ISRO'S Mars Orbiter Mission

06/11/2013 15:00

MARS ORBITER MISSION

 

 

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.

 

 

MISSION PLAN

 

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 is gradually maneuvered into a hyperbolic trajectory with which it escapes from the Earth’s 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.

 

The mission consists of following three 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 918347 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, May2018 etc.

 

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) 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.

 

 

SPACECRAFT INFO

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 1340 -3/+0 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

 

 

NOTE:- MORE INFO WILL BE ADDED LATER!

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