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
Mars Orbiter Mission
Mission objectives
Technological objectives:
•
Design and realisation of a Mars orbiter spacecraft with a capability to survive
and perform Earth bound manoeuvres, cruise phase, Mars orbit insertion and
capture, 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 Martian
atmosphere by indigenous scientific instruments.
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Mars Orbiter Mission
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Spacecraft On-orbit Configuration
Mars Orbiter Mission
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Mars Orbiter Mission
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Science Payloads
Mars Orbiter Mission
Lyman Alpha Photometer (LAP)
Measures the relative abundance of
deuterium and hydrogen from Lymanalpha emission in the Martian upper
atmosphere. allows us to understand
especially the loss process of water from
the planet.
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.
December 17, 2014
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.
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.
Mars Orbiter Mission
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.
4
Salient features of the Spacecraft
Mars Orbiter Mission
Features
Mass
Structures
Mechanism
Specifications
1337 kg
Aluminum and Composite Fiber Reinforced Plastic (CFRP) sandwich constructionmodified I-1 K Bus.
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)
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Mars Orbiter Mission
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Key mission elements ( Structure)
Mars Orbiter Mission
Primary Structure : 1.5m X 1.53m X 1.56m Cuboid with the central thrust
cylinder made of CFRP.
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Key mission elements ( Propulsion )
Mars Orbiter Mission
Propulsion System
• Liquid Engine to be restarted after 10
months for Martian Orbit Insertion
(MOI) manoeuvre.
• To improve safety and redundancy
additional hardware incorporated in
the configuration.
• Component heritage maintained from
INSAT/IRNSS/Chandrayaan-1
missions. 390 litres propellant tanks
accommodates a maximum of 852kg of
propellant.
• A Liquid Engine of 440N thrust to be
used for orbit raising, Trans Mars
Injection (TMI) and Martian Orbit
Insertion (MOI).
Assembling
propellant
tanks to
structure
Liquid
Engine Tests
in progress
• 8 numbers of 22N thrusters for wheel
de - saturation and attitude control
during manoeuvres.
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Key mission elements ( Power )
Mars Orbiter Mission
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).
Perihelion
Direct solar (W/m2)
Aphelion
Mean
Mars
Earth
Mars
Earth
Mars
Earth
717
1414
493
1323
589
1367.5
• The array designed for maximum performance at minimum solar flux
conditions at Mars ensuring minimum of 765 W.
Solar array wing assembly
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Mars Orbiter Mission
Battery
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Solar Panel Primary Deployment Test
Mars Orbiter Mission
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Mars Orbiter Mission
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Solar Panel Secondary Deployment Test
Mars Orbiter Mission
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Key mission elements ( Communication)
Mars Orbiter Mission
Communication system
• Communication management in Earth bound phase, cruise phase, MOI and
Martian 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 diameter
reflector 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).
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Key mission elements ( Communication)
Mars Orbiter Mission
Antenna Range and Margins
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
• 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.
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Mars Orbiter Mission
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HGA 2.2m CFRP Reflector
deployment tests
Mars Orbiter Mission
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Mars Orbiter Mission
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EMI/EMC Tests
Mars Orbiter Mission
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Key mission elements ( Sensors)
Mars Orbiter Mission
Sensors
• Star sensors: provide the inertial attitude
knowledge through identification of star
patterns in all phases of mission. Ensures
satellite pointing accuracy of +/- 0.05 deg
each axis.
• Coarse Analogue Sun Sensor (CASS) and
Solar Panel Sun Sensor (SPSS) are other
sensors providing attitude reference for sun
pointing.
• Due to the reduction in the solar irradiance
in the Martian orbit there is a large variation
in the input signals for which the pre
amplifier gain and sun presence threshold
values were optimized.
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Mars Orbiter Mission
Two star sensors integrated to
anti sun side of spacecraft
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Key mission elements ( Inertial systems)
Mars Orbiter Mission
• Accelerometers
:
For
measuring
the
precise
incremental velocity (∆V) and
for precise burn termination
• Gyros : For measuring rotation
in an inertial reference frame
Inertial Referencing
(gyros)and Accelerometer
package under Hardware Inloop Simulation
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Key mission elements ( Reaction wheels)
Mars Orbiter Mission
Reaction wheels: angular Momentum devices meant for controlling the
orientation of the spacecraft with each wheel capable of storing 6.5NMs
momentum at 4550 rpm and generate a maximum torque of 0.05Nm.
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Key mission elements ( AOCS)
Mars Orbiter Mission
• 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
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Key mission elements( Radiation)
Mars Orbiter Mission
Radiation Environment:
• Designed for interplanetary missions, capable of operating in Earth Burn
Manoeuvres (EBN) for the Van Allen belt crossings , Mars Transfer
Trajectory (MTT) and Martian Orbit (MO) environments.
• Components are selected with respect to a accumulated dose of 9 krads for
packages inside the cuboid and 15krads for outside mounted packages. Bus
Components 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
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Key mission elements ( Thermal)
Mars Orbiter Mission
Thermal Environment:
•
The spacecraft 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 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 is
experienced by an Earth-orbiting spacecraft.
•
As a result of the eccentricity of Mars's 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.
•
Albedo fractions are similar to Earth's, being around 0.25 to 0.28 (average).
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Loading to Large Space Simulation
Chamber for Thermovac test
Mars Orbiter Mission
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MOM Spacecraft under
Thermal Balance
Test
Mars Orbiter Mission
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Spacecraft Autonomy
Mars Orbiter Mission
• Maximum Earth to Mars Round-trip Light Time (RLT) 42 minutes during the
mission. Impractical to Micromanage the mission from Earth with ground
intervention.
• On-board autonomy is implemented thro autonomous Fault Detection,
Isolation and Reconfiguration (FDIR) logics and executed by AOCE and TMTC
packages.
• Continuous Watch, Fault Detection, Isolation and Reconfiguration without
disturbing the Earth Pointing attitude
• Essential during Communication Interruptions during eclipse, whiteouts,
blackouts.
• Safeguard the spacecraft during MOI, TWTA duty cycling and sun pointing safe
mode.
• The
mission
requirements
resulted
in
development of 22 new software modules,
modification of 42 modules and usage of 19
existing modules. These modules were extensively
tested using simulation and flight hardware in
OILS and HILS tests
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Autonomy : Definition
Mars Orbiter Mission
MOM Autonomy
• Definition
o To Monitor health of AOCS systems continuously & take corrective
actions to avoid Attitude loss
o Ensure Power Generation & communication towards Earth and
survive without ground intervention incase of any AOCS system
anomalies
•
Objective
o S/C should recover from failures automatically without ground
intervention by using all available resources and achieve Earth
pointing
o Onboard Autonomy has two major components:
o Failure Detection Isolation & Reconfiguration (FDIR) and
o Master Recovery Sequencer (MRS)
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Levels of Autonomy
Mars Orbiter Mission
• 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)
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Levels of Autonomy
Mars Orbiter Mission
• 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 logic
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MRS Context Diagram
Mars Orbiter Mission
• Master Recovery Sequencer
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Spacecraft Autonomy Realization
Mars Orbiter Mission
Communication System
Power Electronics
Actuators
Li-Ion Battery ( Voltage Levels
Check )
2 Transmitters
Sensors
4 Reaction Wheels
3 Dynamically Tuned Gyros
8 22N Attitude Control
Thrusters
2 Star Sensors
Thermal System
Heaters Temperatures
2 SPDM Motor Coils
2 Receivers
4 Accelerometers
2 TWTA
3 Antennas ( Low Gain ,
Medium Gain , High Gain)
( PATC)
Telemetry
Interface
Telemetry
Interface
TC Interface
TC Interface
Events Commands Execution for
Isolation and Reconfiguration
TC Interface
1.
2.
3.
Events Commands
Execution for Isolation
and Reconfiguration
TC Interface
Tele command Processor ( TCP-1 and TCP-2)
Isolation and Reconfiguration Commands
Execution on the reception of Events Flag ( 64)
from AOCE.
Health Check Using Telemetry and Isolation
and Reconfiguration using EBC’s (40)
Auto Thermal Control ( PATC)
December 17, 2014
1553
Packets
Transfer
Direct
Interface
1553 Interface
Events Flag Interface
TC Interface
16 TM Words ( Battery
Voltage , Current etc)
Mars Orbiter Mission
AOCE Processor ( AOCE-1 and AOCE-2)
1.
2.
3.
4.
5.
6.
Health and Performance Analysis
Fault Detection
Isolation and Reconfiguration Through Events
Raising
Internal Reconfiguration
Battery Voltage and Current Check
Safe Mode Detection and Normalization(thru
Events Flag)
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Flight SW major highlights
Mars Orbiter Mission
•
•
•
•
•
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
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S/W Development process
Mars Orbiter Mission
Req Gen 1
Req Review 1
Module
Design 1
Design
Review 1
Module Dev 1
Req Gen 2
Req Review 2
Module
Design 2
Design
Review 2
Module Dev 2
Req Gen 3
Req review 3
Module
Design 3
Design
Review 3
Module Dev 3
Module Test 3
Q A Clearance
H/W level Package
test
Integrate
Module Test 2
Module Test 1
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Spacecraft in Launch Configuration
Mars Orbiter Mission
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Mars Orbiter Mission
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Spacecraft at Launch pad
Mars Orbiter Mission
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Mars Orbiter Mission
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Launch Vehicle
Mars Orbiter Mission
•
A high performance variant of the four-stage
Polar
Satellite
Launch
Vehicle(PSLV-XL)
developed primarily for launching remote
sensing satellites of 1700kg class in SunSynchronous 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 bipropellant VIKAS engine.
•
Fourth stage employs pressure fed
performance bipropellant liquid engines.
•
Long coasting of vehicle carried out before the
separation of satellite to get the required
Argument of Perigee, one of the critical orbital
elements for mission to Mars.
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Mars Orbiter Mission
high
33
Launch
Mars Orbiter Mission
• Launched on November 05, 2013 by ISRO’s work-horse rocket PSLV in its
extended form PSLV-XL designated as PSLV-C25 Mission.
• The launch day was selected for its minimum energy transfer opportunity to
Mars, which occurs once in 26 months.
 The launch put the MOM spacecraft in 248 X 23,550 km elliptical orbit.
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Mission Profile
Mars Orbiter Mission
• Six Liquid Engine Burns take the
spacecraft gradually into a
departure hyperbolic trajectory.
S/C escapes from the Earth’s
Sphere Of Influence (SOI) with
Earth’s orbital velocity + ΔV boost.
o Spacecraft leaves Earth in a
direction tangential to Earth’s
orbit around sun. Encounters Mars
tangentially to its orbit around
sun. The flight path is roughly one
half of an ellipse around sun.
o The spacecraft arrives at Mars’
SOI in a hyperbolic trajectory.
When the spacecraft reaches Mars
Periapsis, it is captured into the
planned orbit around Mars by
imparting ΔV retro.
24-09-2014
2.5 km/s= 1.5km/s +1.0km/s
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Earth Bound Manoeuvres and Escape
Mars Orbiter Mission
• 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 the
spacecraft 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
20
Injection (TMI)
2013 11 30 19 19 24
Hyperbolic
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Trans Mars Injection
and Heliocentric Phase
Mars Orbiter Mission
• Trans Mars Injection (TMI)
MOM Spacecraft was put on the way to Mars
by performing TMI on 4th December 2013.
• Heliocentric Phase
After TMI the spacecraft entered into an elliptical
heliocentric 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.
December 17, 2014
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
Mars Orbiter Mission
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Ground Segment
(MOM Operations Complex)
Mars Orbiter Mission
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Mars Orbiter Mission
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MOM:MOI Operation
Mars Orbiter Mission
• Pre-MOI Command Uplink and operation execution plan
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Mars Orbiter Mission
TCM#4AOCS
Engines
Update MOI
Plan A or
Plan B
HTR Limits
UpdateReducing
Htr Load
23rd Sep,
2 UT-3UT UT/GDS
Coil-2 Test
22nd Sep,
13 UT/IDSN
Coil-1 Test
22nd Sep,
11 UT/IDSN-CAN
EED
Operation
Conditional
Operation
22nd Sep,
10 UT/IDSN-CAN
SS#1/D
TG#3
ON
Nominal
Operation
22nd Sep,
9UT/IDSN-CAN
TCP
Cmds
Uplink/
HTRSRAM
TEST
Acc Bias
Range 50
mg to 1
mg
/ Refinement
of MOI
Parameter
s
T0- 2 days
22nd Sep,
8UT/IDSN-CAN
T0- 5 days
19th Sep,
6UT-16UT/IDSN
14th Sep,
6UT-16UT/IDSN
T0- 3 days
15th Sep,
6UT-16UT/IDSN
Nominal
Uplinks
LEB
Cmds
Uplink
T0- 1 day
T0-9 days
21st Sep,
6UT-16UT/IDSN
T0-10 days
39
MOM:MOI Operation
Mars Orbiter Mission
• Time Line of Events
GDS
GDS + CAN
Thr Dump
Disable,LEB INIT
~24.23 min
-2.5 hrs
-3 hrs
~ 6 min
Acc Bias Init :
-1.3 hrs
5 min
TES
TBS TOS
TBM
TEE
TBE
~ 10 min
TRMS TOE
Nadir Qs Update : RB=-20 deg
Safe Mode Disable
Earth Pointing to Thrust
vector
MGA Support with 70 mt antenna
pointing
•TFMS
•TES
•TBS
•TOS
•TBM
•TEE
•TBE
• TRMS
•TOE
•TRME
MGA – 250 bps
December 17, 2014
23 min
HK-RT(MGA)+
Doppler
~ 17 min
~4.75 min
-4.5 hrs
TFMS
TM ON-HGA
~3.5 min
-21 min
-4 hrs
TM OFF
TM ON
TM OFF
: Forward Maneuver Start
: Eclipse Start ,@ TBS – 5.2 min
:Burn Start
: Occult Start ,@ TBS + 4.3 min
: Burn Mid
: Eclipse End ,@TBE – 4.75 min
: Burn End
: Rev Man Starts @ TBE +1 min
: Occult End , @TBE + 3.5 min
: Reverse Maneuver End
Back to
Earth pointing
Window Out Time : 46 mins
( 55% Thrust Underperformance
Events are referred wrt T0: Burn Start Time
Mars Orbiter Mission
TRME
HGA – 1 kbps
40
MOM Success: How it happened
Mars Orbiter Mission
• 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 Autonomy
requirements, test case reviews and test result
reviews.
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MOM Success: How it happened
Mars Orbiter Mission
• No compromise on Tests and Development
processes
• Independent V&V team
• Budget , Schedule & Logistic Management : Pre
& Post Launch
• Standing Spacecraft Authorization Board(SSAB):
Review of all mission operations, simulation
results and guided the mission team to carry out
flawless operations
• Management Support
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MOM Success: How it happened
Mars Orbiter Mission
A symbol of Team Effort
• Launch Vehicle
• Satellite
• Main Frame
• Payloads
• Integration and Testing
• Ground System
• Mission Plan & Ops
• Science Team
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Mars Orbiter Mission
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Reduction in Project cost
Mars Orbiter Mission
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%
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Anxious moments
Mars Orbiter Mission
• 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
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Mars Orbiter Mission
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Image of Mars by MOM
Mars Orbiter Mission
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Mars Orbiter Mission
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Image of Mars by MOM
Mars Orbiter Mission
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Mars Orbiter Mission
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Mars viewed by MOM
Mars Orbiter Mission
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Mars Orbiter Mission
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Mars viewed by MOM
Mars Orbiter Mission
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Mars Orbiter Mission
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Full disc of mars by MOM in three colours
Mars Orbiter Mission
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Mars Orbiter Mission
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Deimos in the orbit viewed by MOM
Mars Orbiter Mission
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Mars Orbiter Mission
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Phobos in the orbit viewed by MOM
Mars Orbiter Mission
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Mars Orbiter Mission
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The Firsts to MOM’s credit!
Mars Orbiter Mission
• Unique launch vehicle trajectory and mission concept.
• First interplanetary mission by India.
• First Indian spacecraft to escape the Sphere Of Influence of
Earth and orbit Sun.
• First mars mission in the world to succeed in TMI in first
attempt.
• First time two spacecraft from different agencies(ISRO and
NASA) reaching Mars almost simultaneously .
• Most economical interplanetary mission in the world: Realized
with a budget of INR 450Cr(USD 75 m), including launch
Vehicle, Spacecraft and Ground Segment.
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The Firsts to MOM’s credit!
Mars Orbiter Mission
• Shortest Project Schedule-Realized in a schedule of 15 Monthsfrom conceptualization to launch (space, launch & ground
segments);
• Youngest team accomplished the project in shortest time.
• First Indian spacecraft to successfully survive Van Allen belt
crossing 39 times.
• First Indian spacecraft to incorporate full scale on-board
autonomy.
• First mission to mars with spacecraft not encountering “lost in
space” conditions
• First to use Ship Borne Terminals to track the launch vehicle by
ISRO.
• First ISRO project to go online (face book/twitter) enhancing
outreach!
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Way forward
Mars Orbiter Mission
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
Test Beds : OILS and HILS
Common data base : Bus & Phases
OILS augmented with other systems like S/C
simulator/Command and TM system and
HILS
Production Driving Projects and
Technology Driving R&D
Project Driving R&D
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Acknowledgement
Mars Orbiter Mission
• Acknowledgement to all MOM team members who contributed in
preparing this material
• Sincere thanks to Mr. Arunan S, PD,MOM and Project team supported in
preparation and Dr. Annnaduri M, PRD, IRS&SSS, who gave valuable
guidance throughout.
• I thank Shri. Vasantha E, DD, CDA for his valuable guidance and
suggestions 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. Harmalkar
Subodh, for inviting and giving the great opportunity to participate in
this workshop
• I thank all the distinguished delegates and domain experts who made
me feel great by presenting about MOM
THANK YOU ALL
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