Mission Objectives
One of the main objectives of the first Indian mission to Mars is to develop the technologies required for design, planning, management and operations of an interplanetary mission.
Following are the major objectives of the mission:
A. Technological Objectives:
- Design and realisation of a Mars orbiter with a capability to survive and perform Earth bound manoeuvres, 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.
B. Scientific Objectives:
Exploration of Mars surface features, morphology, mineralogy and Martian atmosphere by indigenous scientific instruments.
Major Challenges
Mangalyaan undergoing Thermal Balance Test
The bus needs to cope with a wide range of thermal environment, from Near Earth conditions with Sun and Earth contributions (hot case) to Mars conditions where eventually eclipses and reduced solar flux give rise to cold case issues.
The average solar flux at Mars orbit is 589 W/Sq.m, or about 42% of what is experienced by an Earth-orbiting spacecraft. As a result of the eccentricity of Mars orbit, however, the solar flux at Mars varies by +/- 19% over the Martian year, which is considerably more than the 3.5% variation at Earth.
Radiation Environment
ISRO Chairman inspecting work on Mangalyaan. Mangalyaan was completed in one- and-a-half years; NASA took five years for MAVEN
The main frame bus elements and payloads are basically designed for interplanetary missions capable of operating in Earth Burn Manoeuvres (EBN), Mars Transfer Trajectory (MTT) and Martian Orbit (MO) environments. The bus unit components are selected with respect to a cumulated dose of 6 krads below 22 AWG aluminium shielding. Parts have been considered as directly suitable, if they have been evaluated successfully up to 12 krads (margin factor of 2).
Communication Systems
The communication systems for the Mars mission are responsible for the challenging task of communication management up to a distance of 400 million km. It consists of Telemetry, Tracking and Commanding (TTC) systems and Data transmission systems in S-band and a Delta Differential One-way Ranging (Δ-DOR) Transmitter for ranging. The TTC system comprises of coherent TTC Transponders, TWTAs (Travelling Wave Tube Amplifiers), a near omni coverage antenna system, a High Gain Antenna system, Medium Gain Antenna and corresponding feed networks.
The High Gain Antenna system is based on a single 2.2 meter reflector illuminated by a feed at S-band.
Power System
One of the major challenges in the design of power system is due to the larger distance of the satellite from the Sun. The power generation in Mars orbit is reduced to nearly 50% to 35% compared to Earth’s orbit.
The power bus configuration comprises of a single wing of solar array with 7.56 m2 area generating about 840 W during sunlit and normal incidence in Martian orbit, and a 36 Ampere-Hour Lithium-Ion battery supports the power load during launch phase, initial attitude acquisition, eclipse, Earth burns, MOI, safe mode and data transmission phases.
Propulsion System
Propulsion System consists of one 440N Liquid Engine and 8 numbers of 22N thrusters. The propellant tanks have combined storage capacity up to 852 kg propellant. The 22N thrusters are used for attitude control during the various activities of the mission like, orbit raising using liquid engine, attitude maintenance, Martian orbit maintenance (if any) and momentum dumping.
As the critical operation of Martian Orbit Insertion with Liquid Engine burn occurs after 10 months of launch, suitable isolation techniques are adopted to prevent fuel/ oxidiser migration issues.
On-board Autonomy
Given that the Round-Trip Light Time (RLT) from Earth to Mars can vary anywhere between 6 to 43 minutes, 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, the configuration includes the use of on-board autonomy to automatically manage both the nominal and non-nominal scenarios on-board the spacecraft.
PSLV C-25 and PSLV Classic Configuration - Image: ISRO
The Launch Vehicle - PSLV-C25 injected 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 was gradually maneuvered into a hyperbolic trajectory with which it escapes from the Earth’s Sphere of Influence (SOI) and arrived 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 80,000 km.
The mission consists of following three phases:
Mangalyaan Trajectory Phases
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 Earth’s Sphere of Influence (SOI) with Earth’s orbital velocity + V boost. The SOI of earth ends at 9,18,347 km from the surface of the earth beyond which the perturbing force on the orbiter is mainly due to the Sun. One primary concern is how to get the spacecraft to Mars, on the least amount of fuel. 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.
2. Helio Centric Phase
The spacecraft leaves Earth in a direction tangential to Earth’s 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, May 2018 etc.
3. Martian Phase
The spacecraft arrives at the Mars Sphere of Influence (around 5,73,473 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) manoeuvre. The Earth-Mars trajectory is shown in the above figure. ISRO plans to launch the Mars Orbiter Mission during the November 2013 window utilizing minimum energy transfer opportunity.
Mars Orbiter Spacecraft
Mars Orbiter Spacecraft
Deployed View of Spacecraft
The spacecraft configuration is a balanced mix of design from flight proven IRS/INSAT/Chandrayaan-1 bus. Modifications required for Mars mission are in the areas of Communication, Power, Propulsion systems (mainly related to Liquid Engine restart after nearly 10 months) and on-board autonomy.
- 390 litres capacity propellant tanks accomodate a maximum of 852 kg of propellant which is adequate with sufficient margins.
- A Liquid Engine of 440 N thrust is used for orbit raising and insertion in Martian Orbit.
- The spacecraft requires three solar panels (size 1800 X 1400 mm) to compensate for the lower solar irradiance.
- Antenna System consists of Low Gain Antenna (LGA), Medium Gain Antenna (MGA), and High Gain Antenna (HGA). The High Gain Antenna system is based on a single 2.2 meter reflector illuminated by a feed at S-band. It is used to transmit/receive the Telemetry, Tracking and Commanding (TTC) and data to/from the Indian Deep Space Network
- On-board autonomy functions are incorporated as the large distance does not permit real time interventions.
Lift-off Mass
|
1,337 kg
|
Structures
|
Aluminum and Composite Fiber Reinforced Plastic (CFRP) sandwich construction-modified I-1 K Bus
|
Mechanism
|
Solar Panel Drive Mechanism (SPDM), Reflector & Solar panel deployment
|
Propulsion
|
Bi propellant system (MMH + N2O4) with additional safety and redundancy features for MOI. Proplellant
mass:852 kg
|
Thermal System
|
Passive thermal control system
|
Power System
|
Single Solar Array-1.8m X 1.4 m - 3 panels - 840 W Generation (in Martian orbit), Battery:36AH Li-ion
|
Attitude and Orbit Control System
|
AOCE (Attitude and Orbit Control Electronics): with MAR31750 Processor
Sensors: Star sensor (2Nos), Solar Panel Sun Sensor (1No), Coarse Analogue Sun Sensor
Actuators: Reaction Wheels (4Nos), Thrusters (8Nos), 440N Liquid Engine
|
Antennae
|
Low Gain Antenna (LGA), Mid Gain Antenna (MGA) and High Gain Antenna (HGA)
|
Launch Date
|
Nov 05, 2013
|
Launch Site
|
SDSC SHAR Centre, Sriharikota, India
|
Launch Vehicle
|
PSLV - C25
|
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.
Payloads
Methane Sensor for Mars (MSM)
MSM is designed to measure Methane (CH4) in the Martian atmosphere with PPB accuracy and map its sources. Data is acquired only over illuminated scene as the sensor measures reflected solar radiation. Methane concentration in the Martian atmosphere undergoes spatial and temporal variations. Hence global data is collected during every orbit.
Mars Color Camera (MCC)
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.
Lyman Alpha Photometer (LAP)
Lyman Alpha Photometer (LAP) is an absorption cell photometer. It measures the relative abundance of deuterium and hydrogen from Lyman-alpha emission in the Martian upper atmosphere (typically Exosphere and exobase). Measurement of D/H (Deuterium to Hydrogen abundance Ratio) allows us to understand especially the loss process of water from the planet.
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.
Lyman Alpha Photometer (LAP)
Lyman Alpha Photometer (LAP) is an absorption cell photometer. It measures the relative abundance of deuterium and hydrogen from Lyman-alpha emission in the Martian upper atmosphere (typically Exosphere and exobase). Measurement of D/H (Deuterium to Hydrogen abundance Ratio) allows us to understand especially the loss process of water from the planet.
The objectives of this instrument are as follows:
- Estimation of D/H ratio
- Estimation of escape flux of H2 corona
- Generation of Hydrogen and Deuterium coronal profiles
Mars Exospheric Neutral Composition Analyser (MENCA)
MENCA is a quadruple mass spectrometer capable of analysing the neutral composition in the range of 1 to 300 amu with unit mass resolution. The heritage of this payload is from Chandra’s Altitudinal Composition Explorer (CHANCE) payload aboard the Moon Impact Probe (MIP) in Chandrayan-1 mission.
Thermal Infrared Imaging Spectrometer (TIS)
TIS measure the thermal emission and can be operated during both day and night. Temperature and emissivity are the two basic physical parameters estimated from thermal emission measurement. Many minerals and soil types have characteristic spectra in TIR region. TIS can map surface composition and mineralogy of Mars.
Ground Segment
ISRO's Deep Space Network at Byalalu near Bangalore
- 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.
Mangalyaan Ground control team
- 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 Alcantara and Cuiaba TTC stations of INPE, Brazil, Hartebeestoek TTC station of SANSA and the DSN network of JPL, NASA. (Text Adapted from ISRO)
