GREEN TEAM Summer School Alpbach 2008 July 30, 2008 Table of - - PowerPoint PPT Presentation

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GREEN TEAM Summer School Alpbach 2008 July 30, 2008 Table of - - PowerPoint PPT Presentation

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Final Presentation GREEN TEAM Summer School Alpbach 2008 July 30, 2008 Table of Contents Table of Contents 1. Science Case Scientific Background Mission Objectives Why Sample Return Scientific Competitivness Mission

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Final Presentation

GREEN TEAM

Summer School Alpbach 2008 – July 30, 2008

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GREEN TEAM – SUMMER SCHOOL ALPBACH 2008 2

Table of Contents Table of Contents

  • 1. Science Case
  • Scientific Background
  • Mission Objectives
  • Why Sample Return
  • Scientific Competitivness
  • Mission Target Selection
  • 2. Mission Requirements
  • 3. Mission Design
  • 4. Management Aspects
  • 5. Conclusion
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GREEN TEAM – SUMMER SCHOOL ALPBACH 2008 3

Solar System Formation I Solar System Formation I

Collapse of Solar Nebula CAI Formation 4.566 ± 0.002 Gyr Small Planetary Bodies accreted

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Solar System Formation II Solar System Formation II

Alteration (petrolo Alteration (petrology) in gy) index: dex: 3 = lowest grade of thermal metamorphism 6 = highest grade of thermal metamorphism

Ordinary Chondrites

Mildthermal Metamorphism Material (~10 Myr)

≥ 50km

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Solar System Formation III Solar System Formation III

Planetesimals differentiated

Production of melted material (2-3 Myr)

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Early Solar System Timeline Early Solar System Timeline

CAIs Differentiation Thermal Metamorphism (OC) Time (Myr) 3 10 Earth Mars 15 30

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Formation of Planets Formation of Planets

Kleine et al. 2002

Runaway growth of planets

Earth in 30-40 Myr

Planetesimals represent the building blocks of the planets

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GREEN TEAM – SUMMER SCHOOL ALPBACH 2008 8

Missions to Asteroids Missions to Asteroids

NASA NASA NASA NASA ESA NASA JAXA (243) Ida, Galileo, 1993 (5535) Annefrank, Stardust, 2002 (25143) Itokawa, Hayabusa,2005 SR SR (4) Vesta, Dawn, 2012 (433) Eros, NEAR, 1998 (9969) Braille, Deep Space 1, 1999 (2867) Steins, Rosetta, 2008

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GREEN TEAM – SUMMER SCHOOL ALPBACH 2008 9

Main questions to be answered Main questions to be answered

What are the building blocks of the terrestrial planets? What are the timescales and nature of differentiation in planetesimals? How are the basaltic asteroids genetically linked to differentiated meteorites? What is the formation history of the basaltic NEOs? What is the composition and mineralogy of Q-type asteroids ? Are they linked to S-type?

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Meteorites Meteorites

Ordinary Ordinary Chondrites Chondrites H, L, LL Differenti Differentiated ated HED Angrites Pallasites Mesosiderites

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GREEN TEAM – SUMMER SCHOOL ALPBACH 2008 11

Choice Choice of the Targets f the Targets

Criteri Criteria NEO Accessibility Taxonomic Type Physical Parameters Final Choi Final Choices ces Sample Object: a V-type NEO (5604) 1992 FE

  • Rot. period: 6.026h

First Flyby: a Q-type NEO (152560) 1991 BN

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GREEN TEAM – SUMMER SCHOOL ALPBACH 2008 12

Differentiated Differentiated Meteorites eteorites

  • 0,2
  • 0,1

0,1 0,2 0,3 3 4 5 6 Eucrites Diogenites Mesosiderites

Mars Moon Angrites Ibitira Pallasites TFL

δ O Δ O

18 17

Franchi et al. (2000), Wiechert et al. (2000,2004) and Greenwood et al. (2005, 2006)

Asteroids: M , A, V Types

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Mission Statements Mission Statements

Return a subsurface sample from a V-type NEO after characterising the body from orbit Rendezvous with a fragment of a Q-type asteroid

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Why Why Sample ample Return? eturn?

Meteorite analysis is limited by lack of geological context The detailed high precision and resolution measurements needed to answer many important questions cannot be performed in situ: Radiometric dating Imaging and petrology on the mm to nm scale Trace element analysis Isotopic analysis

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Key Key Analyses Analyses for for the he Sample Sample

Informa Information ion Example Examples of T

  • f Techniques

echniques Why not in situ Why not in situ Kind of Kind of An Analyses alyses Imaging (mm to nm) Mineralogy Chemical composition Atomic structure Radiometric dating Oxygen isotopes Optical microscopy Electron microscopy Ion microprobe XANES Laser-fluorination mass-spectrometry MC-ICP-MS Sample preparation (chemical and physical) high energy source constraints large magnetic sector geometries needed Most are destructive

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First sampling mission to target a differentiated asteroid Will visit a class (Q-type) of NEO that has not yet been visited by a space mission How does it complement existing and planned missions? Hayabusa and Marco Polo primitive asteroids Dawn Vesta, Ceres but no sample return

Scientific Competitiveness Scientific Competitiveness

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Scientific Payload Scientific Payload

Instrument Instrument Mass (g) Mass (g) Power

  • wer (W)

(W) TRL TRL Da Data Rate ta Rate Herita Heritage Fr Frame ame C Camer mera 5000 12 9 0.9 Mb/s DAWN High Resol. Ca High Resol. Cam. m. 2400 6.5 4-9 100 kb/s Bepi Colombo Close Close Up Cam Up Cam 600 4 4-9 14 Mb/s Marco Polo 2 * Wide An 2 * Wide Angle Cam gle Cam 2 * 350 2 * 4 4-9 100 kb/s Marco Polo Laser Alt Laser Altimeter meter 4000 17 9 15 kb/s Hayabusa Visible + IR Spectr. Visible + IR Spectr. 9300 17.6 9 5 - 20 kb/s DAWN X Ra X Ray Spectr. y Spectr. 3000 18 9 7.5 kb/s SMART 1 Radio Sci Radio Science ence

  • 8
  • Mars Express

Σ 25000 g 83.1 W

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Payload I Payload I

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Payload II Payload II

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1. Return a subsurface sample from a V-type NEO that can be analysed by a number of high resolution and precision techniques on Earth 2. Provide geological context for the returned sample 3. Remote sensing of a Q-type NEO before sampling the target to understand the diversity of differentiated asteroidal material

Scientific Drivers I Scientific Drivers I

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From orbit Imaging of surface features to 1m Macroscale mineralogy and composition (to 20m resolution) Size, shape, mass, density, gravity of asteroidal body Sample site Imaging of the sampling site (FOV ~ 1millirad) Chemical composition and mineralogy (FOV <1millirad) Sample of 150g and at least below 3cm

Scientific Drivers II Scientific Drivers II

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Planetary Protection Planetary Protection

Based on COSPAR Planetary Protection Policy

Category II: impact probability (1) and contamination control measures. Category V Unrestricted Earth return

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Scientific requirements: no contamination by terrestrial material no greater temperature ranges than sample experienced on parent body No planetary protection requirements

Sample Return Container Sample Return Container

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Table of Contents Table of Contents

  • 1. Science Case
  • 2. Mission Requirements
  • 3. Mission Design
  • Mission Scenario
  • Orbit and Trajectories
  • Launch Segment
  • Design of Space Segment
  • Ground Segment
  • Key Technology
  • 4. Management Aspects
  • 5. Conclusion
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S/C Configuration Trade-Off I S/C Configuration Trade-Off I

Case Case 3 3

Case 3 Case 3 Case 2 Case 2

Case 1 Case 2 Case 3 Points 3

  • 11

9

Case 1 Case 1

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S/C Configuration Trade-Off II S/C Configuration Trade-Off II

Hexagonal shape Solar panels configurations Radiators configuration Landing gears Sampling strategy, sample mechanism

  • Config. 1
  • Config. 2
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Mission Scenario I Mission Scenario I

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Mission Scenario II Mission Scenario II

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Mission Scenario III Mission Scenario III

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Mission Scenario IV Mission Scenario IV

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Mission Scenario V Mission Scenario V

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Mission Scenario VI Mission Scenario VI

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Mission Scenario VII Mission Scenario VII

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Mission Scenario VIII Mission Scenario VIII

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Mission Scenario IX Mission Scenario IX

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Mission Scenario X Mission Scenario X

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Mission Scenario XI Mission Scenario XI

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Mission Scenario XII Mission Scenario XII

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Transfer Trajectories I Transfer Trajectories I

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Transfer Trajectories II Transfer Trajectories II

Launch Date June 7th 2018 Stay Time at 1991 BN 1391 days Stay Time at 1992 FE 109.9 days Mission DV (nominal) 25.8 km/s Earth Return Date August 29th 2028 Mission DV (incl. Earth Escape) 29.8 km/s Earth Entry Velocity 12.3 km/s

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Earth Entry, Descent and Landing Earth Entry, Descent and Landing

Peak Peak D Deceleration

  • n Load

ads

10 20 30 40 50 60 70 80 90 8 9 10 11 12 13 14

En Entry a angle [ [°] Pe Peak d decelerati tion

  • n [

[g]

Total H l Heat at L Load ad

50 100 150 200 250 300 350 8 9 10 11 12 13 14

En Entry a angle [ [°] To Total h heat l load [ [MJ/m

2]

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Earth Entry, Descent and Landing Earth Entry, Descent and Landing

Entry mass 50 kg TPS + structure 17.3 kg 10 m parachute 12.8 kg Peak stagnation point heat flux 11.4 MW/m2 Heat shield diameter 1 m Cone angel 45 ° Nose radius 0.25 m

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Launcher: Ariane V ECA Launcher: Ariane V ECA

Dedicated Ariane V launch to C3=0 or higher Shared Ariane V launch to GTO Mass 2600 kg ~ 2830 kg Launch date June 7th 2018 Before August 3rd 2017 Total mission time 10 yrs 86 days 11 yrs 30 days

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Science Orbit Science Orbit

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  • Descent to 5000 m orbit
  • Change in configuration
  • Rehearsals

Proximity Operations Proximity Operations

Descent / Sampling / Ascent

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  • Autonomous descent

Proximity Operations Proximity Operations

Descent / Sampling / Ascent

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  • Autonomous descent

Proximity Operations Proximity Operations

Descent / Sampling / Ascent

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  • Soft landing
  • Surface thrust towards

ground

  • Harpoon anchoring
  • Sampling operations
  • Take-off to safe orbit

Proximity Operations Proximity Operations

Descent / Sampling / Ascent

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Spacecraft Structure I Spacecraft Structure I

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Spacecraft Structure II Spacecraft Structure II

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Attitude and Orbit Control System Attitude and Orbit Control System

Ring Laser Gyro (RLG) MEMS accelerometers Star trackers unit Sun sensor

Earth Autonomous Asteroid Semi‐autonomous

LIDAR Target markers Advanced Microwave Sounding Unit (AMSU)

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Attitude and Orbit Control System Attitude and Orbit Control System

Nominal: Advanced signal processing Backup: Earth-based control Ring Laser Gyro (RLG) MEMS accelerometers Star trackers unit Sun sensor

Earth Autonomous Asteroid Semi‐autonomous

LIDAR Target markers Advanced Microwave Sounding Unit (AMSU)

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Attitude and Orbit Control System Attitude and Orbit Control System

Thrusters Control moment gyros Estimated Total Mass: 110kg Estimated Power: 384W Estimated TRL: 7 Ring Laser Gyro (RLG) MEMS accelerometers Star trackers unit Sun sensor

Earth Autonomous Asteroid Semi‐autonomous

LIDAR Target markers Advanced Microwave Sounding Unit (AMSU) Nominal: Advanced signal processing Backup: Earth-based control

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Thermal Control System I Thermal Control System I

Two major problems Electrical propulsion system generates 4.2 KW heat High temperature difference between Hot Case (0.5 AU and EPS in use) and Cold Case (2.0 AU and EPS off) Special measurements to cope with mission challenges Rotation of solar arrays to reduce effective area at a distance between 0.7 AU and 0.5 AU

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Thermal Control System II Thermal Control System II

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Thermal Control System III Thermal Control System III

  • Two major problems

Electrical propulsion system generates 4.2 KW heat High temperature difference between Hot Case (0.5 AU and EPS in use) and Cold Case (2.0 AU and EPS off)

  • Special measurements to cope with mission challenges

Rotation of solar arrays to reduce effective area at a distance between 0.7 AU and 0.5 AU EPS will be fixed directly to radiators; heaters will be used in Cold Case Louvers will be used to change α/ε-ratio and variable diode heat pipes to reduce radiation heat input Estimated Mass: 60 kg Estimated Power EPS on: 5 W Esitmated TRL: 7 Estimated Power EPS off: 1650 W

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Propulsion Propulsion

Bi-propellant thrust and cold gas Bi-propellant thrust and cold gas thrust system thrust system

15x Bi-propellant Thruster Model: 10 N S10 – 01 3x Cold gas Thruster, Model: CGT1 Sterer

Regulators/Valve

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Propulsion Propulsion

Bi-propellant thrust and cold gas Bi-propellant thrust and cold gas thrust system thrust system El Electric ectrical Propulsion l Propulsion

15x Bi-propellant Thruster Model: 10 N S10 – 01 3x Cold gas Thruster, Model: CGT1 Sterer 5x Ion Propulsion, Model: RIT 22 Misp

Regulators/Valve

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Communications (TT&C) Communications (TT&C)

Estimated Downlink (Ka-Band): 50kbit/s Estimated Total Mass: 55kg Estimated Power: 45W Estimated TRL: 9 Standard Operation High Gain Antenna (HGA) No HGA Antenna due to Trajectories or Emergency 2 Low Gain Antennas (LGA) Radio Science Experiment (Data switched off) HGA 35m Deep Space Antenna

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Data Handling System Data Handling System

Two independent units based

  • n LEON 3FT and RAD6000

SpaceWire bus (up to 200 Mbits/s) 43 GB E-Disk solid-state hard drives ECSS and STD 883 equipment Main CDH Backup CDH Earth remote control Watchdog Estimated Total Mass: 10kg Estimated Power: 25W Estimated TRL: 7

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GREEN TEAM – SUMMER SCHOOL ALPBACH 2008 61 GREEN TEAM – SUMMER SCHOOL ALPBACH 2008

Power Power

AM0 90° AM0 90°,(a) ,(a) 75°,(b) ,(b) 60° 60°,(c) 45° ,(c) 45°

Lithium technology Long life: 18 year GEO at 80 % DOD Selected for Galileo, Optus D3, Alphabus Minimum specific energy 175 Wh/kg 3rd generation bare cells, 28% efficiency Effective up 2,5 AU ESA solar cell development for Mars exploration missions program Estimated Total Mass: 153kg Estimated Generated Power: 15kW Estimated Batteries Power: 4h Estimated TRL: 7

61

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ESTRACK Ground Segments ESTRACK Ground Segments

GREEN TEAM – SUMMER SCHOOL ALPBACH 2008 62

Cebreros, Spain 35m Deep Space Antenna New Norcia, Australia 35m Deep Space Antenna American Longitude Deep Space Antenna Planned (2010-2011)

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How to break the base? What if we’ll find different materials?

Piezoelectric Sampling Mechanism Piezoelectric Sampling Mechanism

SAMPLING CONCEPT

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GREEN TEAM – SUMMER SCHOOL ALPBACH 2008 64

HAMMERING ACTION

How to break the base? What if we’ll find different materials? Self-closing shape Triangular blades forming pyramid

Piezoelectric Sampling Mechanism Piezoelectric Sampling Mechanism

SAMPLING CONCEPT

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Piezoelectric Sampling Mechanism Piezoelectric Sampling Mechanism

  • Lightweight (4 kg),
  • Requires low pre-load (< 5N)
  • Driven at low power (5W)

EQUIVALENT MODEL

ULTRASONIC DRILLS

JPL’S DRILL

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Piezoelectric Sampling Mechanism Piezoelectric Sampling Mechanism

ULTRASONIC HAMMER (UH) SAMPLING PYRAMIDAL HEAD (SPH) NEW DESIGN

ROCK (BASALT)

UH

REGOLITH

ARM

ROCK (BASALT)

UH

REGOLITH

ARM

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Piezoelectric Sampling Mechanism Piezoelectric Sampling Mechanism

HANDLING BY ROBOTIC ARM

a) Deployment b) Coring c) Extraction d) Rotation e) Capsulation f) Separation and closure of the capsule

a) b) c) d) e) f)

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Anchoring System Anchoring System

Absorption of impact energy by landing gear's damping mechanism Special shape of the harpoons allowing anchoring on different type of materials Rewind system for tensioning the cable connected with the Harpoon within few seconds after the firing Compensation of the firing Impulse by thrusters

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  • 1. Science Case
  • 2. Mission Requirements
  • 3. Mission Design
  • 4. Management Aspects
  • Budgets
  • Risk
  • Timeline
  • Cost
  • 5. Conclusion

Table of Contents Table of Contents

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Budgets Budgets

Weight (kg) Power (W) TRL Payload Payload 110 110 133 133 2 Frame Camera 5 12 6 High-res Camera 2,4 6,5 6 Close-up Camera 0,6 4 6 2x Wide Angle Camera 0,7 8 6 VIR 9,3 18 6 Sampling Machine 30 50 2 Return Capsule 50 6 X-Ray 8 18 7 Laser Altimeter 4 17 8 Weight (kg) Power (W) With Margin (20%) 1445 Dry Weight 1205 Propelant 1220 Total* Total* 2665 2665 14587 14587 Weight (kg) Power (W) TRL Spacecraft Spacecraft subs subs 1094 1094 14454 14454 5 Propulsion 410 14000 6 GNC 110 384 6 Communication s 55 45 8 CDH 10 25 7 Thermal 60 1560 8 Power 149

  • 7

Structure 300 5 * @ Earth Escape

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Risk Assessment Risk Assessment

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Mission Development Timeline Mission Development Timeline

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Cost Analysis Cost Analysis

Total Mission Cost M€ 1556.04 Contingency M€ 202.10 Total Nominal Mission Cost M€ 1071.70 Payload Cost M€ 270 Platform Cost M€ 801.70

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GREEN TEAM – SUMMER SCHOOL ALPBACH 2008 74

The Dream The Dream Team thanks eam thanks you you for your

  • r your

attention! attention!

Simone Arloth, Rosalind Simone Arloth, Rosalind Arm Armytage, Uwe Derz, Laurent ytage, Uwe Derz, Laurent Donati Donati, Rene , Rene Duffard, Al Duffard, Alen en Duricic, Jessica Fla ricic, Jessica Flahaut, Stefanie Hem t, Stefanie Hempel, el, Mich ichal Ku al Kubin binyi, Ka i, Kartik rtik Kum Kumar, Simo r, Simone ne Pirro Pirrotta, An tta, Andreas P eas Pollin

  • llinger, Stine

ger, Stine Kildegaa Kildegaard Poulsen, Ma lsen, Mario Sala rio Salatti, Angelo Pio tti, Angelo Pio Rossi,

  • ssi,

Wolf Wolfga gang Seb ng Seboldt.

  • ldt.

Special thanks to all the lecturers and tutors for their precious help… FR FR3OG OG-ASM

  • ASM
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GREEN TEAM – SUMMER SCHOOL ALPBACH 2008 75

Thank you for your attention!

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Mission Architecture Mission Architecture

Weight Criteria Case 1 Case 2 Case 3 1 Mass 1 less because no double systems

  • 1

higher because double systems

  • 1

higher because double systems 1 Structure

  • 1
  • satellite has to be designed

robust enough for landing

  • Landing and touch down

system designed for higher mass

  • -> higher mass

1

  • normal standard satellite

structure

  • landing and touch down system
  • nly has to carry lander mass

1

  • normal standard satellite

structure

  • landing and touch down system
  • nly has to carry lander mass

1 Power

  • 1
  • rotation of solarpanels
  • stiff connection possible
  • dedicated solar panel design

1

  • rotation of solarpanels
  • standard solar panel design
  • power supply through teather is

possible

  • rotation of solarpanels
  • standard solar panel design
  • Lander needs own power

supply 1 Development effort 1 1 spacecraft

  • 1
  • rbiter and lander
  • 1
  • rbiter and lander

3 TRL

  • rotating system for non-

symmetric solarpanels

  • high TRL for all other parts
  • spider tested in microgravity
  • teather also tested
  • high TRL for all parts
  • rendezvous principle tested but

with GPS; no GPS at asteroid

  • high TRL for all other parts

1 Thermal

  • 1
  • complete spacecraft has to

withstand temeratures both during transfer and at asteroid 1

  • only lander has to withstand

temeratures both during transfer and at asteroid 1

  • only lander has to withstand

temeratures both during transfer and at asteroid 2 Communications

  • 1
  • no communication when

spacecraft is landed

  • orbit is restricted to teather

length

  • partly communication with earth

is possible 1

  • special communication orbit

possible

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Mission Architecture Mission Architecture

Weight Criteria Case 1 Case 2 Case 3 3 Risks 1

  • no rendezvous risk
  • same landing risk
  • no special orbit for orbiter

needed

  • same "stay on surface" risk
  • 1
  • no rendezvous risk
  • same landing risk
  • because of teather polar or

equatorial orbit is needed

  • maybe communications issues

should be considers for orbit selection

  • same "stay on surface" risk
  • rendezvous risk
  • same landing risk
  • a communication orbit to lander

is needed

  • same "stay on surface" risk

1 Pollution of surface cold gas, e.g. helium cold gas, e.g. helium cold gas, e.g. helium 2 Autonomy 1

  • simpliest one
  • 1
  • two autonomous systems
  • 1
  • most challenging
  • two autonomous systems

3 Sampling site 1

  • free choice of sampling site
  • 1
  • only polar or equatorial site is

possible

  • sice of possible landing sites

depend on teather length 1

  • free choice of sampling site

2 Anchoring and sampling mechanism 1

  • bigger mass is possible
  • direct contact to re-entry

capsule

  • 1
  • less mass is possible
  • transport to re-entry capsule
  • 1
  • less mass is possible
  • transport to re-entry capsule

2 Landing 1

  • acitive satllite landing
  • 1
  • passive lander

1

  • active lander

Weighted Summ 9

  • 11

3