MarcoPolo-R Mission and Spacecraft Design Lisa Peacocke – 19 th June 2013 IPPW 2013, San Jose,...
-
Upload
martina-burns -
Category
Documents
-
view
221 -
download
5
Transcript of MarcoPolo-R Mission and Spacecraft Design Lisa Peacocke – 19 th June 2013 IPPW 2013, San Jose,...
MarcoPolo-RMission and Spacecraft Design
Lisa Peacocke – 19th June 2013
IPPW 2013, San Jose, USA
MarcoPolo-R Mission ESA Cosmic Vison M-class candidate
Aim: To return a sample from a primitive near-Earth asteroid Currently Phase A, down-selection in Feb 2014 Target is primitive asteroid 2008 EV5
MarcoPolo-R Assessment Study and CCN Demonstrate technical and programmatic feasibility of the
mission Achieve a cost-effective and consolidated mission design Astrium team kicked off in February 2012
Team of 18 engineers currently working on the study
2
Astrium Satellites Ltd
Study Prime, Systems Engineering, Mission Analysis, Sample Handling,
Platform Subsystems, Payload Interface, AIV/Programmatics
Astrium ST S.A.S.Earth Re-Entry Capsule
Astrium ST GmbHLanding/Touchdown System
Astrium Satellites S.A.S.GNC for Proximity Operations
Mission Design Launch on Soyuz-Fregat to 2008 EV5
Launch years 2022, 2023, 2024 All outward trajectories require an Earth GAM Mission durations 4.5-6.5 years
2008 EV5 DeltaV’s relatively low Plasma propulsion architecture becomes feasible Reduced return velocities for Earth re-entry
Other benefits of 2008 EV5 Smaller asteroid => less surface area to map Less extreme orbit with more consistent Sun distance Lower mass asteroid => lower gravity environment
3
1996 FG3 Primary and Secondary
2008 EV5
30
Sun
S/C
305km altitude
60
Science Operations
4
Proximity Phases
Distance to surface (km)
Duration (days)
Arrival at Asteroid 500 0
Close Approach Trajectory 500 to 100 7
Transition to Far Global Characterisation Phase 100 to 10 10
Far Global Characterisation Phase (FGCP) 10 5
Transition to Radio Science (RSE) Phases 10 to 5 7
Gravitational RSE Phase 1 to 2 20
Bistatic Radar RSE Phase 1 to 5 TBC 5
Transition to Global Characterisation Phase 5 1
Global Characterisation Phase (GCP) 5 16
Selection of 5 sampling sites plus transition to LCP 5 7
Local Characterisation Phase (LCP) 0.25 15
Selection of best sampling sites + transition to sampling rehearsal 5 7
Sampling Rehearsals 5km to 100m 7
Transition to SAM Phase 5 7
Descent/Sampling (SAM) Phase 5km to surface 21
Transition to safe high orbit or FF position 10 7
Post Sampling Local Characterisation Phase (PSLCP) 0.25 3
Margin TBC 27
Preparation for return cruise TBC 7
Departure from Asteroid Far distance 1
Total
180
Operations phases give ~140 GB data 8 hours data downlink per day is feasible ESA’s 35 m ground stations
Spacecraft Design
5
Spacecraft Design Mechanical
Solar Orbiter derived structure, modifications to support plasma thrusters
Propulsion Three Snecma PPS1350 plasma thrusters (1.5 kW) with pointing
mechanisms and PPUs – SMART 1 Two Xenon tanks and a high pressure regulator – BepiColombo MTM Aeolus derived monopropellant system with 20N thrusters and hydrazine tanks
Thermal ‘Standard’ design with heaters and MLI, detailed analysis ongoing One panel with embedded heat pipes to aid PPU heat dissipation
AOCS Off-the-shelf European IMU, star tracker, reaction wheels and sun sensors
Electrical Two rotating solar array wings (7.5 m2 each) with drive mechanism – Sentinel 1 & 2 Lithium-ion battery and TerraSAR-X2-based 50V PCDU Mars Express 1.6 m high gain antenna; MGA and LGAs with 80 W RF TWTA and deep
space transponder – BepiColombo/Solar Orbiter/LISA Pathfinder Gaia-based on-board computer with mass memory, and Solar Orbiter RIU
6
Spacecraft Design
7
Payload Accommodation All instruments mounted on same structure panel
Facilitates integration and mutual alignment Accommodated inside spacecraft with views through cutouts
8
Key Technologies Proximity GNC
Visual navigation uses Wide Angle Camera based on NPAL development and a Radar Altimeter
Simulations performed for descent/touchdown
Sample Acquisition, Transfer and Containment Rotary brush sampling mechanism developed and tested Touchdown damping from boom back-driven motor
Minimal forces at 10 cm/s
Earth Re-entry Capsule Hard landing, no parachute or beacons/battery Hayabusa-shape aeroshell
9
Proximity GNC/AOCS
10
Touchdown Dynamics
11
12
Sampling and Transfer
13
Sampling Mechanism Early Testing
14
Sampling Mechanism
25mm
Container Volume
Cone Structure
Sampler ejected cover
Rotary Bristles PrimarySampling system
Sample Container
Earth Re-entry Capsule Main requirements
Maximum entry velocity = 12 km/s Maximum heat flux = 15 MW/m2
Maximum total pressure at stagnation point = 80 kPa Fully passive, no parachute – cost and MSR demonstration Ensure impact loads to sample are less than 800 g No beacon or battery on board Land at Woomera, Australia
Entry flight path angle of -10.8 degrees selected Based on entry dispersion and appropriate landing ellipse
Hayabusa aeroshape selected Stable and meets g-load, aerothermo requirements θc = 45 deg, RN/D = 0.5
3 June 2013 - 15
0
20
40
60
80
100
120
-16 -14 -12 -10 -8 -6
Nominal Flight Path Angle (deg)
Lan
din
g E
llip
se M
ajo
r A
xis
(km
)
DFPA=+/-0.4
DFPA=+/-0.2
DFPA=+/-0.1
DFPA=+/-0.08
Earth Re-entry Capsule Design Properties
Diameter = 0.880 m, Mass = 45.6 kg, Centring = 28.75% D (Max ~33% D) TPS: 56 mm ASTERM on frontshield (low density carbon phenolic, 280 kg/m3) 11 mm Norcoat Liége on backcover (low density cork phenolic, 470 kg/m3) 170 mm Aluminium foam crushable material, PU foam surrounds container
3 June 2013 - 16
Lid TPS Lid structure (Clip door)
Front energy absorbing material
FrontShield Structure
FrontShield TPS
Rear energy absorbing material
Internal structure
Sample
BackCover structure
BackCover TPS
Container
Margin crushable thickness
Earth Re-entry Capsule 2 rpm min spin-up
3 June 2013 - 17
Arm Drive Mechanism
ArmSpin-up & Eject Mechanisms
Mechanicalfuses
Earth Re-entry Capsule Landing ellipse is 68 km along longitudinal axis
3 June 2013 - 18
Conclusions Phase A study finishing at the end of July
Preliminary Requirements Review in Oct/Nov Selection will occur in February 2014
Astrium’s mission & spacecraft design is feasible Key technologies are well into development Extensive unit re-use or modification Keeps development costs to a minimum, reduces cost risk
MarcoPolo-R is a very promising M-class mission candidate
New target has simplified engineering & design significantly Serendipitous short mission trajectories – right time for
asteroid sample return
19
Questions?
MarcoPolo-R Team:
Steve Kemble, Héloise Scheer, Jean-Marc Bouilly, Antoine Freycon, Steve Eckersley, Brian O’Sullivan, Jaime Reed, Martin Garland, Mark Watt, Marc Chapuy, Kev Tomkins, Howard Gray, Bill Bentall, Andrew Davies, Chris Chetwood, Andy Quinn, Alex Elliott, Mark Bonnar, David Agnolon, Remy Chalex, Jens Romstedt
3 June 2013 - 20