The ISRO Mars Orbiter Mission

Mars Orbiter Mission


Subramanya Udupa

Group Director, CEG,ISAC/ISRO

Email: [email protected]

FSW-2014

California Institute of Technology, Pasadena, USA

December 17, 2014


Mission Objectives

Mission objectives

Technological objectives:

• Design and realisation of a Mars orbiter spacecraft with a capability to surviveand perform Earth bound manoeuvres, cruise phase, Mars orbit insertion andcapture, and on-orbit phase around Mars.

• Deep space communication, navigation, mission planning and management.

• Incorporate autonomous features to handle contingency situations

Scientific objectives:

• Exploration of Mars surface features, morphology, mineralogy and Martianatmosphere by indigenous scientific instruments.

2Mars Orbiter MissionDecember 17, 2014


Spacecraft On-orbit Configuration



Science Payloads

4

Methane Sensor for Mars (MSM)

Designed to measure Methane (CH4) in

the Martian atmosphere with PPB

accuracy and map its sources. Global data

is collected during every orbit.

Mars Color Camera (MCC)

This tri-color camera gives images &

information about the surface features

and composition of Martian surface.

Useful to monitor the dynamic events

and weather.

Lyman Alpha Photometer (LAP)

Measures the relative abundance of

deuterium and hydrogen from Lyman-

alpha emission in the Martian upper

atmosphere. allows us to understand

especially the loss process of water from

the planet.

Mars Exospheric Neutral Composition

Analyser (MENCA)

A quadruple mass spectrometer capable

of analysing the neutral composition in

the range of 1 to 300 amu with unit mass

resolution.

Thermal Infrared Imaging Spectrometer

(TIS)

Measures thermal emission. Many

minerals and soil types have

characteristic spectra in TIR region. TIS

can map surface composition and

mineralogy of Mars.

Mars Orbiter MissionDecember 17, 2014


Salient features of the Spacecraft

5

Features SpecificationsMass 1337 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. Propellant Loading : 852 kg

Thermal System Passive Thermal Control System with heat pipe for TWTA panels.

Power System Single Solar Array-1.8m X 1.4m; 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 (SPSS)-1No, Coarse Analogue Sun Sensor (CASS)-9 Heads, and Inertial Reference Unit and Accelerometer Package (IRAP)

Actuators: Reaction Wheels (5Nms, 4Nos), Thrusters (22N-8Nos), 440N Liquid Engine

TTC Baseband and RF System

Telemetry (TM) and Telecommand (TC): CCSDS Compatible

Baseband Data Handling (BDH) and Solid State Recorder (SSR) :16+16 Gb

Communication (RF) Systems:

S-Band for both TTC and Data

Antennae: Low Gain Antenna (LGA), Mid Gain Antenna (MGA) and High Gain Antenna (HGA)



Key mission elements ( Structure)

6

Primary Structure : 1.5m X 1.53m X 1.56m Cuboid with the central thrustcylinder made of CFRP.



Propulsion System

• Liquid Engine to be restarted after 10months for Martian Orbit Insertion(MOI) manoeuvre.

• To improve safety and redundancyadditional hardware incorporated inthe configuration.

• Component heritage maintained fromINSAT/IRNSS/Chandrayaan-1missions. 390 litres propellant tanksaccommodates a maximum of 852kg ofpropellant.

• A Liquid Engine of 440N thrust to beused for orbit raising, Trans MarsInjection (TMI) and Martian OrbitInsertion (MOI).

• 8 numbers of 22N thrusters for wheelde - saturation and attitude controlduring manoeuvres.

7

Key mission elements ( Propulsion )


Assembling propellant

tanks to structure

Liquid Engine Tests in progress


Power Systems

• Reduction in Solar power generation by a factor of 1/3 at Martian orbit.

• Very low temperature of solar panels during eclipse periods (-185oC).

8

Perihelion Aphelion Mean

Mars Earth Mars Earth Mars Earth

Direct solar (W/m2) 717 1414 493 1323 589 1367.5

• The array designed for maximum performance at minimum solar fluxconditions at Mars ensuring minimum of 765 W.

Solar array wing assembly Battery

Key mission elements ( Power )



Solar Panel Primary Deployment Test



Solar Panel Secondary Deployment Test



Communication system

• Communication management in Earth bound phase, cruise phase, MOI andMartian orbit phase.

• The maximum range is 375 million Km ( Range of Mars after 6 months)

• S-Band deep space frequencies for both TTC and Payload Data transfer

• Two coherent transponders with two 230 W TWTAs and 2.2 meter diameterreflector antenna to increase on board EIRP

• Sensitive receiver with -135 dBm carrier acquisition threshold with“Sequential ranging “ compatibility (500 KHz ranging tone).

• Integrated Doppler provision is provided .

• Selectable data rates of 5/10/20/40 kbps (without turbo coding).

11

Key mission elements ( Communication)



12

Antenna Beam

width

Peak Gain Polarization TM Support

(km)/margin (dB)

TC Support (km)/

margin (dB)

LGA ± 90° 0dB LCP&RCP 1.4 million /2.4 dB 30 million /2 dB

MGA ± 40° 7dB;

3dB @ ±40°

RCP 40 million /2.4 dB 110 million /2.3 dB

HGA ± 2° 31dB RCP 400 million /5.4 dB 400 million /10 dB

Key mission elements ( Communication)

• Satellite Recovery is with MGA with 70 m ground antenna.

• High Antenna Gain (HGA) caters to the communication requirement for full mission. Range up to 400 million Km with 32 m ground antenna.

Antenna Range and Margins



HGA 2.2m CFRP Reflector deployment tests



EMI/EMC Tests



Sensors

• Star sensors: provide the inertial attitudeknowledge through identification of starpatterns in all phases of mission. Ensuressatellite pointing accuracy of +/- 0.05 degeach axis.

• Coarse Analogue Sun Sensor (CASS) andSolar Panel Sun Sensor (SPSS) are othersensors providing attitude reference for sunpointing.

• Due to the reduction in the solar irradiancein the Martian orbit there is a large variationin the input signals for which the preamplifier gain and sun presence thresholdvalues were optimized.

15

Key mission elements ( Sensors)

Two star sensors integrated to anti sun side of spacecraft



16

• Accelerometers : Formeasuring the preciseincremental velocity (∆V) andfor precise burn termination

• Gyros : For measuring rotationin an inertial reference frame

Inertial Referencing (gyros)and Accelerometer

package under Hardware In-loop Simulation

Key mission elements ( Inertial systems)



17

Reaction wheels: angular Momentum devices meant for controlling theorientation of the spacecraft with each wheel capable of storing 6.5NMsmomentum at 4550 rpm and generate a maximum torque of 0.05Nm.

Key mission elements ( Reaction wheels)



18

Key mission elements ( AOCS)


• Initial Acquisition

• Sun Acquisition

• Earth Acquisition

• Inertial Attitude Hold

• On Orbit Mode ( Normal Modes)

• Orbit Maneuver operations

• LEB operations

• Momentum Dumping

• Recovery from Attitude Loss

• Attitude Profile Generation

• Maneuvers for Imaging

• Imaging Attitude Reference generation and Control

• Orbit Computation for current and Future Requirements

• Solar Panel Tracking

• Automatic Sequencers and Safety Logics


Key mission elements( Radiation)

Radiation Environment:

• Designed for interplanetary missions, capable of operating in Earth BurnManoeuvres (EBN) for the Van Allen belt crossings , Mars TransferTrajectory (MTT) and Martian Orbit (MO) environments.

• Components are selected with respect to a accumulated dose of 9 krads forpackages inside the cuboid and 15krads for outside mounted packages. BusComponents are latch up immune with a Minimum threshold value LET:

• For Single Event Upset (SEU) LET > 40 Mev.cm².mg-1

• For Single Event Latch ups (SEL) LET > 80 Mev.cm².mg-1



20

Thermal Environment:

• The spacecraft needs to cope with a wide range of thermal environment, fromNear Earth conditions with Sun and Earth contributions (hot case) to Marsconditions where reduced solar flux and longer eclipses give rise to cold case.

• The average solar flux at Mars orbit is 589 W/m2, or about 42% of what isexperienced by an Earth-orbiting spacecraft.

• As a result of the eccentricity of Mars's orbit, however, the solar flux at Marsvaries by +/- 19% over the Martian year, which is considerably more than the3.5% variation at Earth.

• Albedo fractions are similar to Earth's, being around 0.25 to 0.28 (average).

Key mission elements ( Thermal)



Loading to Large Space Simulation Chamber for Thermovac test



MOM Spacecraft under Thermal Balance Test



• Maximum Earth to Mars Round-trip Light Time (RLT) 42 minutes during themission. Impractical to Micromanage the mission from Earth with groundintervention.

• On-board autonomy is implemented thro autonomous Fault Detection,Isolation and Reconfiguration (FDIR) logics and executed by AOCE and TMTCpackages.

• Continuous Watch, Fault Detection, Isolation and Reconfiguration withoutdisturbing the Earth Pointing attitude

• Essential during Communication Interruptions during eclipse, whiteouts,blackouts.

• Safeguard the spacecraft during MOI, TWTA duty cycling and sun pointing safemode.

23

• The mission requirements resulted indevelopment of 22 new software modules,modification of 42 modules and usage of 19existing modules. These modules were extensivelytested using simulation and flight hardware inOILS and HILS tests

Spacecraft Autonomy



MOM Autonomy

• Definition

o To Monitor health of AOCS systems continuously & take correctiveactions to avoid Attitude loss

o Ensure Power Generation & communication towards Earth andsurvive without ground intervention incase of any AOCS systemanomalies

• Objective

o S/C should recover from failures automatically without groundintervention by using all available resources and achieve Earthpointing

o Onboard Autonomy has two major components:

o Failure Detection Isolation & Reconfiguration (FDIR) and

o Master Recovery Sequencer (MRS)

24

Autonomy : Definition



Levels of Autonomy

• Level A Autonomy

o Fault Detection and Isolation(FDI) logic: WDT Based

o Long Pulse Detection(LPD) logic: Thruster Driver stuck high failure detection and Isolation

o EDAC and Memory Scrubbing

o Mil 1553B Bus Change over Logic

o Remote Programming

o AOCE Reset Handling

o E2PROM Management

• Level B Autonomy

o Fault Detection, Isolation and Reconfiguration (FDIR)

Mars Orbiter Mission 25December 17, 2014


Levels of Autonomy

• Level C Autonomy

o Power Safety Logic

o AOCS Safe Mode

o Master Recovery Sequencer (MRS): To handle more than one failure and recovery & Safe Mode recovery

• Level D Autonomy

o Sequencers

Launch phase sequencer

LEB Sequencer

Thruster Augmentation Logic

LEB termination logic

Payload operation Sequencer

o Other Logics

SS occultation handling

Over speed protection logic

Reference check logicMars Orbiter Mission 26December 17, 2014


MRS Context Diagram

• Master Recovery Sequencer



1553 Interface

Events Flag Interface

Tele command Processor ( TCP-1 and TCP-2)

1. Isolation and Reconfiguration Commands

Execution on the reception of Events Flag ( 64)

from AOCE.

2. Health Check Using Telemetry and Isolation

and Reconfiguration using EBC’s (40)

3. Auto Thermal Control ( PATC)

AOCE Processor ( AOCE-1 and AOCE-2)

1. Health and Performance Analysis

2. Fault Detection

3. Isolation and Reconfiguration Through Events

Raising

4. Internal Reconfiguration

5. Battery Voltage and Current Check

6. Safe Mode Detection and Normalization(thru

Events Flag)

Sensors

3 Dynamically Tuned Gyros

2 Star Sensors

4 Accelerometers

Actuators

4 Reaction Wheels

8 22N Attitude Control

Thrusters

2 SPDM Motor Coils

Power Electronics

Li-Ion Battery ( Voltage Levels

Check )

Thermal System

Heaters Temperatures

( PATC)

Communication System

2 Transmitters

2 Receivers

2 TWTA

3 Antennas ( Low Gain ,

Medium Gain , High Gain)

1553

Packets

Transfer

Direct

InterfaceTelemetryInterface

TC Interface

TC InterfaceTC Interface

TC Interface

Events Commands Execution for

Isolation and Reconfiguration

TC Interface

Events Commands

Execution for Isolation

and Reconfiguration

16 TM Words ( Battery

Voltage , Current etc)

TelemetryInterface

Spacecraft Autonomy Realization



Flight SW major highlights

• Developed using ADA for a Mil-Std-1750 processor.

• Sensor processing and Actuator control.

• Autonomy• Onboard orbit models for Heliocentric and Martian Phases

• Onboard attitude steering for all Mission phases including LEBs

• Model based as well as coefficients based

• Ground Orbit determination system considering range, doppler

and DDOR measurements for various phases of the Mission

• Precise Measurement Modeling

• Ground Orbit computation using accelerometer data for Earth

Burns and MOI



S/W Development process

Req Gen 1

Req Gen 2

Req Gen 3

Req Review 1Module Design 1

Design Review 1

Module Dev 1

Module Test 3

Req Review 2Module Design 2

Design Review 2

Module Dev 2

Module Test 2

Req review 3Module Design 3

Design Review 3

Module Dev 3

Module Test 1

IntegrateH/W level Package test

Q A Clearance



Spacecraft in Launch Configuration



Spacecraft at Launch pad



Launch Vehicle

• A high performance variant of the four-stagePolar Satellite Launch Vehicle(PSLV-XL)developed primarily for launching remotesensing satellites of 1700kg class in Sun-Synchronous Polar Orbit.

• 3.5 meter diameter and 44 meters tall .

• Lift-off mass of 320 ton.

• The first and third stages use solid motors.

• The second stage employs turbo pump fed bi-propellant VIKAS engine.

• Fourth stage employs pressure fed highperformance bipropellant liquid engines.

• Long coasting of vehicle carried out before theseparation of satellite to get the requiredArgument of Perigee, one of the critical orbitalelements for mission to Mars.



Launch

34

• Launched on November 05, 2013 by ISRO’s work-horse rocket PSLV in itsextended form PSLV-XL designated as PSLV-C25 Mission.

• The launch day was selected for its minimum energy transfer opportunity toMars, which occurs once in 26 months.

The launch put the MOM spacecraft in 248 X 23,550 km elliptical orbit.



Mission Profile

• Six Liquid Engine Burns take thespacecraft gradually into adeparture hyperbolic trajectory.S/C escapes from the Earth’sSphere Of Influence (SOI) withEarth’s orbital velocity + ΔV boost.

o Spacecraft leaves Earth in adirection tangential to Earth’sorbit around sun. Encounters Marstangentially to its orbit aroundsun. The flight path is roughly onehalf of an ellipse around sun.

o The spacecraft arrives at Mars’SOI in a hyperbolic trajectory.When the spacecraft reaches MarsPeriapsis, it is captured into theplanned orbit around Mars byimparting ΔV retro.

35

24-09-2014

2.5 km/s= 1.5km/s +1.0km/s



Earth Bound Manoeuvres and Escape

36

• Geocentric Phase• In total, there were about 6 Earth Bound Maneuvers carried out, spread

over 19 orbits. These Earth Bound Maneuvers were meant to increase thespacecraft velocity at perigee (nearest orbital point to the earth).

Maneuver Perigee No. Maneuver start time (U T) Apogee Achieved

(km)

Launch 1 2013 11 05 09 52 43 23550

EBN#1 6 2013 11 06 19 47 28 28826

EBN#2 9 2013 11 07 20 48 52 40183

EBN#3 11 2013 11 08 20 40 43 71623

EBN#4 13 2013 11 10 20 37 52 78708

EBN#4A 14 2013 11 11 23 33 50 118632

EBN#5 16 2013 11 15 19 57 17 192918

Trans Mars

Injection (TMI) 20 2013 11 30 19 19 24 Hyperbolic



37

• Trans Mars Injection (TMI)

MOM Spacecraft was put on the way to Marsby performing TMI on 4th December 2013.

• Heliocentric PhaseAfter TMI the spacecraft entered into an ellipticalheliocentric orbit (Hohmann Transfer Orbit)designed to intersect Mars orbit around sun.

• Trajectory Correction Maneuvers (TCM)

MOM was subjected to two TCMs to correct the effects of “small forces “, AMDs and the errors in navigation solution propogation during the heliocentric phase on 11th December 2013 and 11th June 2014 by imparting a total Δ Velocity of 9.33m/s.

TCM No. Maneuver start time (U T) Burn Duration (s) ΔV Imparted (m/s)

1 2013 12 11 00 59 40 40.3 7.74

2 2014 06 11 11 00 00 15.9 1.59

Trans Mars Injection and Heliocentric Phase



Ground Segment (MOM Operations Complex)



MOM:MOI Operation

• Pre-MOI Command Uplink and operation execution plan

Mars Orbiter Mission 39

T0- 3 daysT0- 2 days

LEB Cmds Uplink

TCP Cmds Uplink/HTR-SRAM TEST

SS#1/DTG#3 ON

Acc Bias Range 50 mg to 1 mg/ Re-finementof MOI Parameters

EED Operation

Coil-1 Test Coil-2 TestTCM#4-AOCS Engines

Update MOI Plan A or Plan B

HTR Limits Update-Reducing Htr Load

14

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ConditionalOperation

Nominal Operation

Nominal Uplinks

T0-10 days

T0- 5 days

T0- 1 dayT0-9 days

December 17, 2014


MOM:MOI Operation

• Time Line of Events


-4.5 hrs

TM ON-HGA

Back to Earth pointing

MGA – 250 bps HGA – 1 kbps

•TFMS : Forward Maneuver Start •TES : Eclipse Start ,@ TBS – 5.2 min•TBS :Burn Start •TOS : Occult Start ,@ TBS + 4.3 min•TBM : Burn Mid•TEE : Eclipse End ,@TBE – 4.75 min•TBE : Burn End • TRMS : Rev Man Starts @ TBE +1 min•TOE : Occult End , @TBE + 3.5 min•TRME : Reverse Maneuver End

HK-RT(MGA)+ Doppler

TBS TBMTBETOS TOETES TEETFMS

~24.23 min

~4.75 min

Earth Pointing to Thrust vector pointing

TRME

MGA Support with 70 mt antenna

-21 min

~ 10 min

~ 6 min

-3 hrs~3.5 min

TM OFFTM ON23 min

Window Out Time : 46 mins ( 55% Thrust Underperformance

TRMS

5 min

TM OFF

~ 17 min

Safe

Mo

de

Dis

able

Nad

ir Q

s U

pd

ate

: RB

=-2

0 d

eg

-4 hrs

GDS GDS + CAN Thr Dump Disable,LEB INIT

-2.5 hrs

Events are referred wrt T0: Burn Start Time

Acc Bias Init : -1.3 hrs

December 17, 2014


MOM Success: How it happened

• Quick & Firm Decisions• Finalization of the S/C configuration :

• Chandrayaan-2 : AOCE, TM, TC, Power, Inertial systems

- H/W diversion Vs Procurement Vs New

• Midcourse Corrections• Fixed Solar panel to Rotating panel

• S- Band alone & Modification of Existing TWTAs

• RX – TC cross strap removal

• DDOR package inclusion.

• Standing review committees for Autonomyrequirements, test case reviews and test resultreviews.




• No compromise on Tests and Developmentprocesses

• Independent V&V team

• Budget , Schedule & Logistic Management : Pre& Post Launch

• Standing Spacecraft Authorization Board(SSAB):Review of all mission operations, simulationresults and guided the mission team to carry outflawless operations

• Management Support




A symbol of Team Effort

• Launch Vehicle

• Satellite• Main Frame

• Payloads

• Integration and Testing

• Ground System

• Mission Plan & Ops

• Science Team

December 17, 2014



Reduction in Project cost

• ISRO designs, develops , tests and integrates S/C• No 3rd party in manufacturing

• Industry role is limited

• Manpower cost is less

• Procurement cycle and project cycle are de-linked

• Project execution time is less• Test as you develop

• Utilization of pre used modules >75%



Anxious moments

• Delivery of all systems to flight Deck on time.

• Spacecraft level thermal balance test.

• Precision launch by PSLV (polar Satellite Launch Vehicle).

• LEB operations and fuel conservation.

• Trans Mars Injection.

• Performance of LAM after long duration

• Getting into the mars orbit

• First image of mars by MOM



Image of Mars by MOM



Mars viewed by MOM



Full disc of mars by MOM in three colours



Deimos in the orbit viewed by MOM



Phobos in the orbit viewed by MOM



The Firsts to MOM’s credit!

53

• Unique launch vehicle trajectory and mission concept.

• First interplanetary mission by India.

• First Indian spacecraft to escape the Sphere Of Influence ofEarth and orbit Sun.

• First mars mission in the world to succeed in TMI in firstattempt.

• First time two spacecraft from different agencies(ISRO andNASA) reaching Mars almost simultaneously .

• Most economical interplanetary mission in the world: Realizedwith a budget of INR 450Cr(USD 75 m), including launchVehicle, Spacecraft and Ground Segment.



The Firsts to MOM’s credit!

54

• Shortest Project Schedule-Realized in a schedule of 15 Months-from conceptualization to launch (space, launch & groundsegments);

• Youngest team accomplished the project in shortest time.

• First Indian spacecraft to successfully survive Van Allen beltcrossing 39 times.

• First Indian spacecraft to incorporate full scale on-boardautonomy.

• First mission to mars with spacecraft not encountering “lost inspace” conditions

• First to use Ship Borne Terminals to track the launch vehicle byISRO.

• First ISRO project to go online (face book/twitter) enhancingoutreach!



Way forward

Pre-MOM Post MOM : Proposed

Project Specific systems - Max Standardised Systems - Max

In-house Systems - more Vendors Systems - more

Project Based Procurement - more Programmatic Procurement - more

Project Specific ILDs, Panels, Harness Standard ILDs, Panels, Harness

Made to order - Project Specific Bus system Off the self Bus systems

Made to order - S/C Off the self S/C

Meetings - more Meetings - Min

Individual data basses : Project & Phase Common data base : Bus & Phases

Test Beds : OILS and HILS OILS augmented with other systems like S/C simulator/Command and TM system and HILS

Project Driving R&D Production Driving Projects and Technology Driving R&D



Acknowledgement

• Acknowledgement to all MOM team members who contributed inpreparing this material

• Sincere thanks toMr. Arunan S, PD,MOM and Project team supported inpreparation and Dr. Annnaduri M, PRD, IRS&SSS, who gave valuableguidance throughout.

• I thank Shri. Vasantha E, DD, CDA for his valuable guidance andsuggestions in preparing this material.

• I thank Dr. Shivakumar S K, Director, ISAC and Dr. K. Radhakrishnan,Chairman ISRO for providing opportunity to attend this Work Shop.

• I thank Dr. Alan D. Unell, Chaiman,FSW-2014 and Mr. HarmalkarSubodh, for inviting and giving the great opportunity to participate inthis workshop

• I thank all the distinguished delegates and domain experts who mademe feel great by presenting about MOM

THANK YOU ALL



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December 17, 2014

The ISRO Mars Orbiter Mission

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