on prop

temperature and 28 to 42 Volts on bus voltage. In addition to a propulsive Attitude Control System, the

MOM spacecraft was equipped with four reaction wheels. Attitude and navigation data is provided by two

star trackers and gyros as well as a coarse Sun sensor with nine heads. Attitude data is also provided by an

Inertial Reference Unit and Accelerometer Package. The spacecraft features a dual redundant bus

management unit for attitude control and command processing and execution. The Attitude and Orbit

Control Electronics are centered around a MAR31750 processor.

5.3 Antenna:

Mangalyaan is equipped with a 2.2-meter diameter High Gain Antenna which is a parabolic X-Bandreflector antenna that is used for data downlink and command uplink. Science data and spacecraft telemetry

is stored in two 16Gb Solid State Recorders aboard the vehicle for downlink during regular communications

sessions. Low and Medium Gain Antennas are used for low-bandwidth communications such as command

uplink and systems telemetry downlink.


39/51

39

6. Mission Payloads

The term payload could be defined as the carrying capacity of a spacecraft and this could include cargo,

extra fuel and scientific instruments. These payloads perform the core functions of the satellite.

Over the years, the payload sensors for various satellites have been designed for specific purposes like

photography, meteorological information and reconnaissance.

Optimal selection and specification of satellite payloads is the biggest challenge for satellite mission.

Each payload is designed to perform certain functions over its useful lifetime.

6.1 Complications in selection of payloads

1. Different payload types have different levels of importance

2. Satellite payloads age and deteriorate over time due to the harsh space environment in which they operate3. All payload launch decisions are subject to various constraints form weight to budget

4. The selected payloads must be assigned certain engineering specifications to ensure their compatibilitywith the satellite bus

5. The composition of payload should match with the overall mission objective.

Payloads needs to be decided in such a way that they would help to know more about the structure of the

planet and nature of surface, meteorological conductions in and around the planet, various geological aspects

of the planet, study the surface of the planet and identifying the presence of water.

This being Indias first mission to Mars the real challenge is to reach to such a long distance and then

undertake observations. The main focus of the mission is technological aspect of the travel. The mission also

has a well-developed scientific agenda and five specifically designed payloads have been identified as the

travels for this mission. ISRO had received nearly twenty scientific payload proposals from various Centresof ISRO and department of space. The basic purpose behind all these payloads has been to address the

science of understanding the Mars atmosphere and its dynamics.

The science objectives of the proposals focus on two major aspects: one, to look at the surface features

like morphology, topography and mineralogy and second, to study the atmospheric reservoir, composition

of gas, dust, ice, clouds and their dynamics. Also, the aim is to know the interaction of atmosphere withsolar radiation and the resultant photochemistry, plasma interactions and loss processes. Most proposals

have a mass budget within 23 kg.

6.2 PayloadsBased on the technical and scientific appreciation of their mission and giving due credit to the

pronounced scientific aims of the mission, finally, five payloads have been selected by ISRO to visit Mars.

For the November 2013 mission, there would be three electro-optical payloads operating in the visibleand thermal infrared spectral ranges, a photometer to sense the Mars atmosphere and surface and a mass

spectrometer. These payloads are as follows:

1. Lyman Alpha Photometer (LAP)

2. Methane Sensor for Mars (MSM)3. Mars Exospheric Neutral Composition Analyser (MENCA)

4. Mars Color Camera (MCC)

5. Thermal Imaging Spectrometer (TIS).

The list of payloads indicating purpose, mass and objectives are summarised in the Table6.1.

ISRO has an independent Laboratory for Electro Optics Systems called LEOS. This unit of ISRO is

engaged in design, development and production of Electro-Optic sensors and camera optics for satellites and

launch vehicles. The sensors include star trackers, earth sensors, sun sensors and processing electronics.

The Space Application Centre of ISRO has provided the majority of the payloads for this mission. SAC

focuses on the design of space-borne instruments for ISRO missions and development and operationalisation

of applications of space technology for national development.SAC designs and develops all the transponders

for the INSAT and GSAT series of communication satellites and the optical and microwave sensors for IRS


40/51

40

series of remote sensing satellites.

Table 6.1 payload summary

Science

Theme

payload Primary

objectives

Mass

(kg)

Development

centre

Atmospheric

studies

LAP It measure D/H

ratio.

1.5 (LEOS)

Bangalore

MSM Measure

Methane in the

martian

atmosphere

with high

accuracy.

3.0 (SAC)

Ahmedabad

Plasma

And particle

environmental

studies

MENCA Map neutral

composition in

exosphere,

martian upper

atmosphere

4.3 (VSSC)-

Trivandrum

Surface

imaging

studies

MCC Optical colour

imaging. It will

take pictures in

red, green and

blue colours.

The camera will

help to

understand

Martian dust

storms

1.4 (SAC)

Ahmedabad

TIR Thermal remotesensing. It will

map the surface

and mineral

composition of

Mars.

4 (SAC) Ahmedabad

6.2.1 Lyman Alpha Photometer

Lyman Alpha Photometer is an absorption cell photometer. It measures the relative abundance of deuteriumand hydrogen from Lyman-alpha emission in the Martian upper atmosphere. Measurement of D/ H allows usto understand especially the loss process of water from the planet (Fig.6.1).

The objectives of this instrument are as follows:(a) Estimation of D/H ratio

(b) Estimation of escape flux of H2 corona

(c) Generation of Hydrogen and Deuterium coronal profiles.


41/51

Nominal plan to operate LAP is bet

3,000 km after Mars periapsis. Minimu

per orbit during normal range of operati

Deuterium is an isotope of hydrogen

was producedin the Big Bang. So Deute

6.

This payload is a Methane sensor. MS

ppb (parts-per-billion) accuracy and m

sensor measures reflected solar radiati

spatial and temporal variations. Hence,

less, scanning is essential. Seven Ap

planned as it scans over the Periareion i

Presence of trace amounts of methan

know more about the methane aspects

could be non-biological or biological

cometary impacts, geology and biolog

41

een the ranges of approximately 3,000 km bef

observation duration for achieving LAPs sci

n.

Fig. 6.1 View of LAP

and it is believed by many that nearly alldeute

ium is expected to be inabundance over Mars.

.2 Methane Sensor for Mars

is designed to measure Methane in the Mar

p its sources. Data is acquired only over illu

on. Methane concentration in the Martian at

global data are collected during every orbit. Si

areion Imaging scans of entire disc and Pe

every orbit (Fig.6.2).

Fig. 6.2 MSM view

on Mars has been reported by some of the N

of Mars is important in this mission. The reas

rocesses. There are three possible sources o

. If the reasons are biological, then further res

re Mars periapsis to

ence goals is 60 min

rium found in nature

ian atmosphere with

inated scene as the

osphere undergoes

ce FOV of MSM is

iareion Imaging are

SA missions. So, To

n for Mars methane

Methane on Mars;

earch to know more


42/51

42

about the presence of life on Mars would have to be undertaken.

6.2.3 Mars Exospheric Neutral Composition AnalyserMENCA(Fig.6.3) is a quadruple mass spectrometer capable of analysing the neutral composition in the

range of 1300 amu with unit mass resolution. The heritage of this payload is from Chandrayan-1 mission.

MENCA is planned to perform five observations per orbit.

Fig. 6.3 Details of MENCA

6.2.4 Mars Colour Camera

This tricolour Mars Colour Camera(Fig.6.4) gives images and information about the surface features and

composition of Martian surface. They are useful to monitor the dynamic events and weather of Mars like

dust storms/atmospheric turbidity. MCC will also be used for probing the two satellites of MarsPhobos

and Deimos.MCC would be providing the context information for other science payloads. MCC images are

to be acquired whenever MSM and TIS data is acquired. Seven Apoareion Imaging of entire disc and

multiple Periareion images of 540 9 540 km snaps are planned in every orbit.

Fig. 6.4 MCC equipment

6.2.5 Thermal Infrared Imaging Spectrometer (TIS)TIS measures the thermal emission and can be operated during both day and night. It would map surface


43/51

43

composition and mineralogy of Mars and also monitor atmospheric CO2 and turbidity. Temperature andemissivity are the two basic physical parameters estimated from thermal emission measurement. Manyminerals and soil types have characteristic spectra in TIR region. TIS can map surface composition andmineralogy of Mars (Fig.6.5).

Fig. 6.5 TIS assembly


44/51

44

7. TRAJECTORY, TRACKING AND COMMAND

7.1 THREE PHASES OF MANGALYAAN

Mangalayaan, Indias first planetary missions, journey was a 300 day, 78 crore kms to orbit Mars.

The Launch Vehicle - PSLV-C25 will inject the Spacecraft into an Elliptical Parking Orbit with a perigee of

250 km and an apogee of 23,500 km, with six Liquid Engine firing the spacecraft .It is graduallymaneuvered into a hyperbolic trajectory with which it escapes from the Earths Sphere of Influence (SOI)

and arrives at the Mars Sphere of Influence. When spacecraft reaches nearest point of Mars (Peri-apsis), it is

maneuvered in to an elliptical orbit around Mars by firing the Liquid Engine. The spacecraft then moves

around the Mars in an orbit with Peri-apsis of 366 km and Apo-apsis of about 80000 km this is shown in fig

7.1 This interplanetary mission passed through three phases namely Geo centric, Helio centric and Martian.

This is shown in Fig.7.1

Fig.7.1 TRAJECTORY DESIGN


45/51

45

7.1.1 GEO-CENTRIC PHASE

The spacecraft is injected into an Elliptic Parking Orbit by the launcher. With six main engine burns, the

spacecraft is gradually maneuvered into a departure hyperbolic trajectory with which it escapes from the

Earths Sphere of Influence (SOI) with Earths orbital velocity + V boost. The SOI of earth ends at 918347

km from the surface of the earth beyond which the perturbing force on the orbiter is mainly due to the Sun.

ISRO uses a method of travel called a Hohmann Transfer Orbit or a Minimum Energy Transfer Orbit to

send a spacecraft from Earth to Mars with the least amount of fuel possible.

First Liquid Motor Orbit Raising Burn

The first orbit raising maneuver of India's Mars Orbiter Spacecraft was successfully performed when theLAM (LIQUID APOGEE MOTOR) was fired for 416 seconds raising the spacecraft's apogee to 28825 km.The perigee is 252 km.

SECOND LIQUID MOTOR ORBIT RAISING BURN

The second orbit raising Maneuver was successfully performed with a 570.6 s burn of the liquid motor,which raised the apogee from 28814 km to 40186 km.

THIRD LIQUID MOTOR ORBIT RAISING BURN

The third orbit raising maneuver was successfully performed using a liquid motor burn for 707s IST, whichraised the apogee to from 40,186 km to 71,636 km.

FOURTH LIQUID MOTOR ORBIT RAISING BURN

The fourth orbit raising operation failed to raise the MOM's orbit to the planned 100,000 km. Theincremental velocity imparted by the liquid motor was 35 m/s against the targeted 130 m/s. As a result, theapogee increased from 71,636-km to just 78,276-km. So ISRO plans a supplementary orbit raising operationto raise the apogee to around 100,000-km

FIFTH LIQUID MOTOR ORBIT RAISING BURN

The fifth orbit raising liquid engine burn successfully for 243.5 s raised the apogee from 118642 km to192874 km.

The orbit raising maneuvers are shown in Fig. 7.2

MOM after fifth Liquid Engine burn.

Fig.7.2


46/51

46

7.1.2 Helio-Centric Phase

The spacecraft leaves Earth in a direction tangential to Earths orbit and encounters Mars tangentially to its

orbit. The flight path is roughly one half of an ellipse around sun. Eventually it will intersect the orbit of

Mars at the exact moment when Mars is there too. This trajectory becomes possible with certain allowances

when the relative position of Earth, Mars and Sun form an angle of approximately 44o. Such an arrangement

recur periodically at intervals of about 780 days. Minimum energy opportunities for Earth-Mars occur in

November 2013, January 2016, May2018 etc.

7.1.3 Martian Phase

The spacecraft arrives at the Mars Sphere of Influence (around 573473 km from the surface of Mars) in a

hyperbolic trajectory. At the time the spacecraft reaches the closest approach to Mars (Periapsis), it is

captured into planned orbit around mars by imparting V retro which is called the Mars Orbit Insertion

(MOI) maneuver. The Earth-Mars trajectory is shown in the Fig.7.3

Fig.7.3 MARTIAN ORBIT

7.2 TRACKING AND COMMAND

The Indian Space Research Organization Telemetry, Tracking and Command Networkperformed

navigation and tracking operations for the launch with ground stations at Sriharikota, PortBlair, Brunei

and Biak in Indonesia, and after the spacecraft's apogee becomes more than 100,000 km, two large 18-metre

and 32-metre diameter antennas of the Indian Deep Space Network will be utilised.NASA's Deep Space

Network will provide position data through its three stations located in Canberra, Madrid and Goldstone on

the U.S. West Coast during the non-visible period of ISRO's network. This is shown in Fig.7.4


47/51

47

Fig.7.4 TRACKING AND COMMAND

CONCLUSION:

People want their scientists to achieve stupendous successes with their missions in space.

Also, scientific achieves a significant success they do get major public support for future missions.

The questions like Why Mars? and Why Mars Now? are found being raised in recent times. Perhaps

such type of questions has no straight forward answers. For a nation-state, there could be multiple reasons

for undertaking such challenging and expensive projects. Countries could plan and invest in such missionsfor political, strategic, scientific, social and/ or economic reasons.

The intellect and consciousness of human mind are always urging it to inquire more about the

universe than what is known. In fact, the human fascination with the universe has been there for long and

would always remain. For many years, scientists are working on various aspects of the Big Bang theory a

prevailing cosmological model that describes the early explosive origin of the Universe. Humans have

always kept their interest alive with regard to finding out about the presence of life on other planets or other

solar systems and about the birth of stars and planets. The field of the study of celestial objects commonly

known as Astronomy has attracted the attention of many inquisitive and brilliant minds. It could be traced

back from the Galileo. Since then, significant developments have taken place in the field of space sciences

which essentially involves studying various known and unknown aspects of the outer space. Mainly theissues related to astronomy, planetary sciences and cosmology do get discussed under space sciences.

It is difficult to actually identify the exact reasons for the state to undertake mission to Mars. Space-faring

nations do plan to have their own dream missions in space and Mars is just a step in that journey. Planet

Mars has some peculiar attractions for which the states are dreaming to reach there. In general, the quest for

the red planet emerges out of its own logic.

Mars has an atmosphere which could provide protection from cosmic and the Suns radiation. No other

planet offers such manageable similarities with the Earth for a human stay. Also, Mars has its two Moons

namely Phobos and Deimos and understanding more about them could also help to know more about a

planet and a Moon system, in general. This could help to draw some inferences with regard to EarthMoonsystem.

Human beings are trying to find an answer to a basic query that why Earth is such a beautiful place to live

and would it always remain like that? Why the Earth is just right distance from the Sunnot too far for the

oceans to freeze or not too close for the oceans to boil? Planetary scientists call it Goldilocks Paradox. It

is argued that Earth and its two neighbouring planets Mars and Venus were formed around the same time

about 46 billion years ago, from the same ingredients including water, carbon dioxide and nitrogen. But,

only Earth developed life. Mars is too cold and Venus is too hot, but Earth is just right.


48/51

48

However, humans are yet to find the answer for such paradox. We are yet to find a theoretical backing for

why do the laws of physics seem fine-tuned for life? The Mars exploration could help us to understand about

the evolution of Earth. Probably, the study of other planet which has a matching DNA with Earth could help

us to understand more about the origins of volcanoes, earthquakes and weather. All such studies over a

period of time could directly or indirectly benefit humans to address issues related to global

warming/climate change and also forecasting of probable natural disasters.

It is important to note that climate change happens not only because of the factors originating on the Earth

itself like human-specific impacts on the surroundings, but it could also occur because of the external factorslike solar radiation received by the planet, etc. Also, scientists are trying to understand how the variations in

the Earths climate by the changes in the characteristics of the Earths orbit and axial tilt could take place.

Study of current weather changes on Mars with changes in its atmospheric composition both due to on

planet and outside planet factors could help to develop atmospheric models to understand changes in Earth

weather.

Study of the Moons of Mars could help to know more about the asteroids in particular and the formation of

the solar system in general and about the structural strength of asteroids. Maybe a Mars mission could allow

us to reach to a singular conclusion for various queries. All this would help in better understanding of Earth.

There are many unanswered questions like why Earth is the only water-rich planet, was there water available

on other planets too and if so, why did it disappear? Some answers to such question could help us to know

the future of water on the Earth. Also, mission to Mars could assist in knowing more in various arenas form

the planetary geophysics. It is known that Sun has a direct correlation with the weather on the Earth.

However, because of the particular rotation of Earth and Sun at times, it is not possible to study all the

properties of Sun form Earth. Such study could become possible from Mars.

The question is why Mars now? A simple argument could be since more than 50 years have passed as

humans have succeeded in reaching space if not now then when? Humans could be said to have started

trying to understand the secrets of Mars since 1600s with the invention of telescope. In the space era,

attempts have been made since early 1960s to study Mars by sending probes in the vicinity of Mars. This

was happening as a part of the unmanned spacecraft interplanetary exploration programme undertaken bythe erstwhile USSR and the US.

In early years, Mars was found unfriendly to Earths attempts to visit it. More missions have been attempted

to Mars than to any other place in our Solar System (except the Moon), and almost 50 % attempts have

failed. Various initial failures could have happened probably because Mars was the first planet Earth

attempted to explore. These failures have also taught us many lessons and assisted in making few subsequent

missions more successful. But, still space powers are yet to master the art of reaching Mars and some

disappointments have occurred relatively recently. Luckily, some successful missions since 1996 have

provided important data about Mars helping us to better understanding. This is helping a better planning for

future missions.

One of the questions with regard to quest for Mars is that are humans aiming for the mineral deposits on

the Mars? However, it is still premature to answer this question. No definitive information in regard to the

Mars mineralogy is available. However, some studies are available providing the regional surface material

distributions on Mars. There are indications that the soil on the Mars surface could have volcanic origin.

Also, clay minerals that usually form when water is present for long periods of time covering a larger

portion of Mars than previously thought. Mars has a different crust than Earth, and very different atmosphere

and so the minerals over there are expected to be different than that of Earth.


49/51


50/51

50

REFERENCE

Rocket propulsion by ramamurthi

Space environment and its effect on space system,aiaaa education series

Foot R, Walter A. China, the United States and Global Order, Cambridge University Press, Cambridge;2011

Anderson I. Model atmospheres show signs of life, New Scientist, Jan 7, 1988, p. 41.

Launius RD. Frontiers of Space Exploration. London: Greenwood Press; 1998. p. 56.

Sadeh E, editor. Politics of space. London: Routledge; 2011. p. 3.

Sarathi VP. Ancient Indian mathematics and Astronomy.http://www.indicstudies.us/Astronomy/aimword.pdf.

Inter-University Centre for Astronomy and Astrophysics (IUCAA). http://www.iucaa.ernet.in/Mission.html.

http://www.astron-soc.in/index.html.

http://www.prl.res.in/..

http://www.iist.ac.in/

http://mars.nasa.gov/programmissions/missions/missiontypes/orbiters/http://www.isro.org/mars.aspx

http://www.astrodigital.org/mars/whymars.html


51/51

final report on exergy analysis of power plant

Documents