MSR Precursor Mission M-4 - European Space Agencyemits.sso.esa.int/emits-doc/ESTEC/AO-1-5589-RD12-M4.pdf · MSR Technology Demonstration Overview : LIDAR Introduction 2 4 6 8 10 12

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All rights reserved © 2007, AstriumFP – ESTEC, 24 October 2007

MSR Precursor Mission MMSR Precursor Mission M--44

The The MoonTWINSMoonTWINS mission conceptmission concept

Final Presentation – ESTEC – October 24th 2007

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Agenda

Agenda:• Introduction• Science Objectives Synthesis• Mission Analysis Synthesis• GNC Analyses Updates & Synthesis• Propulsion System• On-Surface System Engineering Synthesis• Power & RF System Synthesis• System Synthesis• Conclusion, discussion & AOB

P. Regnier, 09:45-10:00

K. Geelen, 11:15-11:30

K. Geelen, 11:30-11:45

A. Povoleri, 10:30-10:45

E. Kervendal, 10:45-11:15

D. Ruf, 11:45-12:00

All, 12:45-13:00

P. Regnier, 12:00-12:45

P. Lognonne, 10:00-10:30

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Introduction P. Regnier

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Work Progress since PM3 (August 3rd) :• Preparation of NEXT mission concepts presentation to ESTAG • Study tasks suspended from August to beginning October due to

involvment of key personal in a phase C/D proposal for ESA• Mission Analysis updates : consolidation of baseline ∆V budget• On-surface system engineering updates : thermal control budgets update• Propulsion system analyses : architecture and sizing• Power & RF analyses updates : consolidation of battery and solar array sizing• GNC analyses updates :

landing analyses (reference descent and landing trajectory, ∆V budget) RV analyses (performances at contact)camera accommodation

• Science consultancy : IPGP documentation updated

Introduction

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Work Progress since PM3 (cont’d) :• System Synthesis :

spacecraft configuration updates (solar arrays, radiators)mass and propellant budgets updates, payload mass and power allocations updates

• Documentation : contributions to the Mission and System design Technical Note

Mission analysis (update)On-surface system engineering (update) Power and RF analyses (update)GNC (new)

Work remaining• Documentation delivery :

TN1 (Mission Requirements) and TN2 (Mission Conceptual Design) : combinedTN3 (Mission and Spacecraft Design) TN4 (Programmatics) already delivered

Introduction

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Introduction

Study Organisation

• Astrium-SAS : prime, GNC & technical synthesis, programmatics

• Astrium-Ltd : mission analysis, on-surface system engineering, propulsion

• Astrium-Gmbh : RF & power system engineering

• SENER : mechanisms analyses (landing legs)

• Deimos : hazard avoidance consultancy

• IPGP & DLR : science consultancy

• 230k€, 6-month study

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Study Logic

Introduction

Kick-Off

PM1

PM2

PM3

Mission Objectives & Requirements Definition

Mission Conceptual Design & Major Trade-offs :-landing site latitude

- P/L operations at night

Mission & SpacecraftPreliminary Design Programmatics

FP

MSR Precursor Mission High Level Requirements

MSR TechnologyDemonstration Objectives

ESA Directives(deselect science P/L, more

mission trade-offs)

Mission baseline : -one polar lander

- one non-polar lander- no capture mechanism

• Programmatics fully presented at PM3 (no change)

• Main technical evolutions since PM3 :– ∆V budget (margins for descent & landing)– propulsion mass budget

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The MoonTWINS Mission Concept : ObjectivesTechnology Demonstration Objectives for MSR

Autonomous RendezVous : focused on GNC and vision / LIDAR navigation• Target detection and tracking by an optical camera• Target approach and proximity operations with an optical camera and a LIDAR, in

MSR representative orbit dynamic conditions• Validation of GNC performances at contact through a touch-and-go manoeuvre• RF tracking and capture mechanism not supported (mass constraints)

Soft Landing : full validation of the two planetary landing technologies currently underdevelopment by ESA :

• optical navigation & LIDAR, including hazard avoidance and precision landing• in MSR representative conditions (as far as possible)

Potential for Science & Exploration :launching two landers on one Soyuz-Fregat means a reduced Science payload capacity but at two different landing sites (science network)most appropriate science payload = geophysics, especially seismometersmost favoured site for future exploration : Peak Of Eternal Light at the South Pole

Mission redundancy through twin landers (like US Viking / MER missions)

Introduction

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The MoonTWINS Mission Concept : Assumptions & Constraints Launch by Soyuz-Fregat 2.1b from Kourou

stringent launch mass constraintshared Ariane 5 ECA GTO commercial launch technically feasible but less attractive (flexibility, costs)

Launch opportunities from 2016 onwardsMinimise mission costs :

recurrence among the two landers, no new development (except MSR related technology)

System Approach :identify the system impacts of different landing sites (esp. wrt latitude)

• but landing on the anti-Earth side was not considered (comms relay issue) • at least one lander at a polar PEL is favoured (exploration perspective)

assess the system impacts of payload on-surface mission characteristics (power/survivalat night, RHUs allowed but no RTGs)assess the system impacts of enhancing MSR RV capture representativity (SampleCanister + capture mechanism + RF proximity link)consolidate the payload resulting allocations in terms of mass and power at night

Introduction

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The MoonTWINS Mission Concept : Mission Architecture Trade-off Launcher (S-F 2.1b vs AR5 shared) and launch injection strategy (direct LTO or

GTO) Staging approachTrade-off criteria :

useful massperformance, mission costs, complexity andrisks

Introduction

S-F launch in LTO S-F launch in GTO S-F launch in GTO Shared Ariane 5

commercial GTO launch

Launch performance ~2100kg (incl adapter)

~3200kg (incl adapter)

~3200kg (incl adapter)

typ. ~4000kg (without adapter)

Staging approach No propulsion stage

No propulsion stage

LISA-Pathfinder like propulsion stage No propulsion stage

∆V to Lunar Circular Orbit ~900m/s ~1600m/s ~1600m/s ~1600m/s TBC

Mass in Lunar orbit 2 x ~750kg 2 x ~900kg 2x ~800kg +200kg (LISA-PF) 2x ~1200kg

∆V to Lunar surface ~1900m/s

Lander dry mass allocation ~350kg each ~450kg each ~400kg each ~600kg each

Lander propellant capacity requirement ~650kg each ~1050kg each ~400kg each ~1400kg each

Useful Mass Performance 4th 3rd 2nd 1st

Mission Costs 1st (cheapest) 2nd 3rd 4th (TBC)

Mission complexity and risks

1st (least complex & risky) 2nd 4th (more complex

composite spacecraft) 3rd (more complex trajectory design)

Baseline

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The MoonTWINS Mission Concept : Mission Scenario

Introduction

~6-month in-flight phase,

~a few year-long surface mission

Cluster-likeLEOP operations

150km altitude circular orbit

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MSR Technology Demonstration Overview : RendezVous

Introduction

Objectiveso validate vision-based & LIDAR RV technologies, GNC

algorithms and operations required for MSRo in representative orbit kinematic conditions o but with much more operational flexibility and safety

(no round trip delay, omni-directional TM, high data rate)o but no capture mechanisms nor RF system

Baselined RV technology (same as for landing)o Vision-based navigation : ESA AutNav & HARVD studies heritageo LIDAR (on one lander) : ESA LiGNC study heritage

Target detection and acquisition (50 – 500km)o validation of on-board image processing algorithms for target optical

detection, acquisition and trackingo results can be extrapolated to MSR conditions (NAC at large range)o on-ground restitution of target orbit

Intermediate rendez-vous phase (down to a few km)o autonomous target optical trackingo trajectory guidance from the groundo target orbit rallying manoeuvres

Orbit periods around Mars and the Moon

1.5

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MSR Technology Demonstration Overview : RendezVous

Introduction

Terminal rendez-vouso validation of trajectory guidance

options (V-bar, R-bar hops, football orbits)

o validation of closed loop GNC and collision avoidance manoeuvres

o use of the LIDAR in addition to the optical camera o use of the landing legs footpads at contact (touch-and-go manoeuvre) :

– minimum mass solution (<1kg per leg for design adaptations)– well suited for shock damping & safe contact (enlarged footpads area)– common WAC / LIDAR boresight directions for landing and RV– safety ensured by several step-by-step iterations before final contact

o reconstitution of GNC performance at contact (relative attitude / lateral offset) through inertial sensors and optical camera measurements (useful to specify the MSR capture mechanism)

landing legs footpadsused at contact

Target

R-bar to NadirOptical sensor FOV limit

Approach from loweraltitude phasing orbit

SK1SK2SK3SK4

V-bar towardsorbital velocity

Final Translation

Target

R-bar to NadirOptical sensor FOV limit

Approach from loweraltitude phasing orbitApproach from loweraltitude phasing orbit

SK1SK2SK3SK4

V-bar towardsorbital velocity

Final Translation

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MSR Technology Demonstration Overview : Optical Nav

Introduction

o based on NPAL study heritage : a technological breakthrough for Vision-based Navigation (ESA science Critical Technologies Program, 2001-2006)

o breadboard camera and image processing / navigation algorithms now qualified in real-time environment (TRL 4-5)

o soon to be tested on the ESA Precision Landing GNC Test Facility (TRL 5-6)o based on the extraction and tracking of unknown feature points o assisted by radar altimeter for robustness / faster convergenceo light weight / low cost

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MSR Technology Demonstration Overview : LIDAR

Introduction

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Previous Elevation Map

Current Elevation Map

∆X, ∆Y

∆X, ∆Y, ∆Z

Generation of elevation map

Raw Elevation Map(no regular grid)

Re-sampling of the map

Conventional Image Processing

VerticalCorrelation

Navigation filter (Kalman)

∆X, ∆Y, ∆ZVx, Vy, Vz

o based on LiGNC study heritageo LIDAR breadboard development

on-going in Europeo more robust to illumination

conditions than vision navigation

o used at short ranges onlyo heavier, power hungry

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MSR Technology Demonstration Overview : Hazard Avoidance

Introduction

o based on vision or LIDAR (LIDAR preferred at grazing Sun incidence angles)

o hazard mapping and re-targeting in the last km o very strong background and heritage at

Astrium and Deimos (VBRNAV)

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SLS0 SLS1 SLS3

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o need to be linked to landing site constraint (PEL) and IMU accuracyo based on image correlation techniqueso needs an on-board DEM or 2D

terrain model of the landing areao based on the on-going

Optical Flow Navigation System for Landing ESA study

o landing accuracy < 100m

MSR Technology Demonstration Overview : Precision Landing

These technologies were not further investigated in the frame of the MoonTWINS pre-phase A study, but their applicability to the MoonTWINS mission scenario and spacecraft design was assessed (system design constraints, resource sizing)

Introduction

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Science Analysis : example of payload and science return P. Lognonne (or Mark Wieczorek)

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Science Payload Boundary conditionPayload Mass < 17 kg depending on night operation

20% margins on payload included in payload massStatic payload only ( geophysics/radioastronomy/environement)Payload mass includes deployment systems (robotic arm, ejectionmechanisms)Assume 15 kg with margins.

PowerContinuous operation for SEIS, pulsed for MAG and Geodesy, Snapshotfor other0.80 Watt ~ Night Power 50 Watt ~ Day Power < 100Watt

Landing siteSouth pole and mid latitude

One polar and one non-polar lander

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Science page 2

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Payload information collection

Payload information has been collected in the science community mainly by email exchange

General Science objectives of the payload areMoon Internal structure ( e.g. 8th ILEWG resolution, point 11, ESSC-ESF report)Radio-astronomy pathfinder experiment (ESSC-ESF report)

Geochemistry/Mineralogy science objectives are excluded, being related to either rover or sample return

Payload mass are without margin and 20% margins are added.

Science page 3

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Payload list 1/2Instrument Mass

Kg Mean Power (W)

Science objectives Comments

Moon geophysics (8.6-15.7Kg) 3 axis Very Broad Band Seismometer (VBB)

4.2 0.7 Deep structure of t he moon, analysis ofthe shallow moonquakes, crustal thicknesslateral variations, detection of SQMs

20x more sensitive at t he frequency of Apollo LP (0.5 Hz) and larger dynamic/bandwidth. Based on GEP instruments. Acquisition common to SP. Include I/F and cover

3 axis Short Period Seismometer in single (SP) or local Network (NSP)

0.4 or

3.7

0.2 or

0.5

Crustal and regolith structure in the vicinity of t he landing sites, detection and characterisation of micro-meteorites or Subsurface and regolith structure in the vicinity of t he sites, detection and characterisation of micro-meteorites

10x better at the peaked frequency of Apollo SP (8Hz, 0.5 10-8 ms-2/Hz1/2) and larger dynamic/bandwidth. Based on GEP instruments

or 3 micro-penetrators with SP micro-seismometers and telemetry. New development

Geodesy experiments (GEO)

1.5-5 0-5 Measure parameters of the dynamics of the Earth/Moon system, including Moon librations and tidal deformation with implications for Lunar deep structure.

10x-100x better than results from the Laser Passive detectors, depending on the technology. Possible technologies are Ka-band transponders, passive Laser reflector or Active Laser.

Magnetometer (MAG)

0.75 0.15 Interaction of th e Earth magnetotail and solar wind with the Moon, magnetic sounding of the Moon

20x better resolution than Apollo (0.01 nT). Mass for dual magnetometers depending on t he technology. Magnetometer put on the surface. Either single magnetometer plus dedicated deployment or dual magnetometers using the robotic arm.

Mole/Heatflux/densitoemeter (MOLE)

1.75 0.1 Measurement of the heat flux, determination of the bulk content in radioactive elements, heat conductivity and density of the regolith

5 meter depth penetration instead of 2.3 m (Apollo 17). Based on GEP instruments

Science page 4

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Instrument Mass Kg

Mean Power (W)

Science objectives Comments

Radio-astronomy (2.50Kg) Radio-astronomy Receiver/GPR (RAS)

2.5 1 Regolith structure beneath the landing sites, detection of radio flashes from ultra-high energy cosmic rays and neutrinos hitting the Moon

Passive/active mode in the 0.1-30 MHz bandwidth. Based on Ex oMars WISDOM and GEP a nd Earth LOFAR technology

Sun/Mon Environment (2.55 kg) Solar wind monitoring instrument (WIND)

1.5 1 Solar wind monitoring, composition, energy of solar wind

TBA

Radiation sensor (RAD)

0.55-0.75

0.75 Measurement of the radiation level on the Moon surface

Several Technology available, including those developed by GEP and for human mission

Context/deployment (4.25 Kg) Camera (CAM) 0.75 N/A Verify landing site location and

instrument deployments, study visual characteristics of r ocks and soil at the site.

Micro-camera system based on previous ESA landers technology, in addition to those of t he landing and RDV systems.

Deployment arm (ARM)

3

N/A Deployment of the geophysical instruments on the Moon surface

From ExoMars GEP accomodation studies

TOTAL 18-25 ~10

Payload list 2/2

Science page 5

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Example of Payload (15 kg including margins)• VBB SEISMOMETER

- 4.2 Kg, 0.7 Watt Night• Heat flux

1.75 kg, continuous operations , internal battery for during night ( buriedinstrument)• Robotic Arm+ sensors hardness

4 Kg, used for heat flow and VBB deployment, MAG on the arm• Radio-Astronomy pathfinder experiment

2.5 kg, Day operation only. Snapshot operation during night ( 34 Whours perLunar Night)• Radiation sensor

0.55 kg, day operation only, snapshot operations during night• Magnetometer

0.75 kg, 0.1 Watt Night• Geodesy

1.2 kg, pulsed mode during night

Science page 6

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3 axis VBBInstrument must be deployed on the ground and covered by a sun shield, robotic arm Network operation is requestedScience return is proportional to the cumulative time of operationNear side stations can be used in synergy with Earth observation (detection of light flashes associated to meteorids impacts)Deep Moonquakes data from Apollo can be processed with future Lander data

Æ670 mm

Æ270 mm

245

mm

ĒŹUmbrellaŹČ likeThermal shield

deployment

Radio interferometry (GIN)Less pointing sensitive with omnidirectionnal antenna1.2 kg mass/5 WattRequest two stations at least (interferometry)Sub mm resolution can be achieved

SEIS High level requirements

Geodesy High level requirements

Science page 7

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MagnetometerSeveral groups in Europe with different technology but comparable massesSensitive to DC/AC magnetic field : imply specific shielding of the solar panels and Magnetic cleanliness programNeed to be deployed/ejected away from landerNetwork operation requested for interior soundingIncreased Science return if data from an orbiting magnetometer are available

Magnetometer High level requirements

Heat fluxVery low power needed after deploymentMole can go deep in the regolithDeployment to be performed during the dayBattery inside the buried mole for night operation ( charged during day)

Mole High level requirements

45 mm28 mm

30 mmBaseplate

(CFC)

Coil Supports(CFC)

ThermistorTerminal PCB

Pigtail

Glider assembly

Flat cable storage

Supports

Flat CableTractor Mole

Payload Compartment (HP3)

Image courtesy of Galileo Avionica, ©2005

Science page 8

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Radio-astronomy/GPRDeploy 2-3 antenna by Mortar system, acquisition of 30 Mhz with data processing, active mode for GPR, passive mode for radio-astronomyPathfinder experiment,aiming to characterize the radio environment, including during the nightFarside observations desired, but will request major reduction in data downlink even if an orbiter is available (due to orbiter visibility limitation)

Radiation sensorMonitor the radiation environment of the Moon

Lander facilitiesDeployment arm for deployment of MOLE and SEISCamera on boom, lunar context and Deployment activities

Other payload High level requirements

Science page 9

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Night operationCrucial for Network science (Seismology, Geodesy, Magnetometer)Crucial for radio-astronomy but only part of the nightCrucial for long-term monitoring (e.g. heat flux, tides) but with low data acquisition profile for geodesy and battery built-in the moleseveral profiles defined depending on the resources (average power, with peak power of 20 Watt)One low power night profile, one night for radio-astronomy and one night for VBB works by alternanceNight data are stored in low power mass memory

Night operation

Night 0 Night 1 DayVBB 0 0,7 0,7GEO 0 0,05 5MAG 0 0,05 0,15MOLE 0 0 2RAS 0,7 8RAD 0,1 1WIND 1

0,8 0,8 17,85

In Watt

Science page 10

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GeophysicsDeep Moonquakes seismology

Apollo ~3 stations leading to 6 travel time data on deep moonquakes for 4 focal parameters = > 2 direct informations on the structureApollo + MoonTwiin ~5 stations leading to 10 travel time data on deep moonquakes for 5 focal parameters = > 5 direct informations on the structure Apollo + single MoonTwiin ~4 stations leading to 8 travel time data on deep moonquakes for 5

focal parameters = > 3 direct informations on the structureHeat flow

2 heat flow measurements for Apollo= > 4(3) after MoonTwin/single MoonTwinGeodesy

5 Lunar reflectors left by Apollo and Luna, only 4 used = > 6(5) after moonTwin/single moonTwinHabitability

Radiation monitoringFirst Experiment, never flown on the Moon

Impact monitoringKg mass detection with the VBB,

Science from the MoonRadio-astronomy

First Experiment, never flown on the Moon

Science return/Apollo

Science page 11

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Science focus

Full geophysical exploration of the Moon interior

Will provide information on both thetemperature and mineralogyWill provide the size of the core andprobably content in light elementsUnique information on the core due to South Pole location

Exploration of the Environement on the SouthPole

radiation and micro-meteoritesmonitoring

Pathfinder experiment for future radio-observatory

Periodic observation of the « Earth » far side by using natural Moon libration

Might be an original mission after severalmissions deploying roving elements(ChangE’2, RLEP-2/3, Chandrayan-2, Selene-2)

Science page 12

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Science return

Conclusion• The two additionnal landers of MoonTwin were able to perform a very significant step with respect to the Apollo geophysicalnetwork

VBB GEO MAG MOLE RAS RAD

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Apollo versus MoonTwin Science return

ApolloApollo+MT1Apollo+MT1_2

• Single lander mission willimprove the geophysicalknowledge of the Moon, but an international collaboration withadditionnal landers will berequested for retrieving themoonTwin Science return

Science page 13

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Mission Analysis SynthesisA. Povoleri

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Summary

Transfer to the MoonDirect transferWeak Stability Boundary transfersEarth departure strategyMoon capture strategyEclipses

Operational orbitChoiceStability

Landing siteGround station visibilitySun elevation

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Transfer to the Moon

Two possibilities for the transfer:1. Direct (Hohmann) transfer ~5 days2. Weak Stability Boundary transfer ~100days

1. Direct transfer: essentially two impulsesImpulse at perigee to raise apocentre to Lunar crossing radiusImpulse at apogee for rendezvous with the Moon and captureTransfer duration is 5 days

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Transfer to the Moon

2. Weak Stability Boundary transfer~100days

DeltaV for apogee raising to Earth-Sun Lagrange point (L1 or L2): this manoeuvre costs about 70m/s more than apogee raising to Lunar crossingUse gravitational perturbation to raise orbit perigee: this way relative velocity at moon approach is lower than in direct transferLower relative velocity allows using Lagrange point L1 or L2 of Earth-Moon to assist capture in orbit around MoonPericentre burn is needed in orbit around Moon to lower the moon-relative apocentre and keep the s/c within the SOI of Moon

Note: possible to save DeltaV by performing Lunar Gravity Assist on the way to Lagrange point. Net DeltaV saving is ~50m/s

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Transfers to the Moon

• 2016-2018 transfers• Launch inclination=5.50, argument of perigee=1560

1. Direct transfers: RV with the Moon can only happen at the node

2. WSB transfers: low inclination favours equinoxes launch

WSB transfers allow 100m/s saving on DeltaV->Mission Baseline

1488.35-Jul-1858304797.81888-17909.830-Jun-1858299689.46878.79388941.66

1504.915-Jul-1858314823.51888-15200.711-Jul-1858310680.36872.39353919.12

1492.814-Apr-1858222805.01888-17048.79-Apr-1858217686.76878.79378068.47

1502.431-Mar-1858208816.71888-15826.727-Mar-1858204685.76878.01374703.82

Total DeltaVDateMJDDeltaVPeriApoDateMJDDeltaVPerigeeApogee

1420.1618-May-1858256665.8918881523527.723-Jan-1858141753.306878.2451221376.3

1391.919-Nov-1758076638.15188872367.284-Aug-1757969752.816878.0031206550.4

1388.4124-Jul-1757958634.18188863523.972-Apr-1757845753.266878.0331221376.6

Total DeltaVDateMJDDeltaVPeriApoDateMJDDeltaVPerigeeApogee

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Mission description

Departure strategy:- Launch into GTO (or similar)- Separation of spacecraft- Apogee raise sequence (typically 3 burns) by means of perigee burns

- Most mass-efficient strategy- Several burns in order to limit DeltaV losses (2.5%)- Period of the intermediate orbits chosen in a way such that easy strategy for

recovering missed burns is possible- Launcher dispersion correction manoeuvre incorporated in the sequence- Correction after the last perigee burn

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Mission description

Insertion into operational orbit at the Moon:- Capture into high apocentre Moon orbit

- Achieved for free in the WSB transfer scenario- For direct transfer it is necessary to perform a capture manoeuvre by

means of propulsion system- Apocentre lowered to target 150km altitude value by means of

propelled manoeuvres - Reverse strategy of the Earth departure- DeltaV loss is limited to <1%

Rendezvous experiment performed 2 months after arrival at Moon, in eclipse-free orbit

Landing:- Rehearsal in orbit with pericentre~10-20km- After rehearsal pericentre raised to safety altitude (50km)- Actual landing starts from 50 by 150km orbit

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Eclipses during transfer

•For direct transfers, there are 2 locally optimal opportunities for each lunar month•Occurrence of apogee eclipses has been evaluated for all these locally optimal opportunities in the 2016-2018 timeframe•Only 2 transfers experiencing eclipse have been found: 12-9-2016 (27 hours) and 27-3-2018 (17 hours)•These opportunities should be avoided

•For WSB the problem doesn’t exist if L1 is targeted•If L2 targeted eclipses may happen, but can be avoided with small modification to the transfer

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Relative velocity gives Arrival plane 1

MoonRelative velocity gives Arrival plane 2

Operational orbit choice/ Impact on transfers

Rendez-vous experiment happens in eclipse-free orbit, a good choice is a terminator orbit.

Rendezvous experiment is foreseen 2 months after arrival. Local Solar Time of node of capture orbit has to be roughly 10a.m.-10p.m.

In nominal Hohmann transfer, approach velocity is tangential to Moon’s velocity

Plane is determined by arrival dateFixed launch condition impose arrival at the node of Moon’s orbitGeometry repeats almost unchanged in Earth-centered frameLST of the node of the capture orbit changes by 24 hours over a yearAny LST can be achieved twice a year onlyOptions to modify orbital plane exist, they all imply DeltaV penalty

Figure: orbital planes for arrivals separated by 5 days with nominal Hohmann transfers

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Operational orbit choice/ Launch windows

Fixed launch conditions potentially introduce some launch windowissues, because there are only a discrete numbers of direct transfers in the year producing acceptable arrival conditions at the Moon, in terms of orbital plane

Penalty for delay/advance in the launch can be relevant in case of launcher injecting directly into Moon crossing orbitStrategy with launch into intermediate orbit and apogee raise sequence is much more flexible (the window applies to the last burn in the apogee raise sequence)It is also possible to think of a strategy where, after the last burn, the s/c flies around one and a half revolutions before rendezvous with the Moon

For WSB transfers there is much more flexibility in plane choiceStrategy with apogee raising sequence is, again, beneficial in terms of launch windows

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Operational orbit stability

•Assumptions:150km circular operational orbit. Node longitude arbitrarily set to -16o

Several inclinations investigated (80o-105o range)Harmonics terms up to order 70 have been consideredOrbit propagated for 30 days

•Affected parameters are pericentre altitude and longitude of ascending node

•Compensation of change in pericentre/apocentre asks for 5.48m/s every month in the worst case (90o inclination)•Leaving the s/c uncontrolled for 3 months results in 108*172km orbit (90deg incl.)

pericentre evolution

136138140142144146148150152

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Stability of landing rehearsal orbit

• Assumptions:Elliptical orbit: pericentre 10-20km, apocentre 150kmStability to be guaranteed for 2/3 orbits (i.e. 4-6hours)Plot: pericentre evolution for 10*150km orbit, 90o inclination (worst case) over 72 hours

In the first 6 hours pericentre drops by 1 kmSame behaviour for 15 or 20 km pericentre orbits (i.e. pericentre would drop by ~1km as well

Pericentre altitude

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Landing site: Ground station visibility

• Figures of interest are:Elevation of landing site from ground station (angle above the local horizon to which landing site appears as seen from ground station, i.e. optimal is 90 degrees)Elevation of ground station from landing site

• When both elevations are positive, there is possibility of contact between LS and GS

Quality of contact is determined by the magnitude of elevations (and range)

LS Elevation

GS

LS

GS ElevationGS

LS

GS horizon LS horizon

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Ground station visibility

• Analysis performed on several LS and GS• 4 representative landing sites have been considered:

South Pole83deg N, 0deg longitudeEquator, 0deg longitude45degN 45degE

• 4 representative GS consideredKourou (5.25N, 52.8W)Kiruna (67.85N, 20.96E)Maspalomas (27.76N, 15.63W)Perth (31.80S, 115.88E)

• LS elevation and GS have been plotted over 1 month durationVariation during the year is very limited

• For each LS, sun elevation has also been plottedFor this quantity there can be significant yearly variation (ex: at the Poles)

• Topography not taken into account!

45



•Example 1: South Pole. All figures refer to Sep 2011, except from Kiruna (Dec 2011)

Perth

0

20

40

60

80

0 5 10 15 20 25 30

Time (days)

Elev

atio

n (d

eg)

LS elevation from GSGS elevation from LSSun Elevation from LS

Kourou

0

20

40

60

80

0 5 10 15 20 25 30

Time (days)

Elev

atio

n (d

eg)


Maspalomas

0

20

40

60

80

0 5 10 15 20 25 30

Time (days)

Elev

atio

n (d

eg)


Kiruna

0

20

40

60

80

0 5 10 15 20 25 30

Time (days)

Elev

atio

n (d

eg)


46



• Example 2: 83deg N, 0deg longitude. All figures refer to Sep 2011

Perth

0

20

40

60

80

0 5 10 15 20 25 30

Time (days)

Ele

vatio

n (d

eg)


Kourou

0

20

40

60

80

0 5 10 15 20 25 30

Time (days)

Ele

vatio

n (d

eg)


Maspalomas

0

20

40

60

80

0 5 10 15 20 25 30

Time (days)

Ele

vatio

n (d

eg)


Kiruna

0

20

40

60

80

0 5 10 15 20 25 30

Time (days)

Ele

vatio

n (d

eg)


47



• Example 3: Equator, 0deg longitude: All figures refer to Sep 2011

Perth

0

20

40

60

80

0 5 10 15 20 25 30

Time (days)

Elev

atio

n (d

eg)


Kourou

0

20

40

60

80

0 5 10 15 20 25 30

Time (days)

Elev

atio

n (d

eg)


Maspalomas

0

20

40

60

80

0 5 10 15 20 25 30

Time (days)

Ele

vatio

n (d

eg)


Kiruna

0

20

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60

80

0 5 10 15 20 25 30

Time (days)

Elev

atio

n (d

eg)


48



• Example 4: 45degN, 45degE: All figures refer to Sep 2011

Perth

0

20

40

60

80

0 5 10 15 20 25 30

Time (days)

Elev

atio

n (d

eg)


Kourou

0

20

40

60

80

0 5 10 15 20 25 30

Time (days)

Ele

vatio

n (d

eg)


Maspalomas

0

20

40

60

80

0 5 10 15 20 25 30

Time (days)

Ele

vatio

n (d

eg)


Kiruna

0

20

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60

80

0 5 10 15 20 25 30

Time (days)

Ele

vatio

n (d

eg)


49


Conclusions

WSB transfers selected as mission baseline because of lower DeltaV Best opportunities around the equinoxes

Departure strategy is launch into GTO and apogee raise sequence (3 burns)

Eclipses during transfer can be easily avoidedRendezvous experiment places constraints on plane of operational

orbitWSB transfers are more flexible than Hohmann transfers

Stability of operational orbit is not an issueStability further improved with DeltaV allocation

Ground station visibility: different landing sites and ground stations have been considered. Low latitude ground stations provide the best coverage in any case

50


DeltaV budget

Delta-V Budget

Departure DeltaV 755 m/sLoss (2.5%) 18.9 m/sLauncher dispersion correction 30 m/sMoon capture 665 m/sLoss (1%) 6.7 m/sNavigation Delta-V 20 m/sOrbit maintenance (2 months) 11 m/sDescent Rehearsal delta-V 37 m/sRendezvous delta-V 10 m/s

Total 1543.53 m/sTotal Including Margin (3%) 1589.83 m/s

Worst case transfer

51


GNC Analyses Updates & Synthesis E. Kervendal

52


1. Landing UpdatesPreliminary baseline

2. Rendezvous UpdatesFinal Approach

3. SynthesisCamera ImplementationLandingRendezvous

53


1. Landing Updates: Preliminary baseline for landing scenario

54


Landing Preliminary baseline updated:

Improved vision-based navigation performances during VGP

Small incidence variationSmall Viewing Angle variation

Cope with SC capabilitiesThrottability Maximal angular acceleration

Cope with MSR scenarioFinal vertical landing (as far as possible)Compatible with LIDAR technological constraints (FOV, incidence)

Angles definition

Viewing Angle

24.0 degMaximal Viewing Angle Variation

1.32 NmMaximal Torque

1526.27 NMinimal Thrust

2500 NMaximal Thrust

5.40 degMaximal Incidence Variation

60.59 m/sVelocity

-68 degFPA

32.57 sTime till touch down

1000 m conditions

Preliminary baseline characteristics

540.8 sDuration

414.42 kgPropellant mass

1978.88 m/s∆V budget

Total budget

Landing Baseline Budget

55


Landing baseline comprises:

Two main phases:Inertial Guidance phase(absolute navigation from IMU, predefined thrust profile)

Visual guidance phase (relative navigation w.r.t. LS, possible retargetings, dispersions compensation)

Four main eventsPDI: beginning of powered descent phaseHigh gate: acquisition of landing site, initialization of relative navigationVGPEP: beginning of hybrid navigation (camera / Lidar + IMU)MECO: engines cut-off, end of GNC

Landing Baseline description

IGP VGP

PDI

VGPEP

HighGate

MECO

IGP VGP

PDIPDI

VGPEP

HighGate

MECO

Preliminary baseline scenario: phases and events

56


Landing Baseline Profiles

Thrust (N) vs. time (s) transition = 132.76 N (6.63 % full throttability)

Altitude (m)

Range to LS (m)

FPA (deg)

Total Velocity (m/s)

57507.520000

0-8991.2-463774

-90-20.900

1267.421699.94

MECOVGPEPPDI

Preliminary baseline characteristics

IGP Trajectory baselineVGP trajectory baseline

Total velocity (m/s) vs. time (s) Pitch, FPA and VA (deg) vs. time (s) transition = 10 deg (within angular acceleration

capability)

57


2. Rendezvous Updates: Final approach

58


Final approach is a forced translation:From -100 m to targetAssumed entirely powered (no final free drift to target)Starts with differential velocity of 0.2 m/s and slow down to 0.1 m/s at contact

Assumed to have several, typical perturbations:

Instruments noise and bias (Lidar, camera, star tracker, IMU)Errors on Thruster directions and amplitudeDispersions at initial point

Based upon a simulator developed for LiGNC* study, with complete GNC loop

Final Approach baseline

* From LiGNC, ESA contract 17389/03/NL/AG

R-bar

V-bar

100 m

Hunter Target

Direction of final translation

R-bar

V-bar

100 m

Hunter Target

Direction of final translation

Final translation along V-bar

Simplified simulator for Rendezvous

59


Complete GNC loop: Simple guidance:

Initial velocity = 0.2 m/sAt t = 250s, V = 0.1 m/s

Simple control based on PD approach (RV performances can be improved)Independent Attitude and position controlThrusters configuration representative of MoonTWINSbaseline

Errors observed on local frame:X-axis along V-bar directionY-axis ~ R-barZ-axis ~ N-bar (perpendicular to relative orbit)

Small performances dispersions observed, compatible with landing legs dimensions (diameter = 0.3 m)

Final Approach results

Trajectory and velocity (m/s) vs. time (s) for ideal case (no dispersions)

Sample of trajectories for real cases

0.051.29∆V (m/s)

0.0010.0015Vz (m/s)

0.001-0.001Vy (m/s)

0.0010.0904Vx (m/s)

0.030.04Z (m)

0.01- 0.03Y (m)

00X (m)

1σMean

Performances at contact

60


3. Synthesis

61


Design of camera shall fulfill constraints from landing and rendezvous phases:

Velocity visibility during VGPEP (given by maximal variation of incidence)Landing site visibility (given by maximal variation of viewing angle)Accuracy for retargetings and final approach

Target detection rangeVisibility of target during closing maneuvers and final approach (with Moon exclusion angle)

Camera with FOV = 40 ° and offset = -10°matched both landing and rendezvous constraintsGood compromise

Large RV detection rangeArea coverage / accuracy / image overlapping for Landing

Camera Implementation

Final map (FOV°, offset°) for camera design and implementation

OKRV

OKLanding

75.9 % (minimal overlapping between two consecutives images during landing sequence)

Image overlapping

0.14 m/pixel100 m

1.41 m/pixel1 km

142.1 m/pixel100 km

710.87 m/pixel500 kmAccuracy

Range ~392 kmDetection

FOV = 40°, offset = -10 °

Characteristics of preliminary camera design

62


Landing

Preliminary baseline for landing and camera design allow for:

High performances navigationContinuous visibility of velocity vector during VGP (and 70.3 s prior to VGPEP)

Continuous visibility of landing site during VGP (and 155.8 s prior to VGPEP)

Demonstration of MSR-like scenarioQuasi-vertical landingDispersions compensationRetargetingsUse of Lidar at 1000 m-altitude point or before (for one SC)

63


Dispersions at VGPEP (Visual Guidance Phase Entry Point):

Result from typical errors induced by Inertial Guidance Phase (dispersions from IMU, thrust errors…)

Typical value: [350 m, 250 m, 1 m/s, 0.3 m/s] at VGPEP*

MBTL allows for compensating these errors between VGPEP and 1000 km point while:

Throttability and maximal angular acceleration are within feasible domainNavigation performances ensured

Small ∆V budget necessary for dispersions compensation

Dispersions at VGPEP

* From LULA, ESA contract 9558/91/NL/JG

2.73∆V (m/s)1.6Time till MECO (s)

0.48Propellant mass (kg)σ

Budget for typical dispersions compensation

Examples of dispersions compensation

64


Retargetings

Retargetings can be performed:At 1000-meter-altitude point (identification of slope)At 500-meter-altitude point (identification of hazardous boulders)If necessary, before 150-meter-altitude point (LS freezing point)

The new LS can be reached:Under visibility constraintsSC capabilities (min/max thrust, torque manoeuvrability)

+ 57 (max)-55 (min)+140 (max)-100 (min)Range (m)

79.7111.277.24111.4310.91114.71DV (m/s)

11.821.711.616.851.6817.35Propellant mass (kg)

22.596.3122.3332.56.1835.08Time (s)

1σMean1σMean

BaselineRetargetingBaselineRetargeting

500 m1000 m

Retargetings and preliminary costs

Example of retargeting at 1000 m and 500 mTotal retargeted range = 170 m for 133.3 m/s

Hazard avoidance can be performed with actual camera /

LIDAR design and reference trajectory.

65


Rendezvous

Preliminary baseline for rendezvous and camera design allow for:

High performances navigationDemonstration of MSR-like scenario

Typical MSR scenario (far range, close proximity operations) Far range target detection (500 km)Target orbit restitution (on-ground), through adequate orbit relative kinematics Demonstration of final GNC performance through contact at end of final approach (with Lidar on one SC)

66


Orbit restitution

Relative orbit can be estimated by chaser

During far range rendezvous (between 400 km to 10 km) and close range rendezvous (10 km to 100m),Based on station keeping maneuvers (impulse in radial direction, relative orbit stability, relative los variation)Improved navigation performances∆V cost is function of R-bar amplitude (∆Z) and orbital period

Evolution of relative position, estimation error (m) vs time (s)

Station keeping at 400 km from target in (R-bar, V-bar) frameR-bar amplitude = 2 km, T = 7357s, ∆V = 1.708 m/S

Rbar

Vbar

∆X

∆Z

Rbar

Vbar

∆X

∆Z

Principle of station keeping for orbit restitution

67


Closing maneuvers

Design of closing maneuversDuring far range rendezvous (between 500 km to 10 km) and close range rendezvous (10 km to 100m),Based on ∆V impulse in V-bar directionConstrained by Camera/Lidar FOV and ∆V budgetUse of Lidar once relative range about 5 km

Final maneuvers with current camera design can be done

* K. Yamanaka, F. Andersen, “New State Transition Matrix for Relative Motion on an Arbitrary Elliptical Orbit”, J. of Guidance, Control and Dynamics, Vol 25 (1), January-February 2002

Principles of closing maneuvers and station-keeping orbit, constrained by instrument Field Of View

Closing maneuvers between 7 km and 100 m, for FOV = 40 °

- Based on Yamanaka-Andersen equations*, ∆V = 1.93 m/s -

Closing maneuvers between 10 km and 100 m, for FOV = 20 ° (Lidar)

- Based on Yamanaka-Andersen equations*, ∆V = 2.76 m/s -

68


Conclusion

• Current baseline for landing and rendezvous designed to demonstrate critical aspects of MSR typical scenario

Rendezvous phaseOrbit restitution with football-like orbit inducing LOS variationTarget detection in far range rendezvous typical domain (500 km)Typical closing maneuvers (V-bar hops)Demonstration of GNC performance through demonstration of contact at end of fully powered, final approach

Landing phaseSimulation of quasi-vertical landing (Martian-like)Precision landing through High performances navigation and Dispersions compensationsSafe landing with hazard avoidance and retargetings

• Demonstration of Hybrid navigation (vision-based navigation) and Lidar based navigation for highly critical phases

69


Propulsion System Analyses K. Geelen

70


Propulsion System

• Requirements • Trade-Offs

Propulsion TypeThruster Selection and ConfigurationPropellant SizingTank Selection and Configuration

• Propulsion System Architecture• Propulsion system Mass Budget• Conclusion

71


Requirements

• Primary Functions:Provide thrust for attitude control and orbit maintenanceProvide thrust for entry into LTO and LOI (WSB baselined)Provide thrust for the controlled descent

Delta-V Budget Delta-V Incl Margin Departure DeltaV 752.80 775.38 m/s Loss (2.5%) 18.82 19.38 m/s Launcher dispersion correction 30.00 30.90 m/s Moon capture 634.20 653.23 m/s Loss (1%) 6.34 6.53 m/s Navigation Delta-V 20.00 20.60 m/s Orbit maintenance (2 months) 11.00 11.33 m/s RDV Rehearsal delta-V 37.00 38.11 m/s Rendezvous delta-V 10.00 10.30 m/s Landing delta-V 1900.00 1995.00 m/s Total 3410.16 3560.77 m/s

72


Propulsion Type

• Electrical propulsion not considered:Thrust level not sufficient for landingEP for transfer is costly and results in long transfer times

• Monopropellant versus BipropellantE.g. 450 kg dry mass and 3685m/s delta-VMonopropellant propellant mass: 2033 kg (I sp = 220s)Bipropellant propellant mass: 980 kg (I sp = 325s)Difference in propellant mass cannot be recovered by dry massBipropellant most mass efficient

• Achievement of ThrottlePulse Width Modulation since no throttleable European engines

• Regulated System

73


Bipropellant Type

• There are six main technologies available:Liquid Oxygen (LO2) and Refined Petroleum (RP-1)Liquid Oxygen (LO2) and Liquid Hydrogen (LH2)Nitrogen Tetroxide (N2O4) and Hydrazine (N2H4)Nitrogen Tetroxide (N2O4) and Monomethylhydrazine (MMH)Liquid Fluorine (LF2) and Hydrazine (N2H4)Chlorine Trifluoride (CIF3) and Hydrazine (N2H4)

• Nitrogen Tetroxide with MMH or Hydrazine preferred:Greater maturity within a European contextEasier storability and handling than cryogenic solutions

• Slight preference of N2O4/MMH Much greater experience within EuropeMixture Ratio allows identical tank sizes for both propellants

74


Thruster Type

• Two main European thruster suppliers• Astrium ST & MT Aerospace Satellite Products existing thruster range

ThrustersS-4 (4 N, development), S-10 (10 N, flown), S-22 (22 N, development)

Main EnginesATV (200-250 N, development), S-400 (400 N, flown), EAM (500 N, development)

75


Selection Criteria

• Thrust RangeA high and variable thrust to weight ratio is preferred to minimise losses during landing

• Specific ImpulseHigher specific impulse means less propellant needs to be carried

• Maximum ThrustHigh maximum thrust means less engines, more simplicity

• Minimal CostSmallest number of main engines to achieve maximum thrustReuse either existing engine/thrusters or those under active developmentSingle supplier for all engine/thrusters to reduces procurement costs

• European Sourcing

76


Baseline Selection

• Two engines baselined:A single main engine: 500N apogee motorEight 220N ATV thrusters currently under development

• European technology from single supplier• Balance between performance and number of engines• Characteristics:

500N engine: <5 kg, 325 Isp220N Thruster: 2.4 kg

Specific impulse/thrust dependent on inlet pressure (thrust level), nozzle design, material choices

77


220N ATV Thruster

• Nominal thrust of 220N occurs at 17bar & 20°C (Isp = 286.5s)

• Higher inlet pressure not preferred:Main engine with valves and other hardware qualified for 17bar.

• Isp improvement possible:By using platinum alloy for the chamber & nozzle (and 1:50 nozzle expansion ratio instead of 1:40). Increases mass by 0.4kg, but a supplier-estimated Isp would be about 300s-305 s in steady state mode. 300s used to dimension propellant because of mass criticality

78


Propellant sizing

Lower LANDER Polar LANDERLander dry mass including 20% margin 500.0 500.0 kgLanding delta-V including 5% margin 1942.5 1995.0 m/sI sp landing 305.3 305.3 sOACS propellant 2 2 kgLanding propellant 456.7 473.6 kgMass before landing 958.70 975.62 kgRendezvous manoeuvres + rehearsals delta V 49.44 49.44 m/sOrbit maintenance delta-V 11.55 11.55 m/sI sp lunar orbit 323 323 sOrbit maintenance propellant 18.6 19.0 kgLander mass after lunar orbit insertion 977.3 994.6 kgLOI delta-V 692.41 692.41 m/sAOCS propellant 4 4 kgIsp 323 323 sLOI propellant mass 242.8 247.0 kgLander mass before LOI 1220.1 1241.6 kgLEOP and Transfer delta-V 846.8 846.8 m/sISp 323 323 sPropellant mass 373.93 380.51 kgAOCS propellant 4.00 4.00 kgPropellant 1,102 1,130 kgResiduals 11.02 11.30 kgTotal Propellant 1,113 1,141 kg

79


Propellant Sizing

• 1141kg of propellant required assuming I sp for 500N engine to be 325 s and Isp of 300s for 220N thruster for polar lander

Combined specific impulse for descent: 305s• This results in 711 kg of NTO and 431kg of MMH for a

mixture ratio of 1.65. • Equivalent propellant volumes with NTO at 1444 kg/m3 and

MMH at 875 kg/m3 are 507 litres for both oxidiser and fuel assuming 3% ullage.

• Bipropellant tank options, fuel and oxidant can be stored in:A single 507 litre tank (2 tank configuration)A pair of 254 litre tanks (4 tank configuration)Three 169 litre tanksFour 127 litre tanks

80


Tank Selection Criteria

• Minimal MassMission mass critical and less dry mass requires less propellantMinimise tank number to reduce residuals (static and tank to tank mismatch)

• Minimal CostReuse of existing product familiesSingle supplier: Reduces procurement costs

• Maximise tank commonality (size, type and mounting)First preference common tank volumes for fuel and oxidantSecond preference common mounting, height or diameter to ease structural accommodation

• European sourcingCurrently there are two main European propellant tank suppliers

Astrium ST, existing bipropellant tank family: OST tanks MT Aerospace Satellite Products Ltd., existing tank families(Eurostar tanks)

81


Tank Selection

• 2 tank configuration (1 fuel tank and 1 oxidiser tank):OST06/0 or Eurostar 3000 tank are candidatesResults in a tank mass of between 50 and 52.2 kg

• 4 tank configuration:Eurostar 2000 tanks results in a tank mass of 54kg.

• 6 tank configuration: T11/4 tank, (over dimensioned)Results in a tank mass of 66 kg.

When considering the tank dimensions and accommodation issues, the 4 tank configuration using Eurostar 2000 tanks is the preferred option.

82


Pressurant

• For a regulated system, the pressurant mass is estimated to be about 3.3kg.

• Two most suitable options with respect to volume and tank mass is a tank from Pressure Systems Inc. (80465-1) and the EADS ST 90l tank.

83


System Architecture

84


Mass BudgetHardware subtotsl Margin 113.2 kgMajor components 94.60 99.33Pressurant tank 12.0 kg 5.00% 12.60Oxidant Tank 27.0 kg 5.00% 28.35Fuel Tank 27.0 kg 5.00% 28.35Main engine 500N thruster 5.0 kg 5.00% 5.25220N thrusters 19.2 kg 5.00% 20.16Reaction Control thrusters 4.4 kg 5.00% 4.62High pressure pressurant assembly 0.73 0.76Fill & Vent Valve (HP) 0.120 kg 5.00% 0.13Pressure Transducer (HP) 0.286 kg 5.00% 0.30Normally Closed Pyro Valve 0.320 kg 5.00% 0.34Low pressure pressurant assembly 2.43 2.55Pressure Regulator 1.150 kg 5.00% 1.21Non Return Valve 0.340 kg 5.00% 0.36Fill & Vent Valve (LP oxidant) 0.120 kg 5.00% 0.13Fill & Vent Valve (LP fuel) 0.180 kg 5.00% 0.19Normally Closed Pyro Valve 0.640 kg 5.00% 0.67Propellant distribution Assembly 3.42 4.27Fill & Drain valve (LP Oxidant) 0.120 kg 5.00% 0.13Fill & Vent valve (LP Oxidant) 0.060 kg 5.00% 0.06Fill & Drain valve (LP fuel) 0.120 kg 5.00% 0.13Fill & Vent valve (LP fuel) 0.060 kg 5.00% 0.06Pressure Transducer (LP) 1.144 kg 5.00% 1.20Single flow Latch Valve 1.280 kg 5.00% 1.34Liquid Filter 0.640 kg 5.00% 0.67Normally Closed Pyro Valve 0.640 kg 5.00% 0.67Main Engine Assembly 0.12 0.13Fill & Vent Velve (LP Oxidant) 0.060 kg 5.00% 0.06Fill & Vent valve (LP fuel) 0.060 kg 5.00% 0.06Pipework supports and fittings 5.87 6.16Pipework 2.200 kg 5.00% 2.31Pipe and component supports 3.666 kg 5.00% 3.85

85


Conclusions

• Regulated bipropellant system most promising solution for a lunar lander

• A 500N main engine with 8 ATV thrusters is a feasible solution• Pulse Width Modulation needed for landing• An engine for a lander project would need a delta development for

optimization of the engine to the particular operation domain / application

86


On-surface System Analyses SynthesisK. Geelen

87


Outline

• Introduction: Mission Options• Landing Sites• Day Time Operations• Night Time Operations• Thermal• Conclusions

88


Introduction

• Different mission scenarios depending on a combination of:The landing site

At Peak of Eternal Light (PEL) : near polesEquatorial or equivalentOther?

The mission durationNo survival at night: Hibernation at night (no science operation): Science operation at night?

89


Landing Site Trade-Off

•More variations and extremes then poles with minima of 84K and maxima of 395K

• Daylight temperatures around 230K that drop to 70K in shadow• Varied local topography lead to short periods of shadowing (thermal cycling)

Thermal

• ~2 weeks of daylight • Sun elevation sinusoid with amplitude of (90 º -latitude) varying with Lunar Summer/Winter• Gentle local topography predictability in illumination conditions

• During summer: ~95% of illumination but uncertainties/risk• 8 terrestrial days of darkness assumed per month (TBD)• Sun is very low in the Lunar sky (~degrees)

Sun Illumination

• Earth continuously visible between:< 83.13° N or S and< 81.84° E or W (near side)

• Earth occultation for two weeks every month

Earth visibilityLower latitudePoles (PEL)Characteristic

Trade-Off performed on systems level (incl. science, mass, complexity, etc ):One Lander targets PEL, other lander flexible for lower latitude

90


Day Time Operations

Different modesPayload Deployment Mode:

Short duration, high power: battery usedNominal Science Operations

Instruments operating: parametric analysisOnly receiver on

Communication Mode, No ScienceAt dawn before switching instruments on

Communication Mode, Including ScienceCommunication Mode, High Data Rate

High Data rate: amplifier on when excess power from solar arrays available after batteries are chargedOnly viable for polar lander

91


Day Time Power Budgets

Equipment Basic Power [W]

Incl. Margin (%)

Basic Power [W]

Incl. Margin (%)

Basic Power [W]

Incl. Margin (%)

Basic Power [W]

Incl. Margin (%)

Basic Power [W]

Incl. Margin (%)

Data Management 36 30 30 30 30CDMU 15.00 18 15.00 18 15.00 18 15.00 18 15.00 18EIU in deployment mode/hopping 15.00 18 off 0 off 0 off 0 off 0EIU surface mode off 0 10.00 12 10.00 12 10.00 12 10.00 12Power Subsystem 36 36 36 36 36PCDU 30.00 36.00 30.00 36.00 30.00 36.00 30.00 36.00 30.00 36.00Communications 54.6 54.6 54.6 117.6 21Receiver 20.0 21 20.0 21 20.0 21 20.00 21 20.00 21Transmitter 32.0 33.6 32.0 33.6 32.0 33.6 32.00 33.6 off 0Amplifier off 0 off 0 off 0 60.00 63 off 0Payload 36 0 24 24 24Payload Day power 0.00 0 0.00 0 20.00 24 20.00 24 20.00 24Payload deployment power 30.00 36 off 0 off 0 off 0 off 0Thermal Subsystem 0 0 0 0 0Miscellaneous 0.00 0 0.00 0 0.00 0 0.00 0 0.00 0Total User Load 163 121 145 208 111Power Harness Losses 3.25 2.41 2.89 4.15 2.22PCDU Conversion Losses 9.76 7.24 8.68 12.46 6.66Total Power 176 130 156 224 120System Margin 20.00% 20.00% 20.00% 20.00% 20.00%

Total Power with Margin 211 W 156 W 187 W 269 W 144 W

Payload deploymentCommunication

(no science)Communication (plus science) High rate comms Science Mode

Not sizing solar array!

92


Night Time Operations

• Trade-Off between:1. No Survival at night: Not further considered2. No science: Survival Mode only

• Minimum of equipment used to ensure survival3. Limited science during the night

a) Centralised architecture: Lander CDMU and PCDU switched on, EIU switched off

b) Decentralised architectureInstrument has internal mass memory and processing powerLow power PCDU mode: under-voltage- & over-current-protection and timer

93


Equipment Basic Power

Margin (%)

Basic Power [W]

Incl. Margin (%)

Basic Power [W]

Incl. Margin (%)

Basic Power [W]

Incl. Margin (%)

Data Management 0 18 0CDMU 15.00 20.0% off 0 15.00 18 off 0EIU 10.00 20.0% off 0 off 0 off 0Power Subsystem 1.2 36 0PCDU decentralised option 1.00 20.0% 1.00 1.2 off 0 off 0PCDU centralised option 30.00 20.0% off 0 30.000 36 off 0Communications 0 0 0Receiver 10.00 5.0% off 0 off 0 off 0Transmitter 30.00 5.0% off 0 off 0 off 0Payload 3 1.2 0Payload 1.00 20.0% 1.00 1.2 1.00 1.2 off 0Payload CDMU 1.50 20.0% 1.50 1.8 off 0 off 0Thermal Subsystem 0 0 0Miscellaneous 0.00 20.0% 0.00 0 0.00 0 0.00 0Total User Load 4.2 55.2 0Power Harness Losses 2.00% 0.084 1.104 0PCDU Conversion Losses 6.00% 0.252 3.312 0Total Power 4.54 59.62 0.00System Margin 20.00% 20.00% 20.00% 20.00%Total Power with Margin 5 W 72 W 0 W

Survival Night ModeDecentralised Science Centralised Science

Night Time operations

Power Budgets: Minimum power options

Assumptions:No electrical heaters required, autonomous switch on for survival

mode, high level command lines to CDMU with EIU switched off

94


Equipment Basic Power

Margin (%)

Basic Power [W]

Incl. Margin (%)

Basic Power [W]

Incl. Margin (%)

Basic Power [W]

Incl. Margin (%)

Data Management 0 18 0CDMU 15.00 20.0% off 0 15.00 18 off 0EIU 10.00 20.0% off 0 off 0 off 0Power Subsystem 1.2 36 0PCDU decentralised option 1.00 20.0% 1.00 1.2 off 0 off 0PCDU centralised option 30.00 20.0% off 0 30.000 36 off 0Communications 0 0 0Receiver 10.00 5.0% off 0 off 0 off 0Transmitter 30.00 5.0% off 0 off 0 off 0Payload 3 1.2 0Payload 1.00 20.0% 1.00 1.2 1.00 1.2 off 0Payload CDMU 1.50 20.0% 1.50 1.8 off 0 off 0Thermal Subsystem 0 0 0Miscellaneous 0.00 20.0% 0.00 0 0.00 0 0.00 0Total User Load 4.2 55.2 0Power Harness Losses 2.00% 0.084 1.104 0PCDU Conversion Losses 6.00% 0.252 3.312 0Total Power 4.54 59.62 0.00System Margin 20.00% 20.00% 20.00% 20.00%Total Power with Margin 5 W 72 W 0 W

Survival Night ModeDecentralised Science Centralised Science

Night Time operations

Power Budgets: Minimum power options

Centralised: Battery is 160kg for pole and 275kg for lower lander!Decentralised option = baseline

95


Additional Options

1. Receiver On Continuously :• Discarded, energy requirement driven by PCDU on

2. Status Sampling and Data Storage: • CDMU, PCDU and EIU have to be switched on for a short duration

3. Status Sampling and Communications: Full System On4. Timer to Switch Receiver On Regularly

Additional battery mass for sampling once per day and frequent near dusk/dawn and transmitting once per day implemented

Additional 1.5 kg for the polar lander Additional 2 kg for a lander between +/-83° latitude.

! For lander at the pole: no comms possible for ~ 2 weeks / month !

96


Risk Reduction

The MoonTWINS on surface mission can be designed very robust:

Undervoltage protection implemented in the PCDU to avoid total discharge of the battery. In case of one failure the battery will not be totally depleted, but everything will be switched off before. Solar array regulator (MPPT) can be implemented such, that the battery can be charged as soon as the SA receives solar flux (without booting the system!). As soon as battery voltage and solar array power reach a dedicated level the system will be booting up.For night time operations and as a back up, a timer is implemented. As part of the PCDU to minimise any power losses. This timer canbe used to switch on the receivers temporarily or to perform health checks of the lander.In the meantime thermal survival is guaranteed by RHU(s)Contingency modes can only be switched out of by ground intervention (lessons learned from Beagle2)

97


Thermal: Day Survival

• Very high temperatures during the day ( up to 395K)• 1380W/m2 if IR flux is radiated towards the radiators.

Conventional radiators absorb more heat than they can reject due to the high IR flux.

• Parabolic radiators required for lunar day survivalLower Lander: radiator placed on side never facing the sunPolar lander: radiator designed to minimise FV to Sun and indirect solar flux and IR flux onto the radiator from the reflector taken into account

High EmittanceRadiator Fin

Specular Parabolic Reflector

0.25 m20.23 m2-X side0.30 m20.27 m2+X Side

PoleEquator

98


Thermal: Night Survival

• Heating required to survive lunar nights• LHP with heat switches for night time survival to avoid heat loss

through radiators (ExoMars Rover design)• Thermal power requirement with heat switches

• RHU versus electrical heaterRHU lightest and most reliableIncrease of battery mass of 27kg to 60kg for electrical heater

Additional heating using RHUs at night Critical items located together in a “warm box”

5.9 W7.0 W-X side

6.2 W8.5 W+X Side

PoleEquator

99


Main Conclusions

• Proposed Landing Sites:1 lander at PEL: favourable illumination but Earth Occultations1 lander at lower latitude: 2 weeks night time, no Earth Occultations

• Night Time OperationsIf mass budgets allows for it, science can be continued during the night, limited to low power consumption instruments. However, instruments operating at night need low power decentralised datahandling systemFull system switch on only feasible for short durations several times in the night with limited communications

• ThermalParabolic Radiators needed to survive dayRHUs and heat pipes with heat switches needed for survival during the night

100


Power & RF Systems Analyses Synthesis D. Ruf

101


Outline

• Power Budget• Power System Architecture• Solar Array Design

For Equatorial or other LanderFor PEL Lander

• Battery Sizing

• RF System Architecture• Link Budgets

102


In-flight Power Budget

LIDAR-equipped lander : ~250W during RV (no eclipse)outside RV phase : ~210W during eclipses, ~185W outside eclipse

(Sun incidence effects resulting from in-orbit attitude constraints also accounted for)

System Assumptions and Constraints

Mode description Normal Mode Normal Mode Manœuvre RendezVous Safe Descent &outside during Mode Mode Mode Landingeclipses eclipses Mode

Duration - 40 minutes max 20 minutes max - - 10 minutes max

Data Handling CDMU 15.0 15.0 15.0 15.0 15.0 15.0EIU 15.0 15.0 15.0 15.0 15.0 15.0

Power PCDU 30.0 30.0 30.0 30.0 30.0 30.0

AOCS / GNC STR 5.0 5.0 0.0 5.0 0.0 0.0IMU 15.0 15.0 15.0 15.0 0.0 15.0MEMs 0.0 0.0 0.0 0.0 4.0 0.0NAVCAM 0.0 0.0 0.0 5.0 0.0 5.0LIDAR 0.0 0.0 0.0 50.0 0.0 50.0Main Engine 0.0 0.0 80.0 0.0 0.0 80.0250N thrusters 0.0 0.0 160.0 0.0 0.0 160.0RCS thrusters 10.0 10.0 10.0 10.0 10.0 10.0

RF TRSP 40.0 40.0 40.0 40.0 40.0 40.0

Harness 3.8 4.3 9.6 5.1 3.4 11.0

Total Bus 133.8 134.3 374.6 190.1 117.4 431.0Thermal Control 20.0 40.0 20.0 20.0 20.0 20.0

Payload 0.0 0.0 0.0 0.0 0.0 0.0

TOTAL 153.8 174.3 394.6 210.1 137.4 451.0System Margin 20%TOTAL with margins 184.5 209.1 473.6 252.2 164.8 541.2

Current Consumptions

103


Power System Architecture

• Unregulated 28V Power BusIn night mode: PCDU completely OFF to save powerDecentralized instrument connected to battery via safety switch (protection against short circuit failures and total discharge of the battery). PCDU power demand in night mode:1WPCDU provides autonomous power ON:

When solar array gets illuminated in the morningControlled by a timer

• Primary Power provided by solar arraysSolar array regulation by MPPT – due to high temperature variation of the SA during on-surface mission (up to 150°C)

104


Solar Array Design

• SA alignment depends on latitude of the landing siteLatitude has only minimum influence on SA size due to

low axial tilt of the Moon

• Alignment of SA panels in eastwards and westwardsdirections optimized to allow maximum utilizationof daylight and operation in day mode as long as possible

• Sizing of SA done to meet in-flight- and on-surface requirements with the same SA

• Application of 28% efficiency class GaAs solar cells (e.g. RWE 3G-ID2*)

105


SA for Equatorial Lander

0 0.5 1 1.5 2

x 106

0

50

100

150

200

250

300

Time [s]

Pow

er [W

]

Solar Array- and Load Power, OPTION3

Solar Array PowerLoad Power

Trade off betweenthree options:

• Option 1 provides poor utilization of daylight• Options 2 and 3 have similar perfomances, with smaller size possible for

option 3• Option 3 selected as baseline:

Allows day-mode operation withscience and communicationduring the whole lunar daySize: 1.2m² | 0.5m² | 1.2m²

2.9m² in totalMass: 14.5kg

106


SA for PEL Lander

0 0.5 1 1.5 2 2.5

x 106

0

50

100

150

200

250

300

350

400

450

Time [s]

Sola

r Arr

ay P

ower

[W]

Solar Array Power

SA1SA2SA3SA4SA5SA6all SA

• Lightning conditions different compared to equatorial lander:Sunlight can come from all around (360°)Maximum night duration 8 days

Requires different SA design

• Trade off between three different options:1) Four SA panels around the spacecraft body (quadrangular)2) Six SA panels around the spacecraft body (hexagonal)3) Deployable and rotating SA panel

Option 2 selected: (best size even withover-sizing of one panel due to in-flightrequirements:4x0.7m² + 2x1.1m² 5m² (24.5kg)

107


Battery Sizing

• Battery size is driven by the night-mode operation – even with a discharge power of a few watts only:

• During night the battery is discharged under extreme conditions:Very low temperatures (-20°C): Leads to an increased internal resistance of the batteryVery low discharge currents (<180mA)

New Li-Ion cell technology available in the near future (“ABSL LVF”) optimized for low temperatures: Provides around 108Wh/kg EOL (according to ABSL analysis). e.g.: around 20kg for 5W discharge

BUT: No heritage on the performance of the LVF cells under given conditions available yet. Performance needs to be confirmed by a dedicated life-cycle test

Normal Mode Normal Mode Maneouver RendezVous Safe Mode Desc/Landing Day Mode Night ModeIn Eclipse: No Yes Yes/No Yes/No Yes Yes No Yes

Power Demand [W]: 200 220 400 260 150 560 200 5Disc.Duration [min]: 0 40 20 80 40 10 0 20160

Bat Disc.Energ [Wh]: 0 147 133 347 100 93 0 1680

Flight Surface

108


RF System Architecture

X-Band selected as baseline:Wide selection of hardware (transponders, antenna, RF-switching units etc) and ground stations availableProvides high data rates with small antenna sizes

• Two low gain antennas (LGA) and one medium gain antenna (MGA)• Transponder from GAIA mission: Provides 5W RF output power

If higher data-rates arenecessary an additionalHPA can be added (e.g.to utilize excess powerfrom the SA during noon)

109


Link Budgets

Assumptions:• 15m X-band antenna ground station• Turbo Coding ¼ (like Bepi Colombo)• Increase of system noise temperature by up to 190K if the GS antenna is

pointed towards the center of the Moonleads to a reduction of the available data-rate by 50% (!)

(based on NASA JPL papers, should be verified by comparison with ESA SMART-1 link performance data)

Data Rates (with 5W GAIA transponder or additional HPA):• 5 W RF Power, Turbo ¼, LGA: 4.4 kbps• 5 W RF Power, Turbo ¼, MGA: 27.4 kbps• 10 W RF Power, Turbo ¼, MGA: 54.8 kbps• 15 W RF Power, Turbo ¼, MGA: 82.2 kbps• 20 W RF Power, Turbo ¼, MGA: 110 kbps• 25 W RF Power, Turbo ¼, MGA: 148 kbps

110


System Synthesis P. Regnier

111


Outline

• Landing Legs Analyses Synthesis• Avionics & AOCS• Spacecraft Configuration Updates

• Mass and Propellant Budgets

112


Landing Legs• Analyses & Trade-offs (SENER)

3 versus 4 legs trade-off : not conclusive without full re-design of lander configuration, but probably not determinant (6kg difference)landing legs design trade-off driven by requirement for multiple RV touch-and-go manoeuvres and two surface landings (optional hoping manoeuvre) :

– spring attenuation device for RV touch-and-go (~220gr)– enlarged footpads for robust and safe RV experiment (+300gr wrt normal)– double length crushable material for landing (+~70gr wrt single landing)

configuration and mass budget

Other SubSystems

113


PCDU

Distribution (LCL)I/FI/F…

Pyro…

Pyro

RF Communications

REGULATIONREGULATION

MEA

MEAMEA

Battery

IMU

STAR TRACKER

Attitude and orbit Control

Electrical Power

Power lines

Propulsion

NAV CAM

HTR Th

… …TCS

Control & Data Management System

TC decoderCPDUTFG

TC decoderCPDUTFG

TC decoderCPDUTFG

TC decoderCPDUTFG

OBTReconfigurationSGM

OBTReconfigurationSGM

ProcessorProcessor

PowerPower

ProcessorModule

ProcessorModule

1553 I/FPower 1553 I/FPower 1553

I/FPower 1553 I/FPower

RCS I/FRCS I/F

I/OI/OI/OI/O

PowerPower

Analog I/FAnalog I/FAnalog I/FAnalog I/F

Solar Array

Thermal Control

CDMU

EIU

SASSAS

Redunded Unit

Self redundant unitRegulated supply

Power lines TM/TC link

reg 28V

Payload 1

SpWi

Not redunded unit

MIL-1553B bus

Direct Commanding

Acquisitions

Int Bus

MMH MMH NTONTOMMH MMH NTONTO

LGA

1

MG

A

LGA

2

OSC

X-SSPA 2

DIPLX

DIPLX

RFDU

Deep SpaceTransponder 1

X-RxX-Tx

Deep SpaceTransponder 2

X-RxX-Tx

X-SSPA 1

MIL-STD-1553 B data bus

LIDAR

MEM Gyros

Mass Memory

Payload Processing Unit

PowerPower

BCRBCRBCRBCRBCRBCRBDRBDRBCRBDRBCRBCRBDRBDR

APRAPRAPRAPRAPRAPR

Radar Altimeter

Payload 2


PowerPowerPayload 3


PowerPower

Avionicscentralised architecture

CDMU : processor (ERC32 or LEON), Mass Memory,SafeGuard Memory, Reconfiguration Module, Transfer Frame GeneratorElectrical Interface Unit : I/Os to AOCS sensors, propulsion, thermal controlMil-1553B data bus + SpaceWire link to the camera

heritage = GAIA, Bepi-Colombo, or new generation LEO platforms redundancy approach :

ensure safe mode successessential uits for the on-surface mission are redunded

Other Subsystems

114


AOCSclassical gyro-stellar attitude estimation (1 STR, 1 3-axis IMU)no reaction wheels6dof thruster configuration (RV & attitude control, no ∆V)

eight 10N thrusters, no redundancyonly four are sufficient for attitude control (safe mode, descent & landing)

Safe Mode relies on Sun Acquisition Sensors and MEMs gyros (Earth comms with omnidirectional LGA coverage)

Other Subsystems

Thrusters

Wide Angle Camera (WAC)

Star Tracker Inertial Measurement

Unit

Radar Altimeter

Image Acquisition

Image Processing

Relative Navigation

Orbital maneuvers

Computation

Position Guidance

Attitude Guidance

Attitude Estimation

∆V Generation

∆V Estimation

Attitude Control

Rendezvous scenario

Landing scenario

Sensors

RDV/Landing S/W AOCS S/W

LIDAR

115


Assumptions & updatesSolar array configurations :

• polar lander : six-panel all-around solar array with two enlarged panels for the flight phase

• non-polar lander : two extreme cases : equatorial landing site or 83° latitude

• launch configuration : polar lander is on topparabolic radiators layout : as per sizing assessmentelectronic units layout as per thermal control preference

Spacecraft Configuration

116


Common central structure


900mm square tube withre-inforced columns

four external Eurostar 2000 tanks mountedon lower floors with struts

one central main engine and 8 ATV thrusters mounted laterally

117


Launch Configuration


Footpads externaldiameter tangential to

ST fairing allowedvolume.

Note : MLI notshown

Bepi-Colombo likelauncher adapter (struts supportingfour rigid corners)

118


Launcher Separation Sequence (initiated by S-F commands)


Clearance constraints fulfilledNote : MLI notshown

119


The three potential lander configurations


Clearance constraints fulfilledNote : MLI notshown

equatorial lander

polar lander

lander at 83deg latitude

electronic units thermal enclosure, with parabolic

radiators on each side

120


Equipment Layout


Note : MLI notshown

STRCDMUPCDU

BatteryP/L 1P/L 2

RFDUTranspondeur x2

EIUIMU

MEMs

SSPA x2

Lidar

Camera

MGA

121


AOCS Sensors FOV


Note : MLI notshown

LIDAR and camera

Sun sensors

122


Assumptions & updates∆V updates :

Hohman transfer case : 1686m/s at PM3, 1665m/s at FP (with representative gravity losses + 3% margins + launcher correction + navigation + orbit maintenance + descent rehearsal + RV allocation)Descent phase :

• previously assumed ∆V was 1900m/s (from previous ESA Lunar Landing studies), but obviously too optimistic (no margin, and not from a 150km orbit)

• based on latest optimal trajectory generation, assume 2000m/s to match a MSR-like vertical descent, 1950m/s without (includes ~3% margin + allocation for dispersion correction & retargeting)

Dry mass budget : + ~22kg since PM3 (propulsion, solar arrays, various points)ATV thrusters Isp misunderstanding : 300s Isp assumed before would need re-qualification (expansion ratio of 50 instead of 40). Otherwise would be 285s. Baseline shown with 300s Isp

Mass & Propellant Budgets

123


Situation with up-to-date assumptions, no payload

recall at PM3 : 17kg payload achievablewith 5% launch mass margin

now : not enough launch mass margin to support science

descent ∆V increase responsible for +100kg at launch mass.


Lower lander Upper landerestimated

massmaturity margin

estimated mass

maturity margin

Payload 0.0 kg 20.0% 0.0 kg 0.0 kg 20.0% 0.0 kgStructure 72.9 kg 13.5% 82.7 kg 67.3 kg 13.3% 76.3 kgPropulsion 119.4 kg 6.9% 127.7 kg 119.4 kg 6.9% 127.7 kgCDMS 11.0 kg 20.0% 13.2 kg 11.0 kg 20.0% 13.2 kgRadar altimeter or LIDAR 0.4 kg 20.0% 0.5 kg 8.3 kg 20.0% 10.0 kgTTC 15.0 kg 11.3% 16.7 kg 15.0 kg 11.3% 16.7 kgAOCS 8.3 kg 7.9% 9.0 kg 8.3 kg 7.9% 9.0 kgPower (incl PCDU) 18.4 kg 20.0% 22.1 kg 18.4 kg 20.0% 22.1 kgSolar Array 14.5 kg 20.0% 17.4 kg 24.5 kg 20.0% 29.4 kgThermal 18.1 kg 20.0% 21.7 kg 18.7 kg 20.0% 21.7 kgHarness 22.4 kg 10.4% 24.8 kg 24.6 kg 11.3% 27.4 kgLanding gear 36.4 kg 10.0% 40.0 kg 36.4 kg 10.0% 40.0 kg

TOTAL 375.8 kg 393.4 kgSystem Margin 20.0% 75.2 kg 20.0% 78.7 kgTOTAL with System Margin 451.0 kg 472.1 kgLanding DeltaV 2000 m/s 2000 m/sIsp 305 s 305 sPropellant Mass 431 kg 451 kgMass before landing (+2 kg AOCS) 882.12 kg 923 kgLunar Orbit Insertion and Maintenance DeltaV 914 m/s 914 m/sIsp 323 s 323 sPropellant Mass 299 kg 313 kgMass at Moon arrival (+4kg AOCS) 1181 kg 1236 kgLEOP and Lunar Transfer DeltaV 751 m/s 751 m/sIsp 323 s 323 sPropellant Mass 320 kg 335 kg

Masses in GTO (+4 kg AOCS) 1501.27 kg 1570.73 kg

3665.7 kg

SOYUZ adapter mass estimated mass 42.0 kg maturity mar 20.0% 50.4 kgLaunch mass 3122.4 kgLauncher capacity in GTO 3230.0 kg

Launcher capacity margin 3.4%

124


Recommendations for Improvement of mission mass performance

use Weak Stability Boundary Transfer :∆V savings ~65m/stransfer duration ~3 months

use an optimal lunar landing trajectory for the non-polar lander (no MSR-like vertical descent) : ∆V savings ~50m/s. This is acceptable because this lander supports the NPAL-like optical navigation technology soft landing demonstration, not the MSR-like LIDAR technology demonstrationreduce the required battery mass by operating the payload at night only on the non-polar lander

continuous P/L operation for the non-polar lander (including during 2-week long nights) quasi-continuous P/L operations on the polar lander (except possibly during 1-week long nights in winter, TBC from PEL characteristics)


125


Mission performance synthesis

Science payload mass allocation : ~ 10-15kgScience payload power allocation at night (non-polar lander only) : ~1-2W


PM3 update WSB transfer,

modified descent trajectory for the non-polar lander

Payload Mass Allocation (on each lander) 0kg 0kg

Payload Power Allocation at night

(for the non-polar lander) NA NA

Launch Mass Margin 3.4% 6.5%

0

5

10

15

20

25

0 1 2 3 4 5 6

Payload power allocation at night for the non-polar lander (W)

payl

oad

mas

s al

loca

tion

per l

ande

r (kg

)

targeting a 0% launch mass margin

targeting a 5% launch mass margin

126


Programmatics

Main Programmatics driversLaunch date: 2016 latest

Critical technologies to be demonstrated for MSR: automaticRendezvous, and soft/precision landing

• MoonTwins critical technologies should be mostly associatedwith these two topics.

Launch date: 2016 latest

ESA long-lasting investments in GNC technology for rendezvous and landing should benefit to MoonTwinsdevelopment

127


MoonTwins Technology pre-development

Pre-development with a bread board model.Parabolic radiator to minimise IR heating from lunar soil

Delta-development from ExoMars Rover thermal switches; included in MoonTwins phase B.

Thermal switches in conjunctionwith RHU usage

Thermal control

Delta-qualify and characterize the thrustersbehaviour (e.g. MIB, Isp) for MoonTwins GNC.ATV 250N thrustersPropulsion

Pre-development with a bread board model; includedin MoonTwins phase B.Landing legsMechanisms

High-fidelity spacecraft dynamics and scene imagingsimulator to validate all AOCS modes (nominal &

back-up); included in MoonTwins phase B.

GNC algorithms for rendezvous andfor soft / precision landing

Pre-development of a LIDAR model for sensorqualification and for navigation system validation in

dynamics conditions. LIDAR

Continue pre-development activities initiated for NPAL and implemented within the Aurora core-

programme.Navigation camera

Avionics

Technology validation approachTechnologySystem

Programmatics

128


ESA GNC investments used by MoonTwins (1/2)NPAL: TRL 4-5 demonstration of autonomous navigation for landing based on a vision camera, including the prototyping of the navigation camera (2006);

LiGNC : TRL 2-3 demonstration of Lidar-based GNC for rendezvous and landing (2006);

PLGTF : TRL 5-6 demonstration of NPAL in open-loop, through a drone demonstration (2008);

HARVD : TRL 4-5 demonstration of autonomous rendezvous GNC for both MSR-like and servicing missions, based on a navigation camera and/or Lidar (2008), including hardware-in-the-loop ground demonstration;

Programmatics

129


ESA GNC investments used by MoonTwins (2/2)

Hazard Avoidance : TRL 5-6 demonstration of NPAL in closed-loop, through a drone demonstration (2009);

LAPS : TRL 5-6 demonstration of LiGNC for landing in closed-loop (with ABSL Lidar BB, see below), through a drone demonstration (2009);

ILT : Design & Breadboarding of two different Lidar concepts (end 2008):• Landing Lidar (ABSL)• Rendez-Vous Lidar (Jena Optronik)

Programmatics

130


Rendezvous & Landing GNC validation approachSensors hardware qualification in flight representative dynamicsconditions using dedicated measurement benches (e.g. PLGTF, HARVD);

Intense system modelling activity (including Monte-Carlo statistics), using sensors & actuators numerical models correlated with measured data;

GNC flight software validated on numerical system simulator, then on Avionics real-time bench with hardware in the (closed) loop, then finally on the Flight models benches.

Dedicated GSE is foreseen to simulate in real-time the imaging of the scenes by the navigation optical sensors (Moon ground for landing, companion spacecraft for rendezvous).

Programmatics

131


Specific Test facilities & EGSE

Thermal test facility with Moon surface simulator (black body, thermally controlled)

Moon yard for landing & landed configuration test

Scene generator for replacing the Lidar and the Navigation camera during avionics validation & PFM qualification campaigns:• Electrical stimulators analog to STOS used for star trackers in

closed-loop test• Enhanced PANGU scene generator foreseen, enabling real-time

bench operation (target = 10 to 20 Hz)• Validation of this scene generator via PLGTF & HARVD campaigns

Programmatics

132


Master schedule

PROJECT PHASES & M ILESTO NES ξ ξ ξ ξKO PDR CDR FAR

Phase B

Phase C /D

Schedule system m argin

Launch Cam paign

PAYLO ADS DELIVERIES ξ ξSTM & EM PFM

SPACECRAFT SM SEQUENCE

PFM S tructure m anufacturing (S & N)

P ropulsion AIT on structure (S & N)

Landing legs Q M production

Spacecraft SM AIT (S & N)

Com posite SM AIT

Q M Landing test on spacecraft N

CENTRAL SO FTW ARE DEVELO PM ENT CSW Maintena Version V1 V2 V3

SPACECRAFT FLATSAT SEQUENCE

BB / EM / EQ M / PFM units production

F latSat A IT cam paign

SPACECRAFT PFM SEQ UENCE

Spacecraft S PFM integration

Spacecraft N PFM integration

Spacecraft S Therm al test

Spacecraft N Therm al tests

Spacecraft AIT com plem ent

Com posite m echanical & EMC test

2013 20142009 2010 2011 2012

Programmatics

133


Conclusions P. Regnier

134


MoonTWINS pre-phase A study main outcomes :Consolidation of MSR technology demonstration objectives

concerns only soft landing and RV technologies !soft landing technologies demonstrations - vision-based & LIDAR navigation, hazard avoidance, precision landing - confirmed to be possible and representative of MSR (especially the polar landerequipped with the LIDAR) autonomous RV technologies demonstration limited by mass constraints : optical detection, close proximity operations (same camera & LIDAR as for landing), GNC performance at contact through touch-and-go manoeuvre, but no capture mechanism, no RF proximity link

Consolidation of Moon Science perspectives for MoonTWINS Payload class (10 – 20 kg) : appears indeed attractive

focused on geophysics / Moon interior (seismometry network)radio-astronomy pathfinder experiment ?South Pole PEL characterisation (manned exploration perspective)

Conclusions

135


MoonTWINS pre-phase A study main outcomes :study team efforts to determine best estimates of payload mass / power allocations

mission analysis investigations (S-F launcher constraints, launch and transfer strategies, launch windows, lunar orbit stability…)on-surface system engineering (thermal control, system architecture and operations,…)power and RF systems engineering (solar array and battery sizing, RF link budgets…)propulsion system architecture and performanceGNC : descent trajectory optimisation, camera implementation, (RV GNC performance, optical detection range)spacecraft configuration : minimum mass solutions, fulfilment of launch, flight and landing applicable constraints (landing legs engineering)

best estimates of payload mass and power allocations : ~10-15kg per lander, ~1-2W for night operations on the non-polar lander

Conclusions

136


Conclusions on the MoonTWINS Mission Concepttechnical and programmatic feasibility is satisfactory at pre-phase A level, but several difficult points deserve further investigations (common to any lunar lander mission) :

descent trajectory optimisationon-surface thermal control performance and RHU implementationpropulsion system performance (more at Isp level than at thrust amplitude and modulation levels)battery (low temperature) and solar array (dust) sizing / performances stringent mass constraint (more specific to MoonTWINS)

despite its modest science payload mass / power allocations, the MoonTWINSconcept raised a strong interest among the scientific community, especially for its “network geophysics” valuestill a very attractive mission at technological level, especially for soft landing with two different demonstrations, and autonomous RV experiment with no major system impact

Conclusions

MSR Precursor Mission M-4 - European Space Agencyemits.sso.esa.int/emits-doc/ESTEC/AO-1-5589-RD12-M4.pdf · MSR Technology Demonstration Overview : LIDAR Introduction 2 4 6 8 10 12

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