CDF Study Report LUMETTOemits.esa.int/emits-doc/ESTEC/AO6791_RD1-Lumetto_v1... · 2011-06-08 ·...

LUMETTO

CDF Study Report: CDF-103(A) May 2010

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CDF Study Report

LUMETTOLunar Mapping and Exploration Technology &

Telecommunication Orbiter

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FRONT COVER

Study logo showing the Moon and a candle representing the details of the lunar surface

that will be highlighted following the mapping.

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STUDY TEAM

This study was performed in the ESTEC Concurrent Design Facility (CDF) by the following interdisciplinary team:

TEAM LEADER T. Bieler, TEC-SYE

CONFIGURATION R. Klotz, TEC-MCS POWER Gilles Beaufils, TEC-EPS

COST C. Runciman, TEC-SYC

PROGRAMMATICS/AIV

O. Brunner, TEC-TCC M. Gaido, TEC-TCC

GS&OPS F. Bosquillon de Frescheville, OPS-HSA

CHEMICAL PROPULSION

R. Schonenborg, TEC-MPC

OPTICS J. Perdigues, TEC-MMO

ELECTRICAL PROPULSION

D. Di Cara, TEC-MPE

MISSION ANALYSIS

M. Khan, OPS-GFA RISK A. Harrison, TEC-QQD

NAVIGATION M. Powe, TEC-ETN RADIATION H.Evans, TEC-EES

SYSTEMS S. Gerene, TEC-SYE D. de Lange, TEC-SY

Under the responsibility of:

B. Hufenbach, HSF-EA, Study Manager

M. Wittig, TIA-TF, Study Manager

With the scientific assistance of:

W. Carey, HSF-EA, Instruments and Customer Advisor

O. Mongrard, HSF-EA, Surface Scenarios

Z. Sodnik, TEC-ETC, Customer Advisor

With the technical support of:

O. Alvarez, TEC-ETC, Telecommunications Consultant

S. Kowaltschek, TEC-ECC, AOCS Consultant

J-F. Dufour, TEC-EDD, DHS Consultant

The editing and compilation of this report has been provided by:

A. Pickering, TEC-SYE, Technical Author

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This study is based on the ESA CDF Integrated Design Model (IDM), which is copyright © 2004 by ESA. All rights reserved.

Further information and/or additional copies of the report can be requested from:

B. Hufenbach ESA/ESTEC/HSF-EA Postbus 299 2200 AG Noordwijk The Netherlands Tel: +31-(0)71-5655075

Fax: +31-(0)71-5654499

[email protected]

For further information on the Concurrent Design Facility please contact:

M. Bandecchi ESA/ESTEC/TEC-SYE Postbus 299 2200 AG Noordwijk The Netherlands Tel: +31-(0)71-5653701

Fax: +31-(0)71-5656024

[email protected]

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TABLE OF CONTENTS

1 INTRODUCTION................................................................................... 9 1.1 Background .............................................................................................................. 9 1.2 Scope ........................................................................................................................ 9 1.3 Document structure ................................................................................................. 9

2 EXECUTIVE SUMMARY ...................................................................... 11 3 MISSION OBJECTIVES AND REQUIREMENTS....................................13

3.1 Objectives ................................................................................................................13 3.2 Surface Scenarios ................................................................................................... 14 3.3 Requirements ..........................................................................................................15

3.3.1 Science Requirements ......................................................................................15 3.3.2 Launch Requirements ......................................................................................15 3.3.3 Operational Requirements ...............................................................................15 3.3.4 Communication Requirements .......................................................................15 3.3.5 Orbit Requirements ..........................................................................................15 3.3.6 AOCS Requirements .........................................................................................15

4 INSTRUMENT PAYLOAD..................................................................... 17 4.1 Overview..................................................................................................................17 4.2 Core Optical Communications Payload..................................................................17

4.2.1 Ka-Band LUMETTO-Moon and LUMETTO-Earth Payloads..........................17 4.2.2 Optical Communications Payload ................................................................... 19 4.2.3 Navigation Payload..........................................................................................20 4.2.4 Optical Imaging Payload ................................................................................. 24

4.3 Additional Instruments.......................................................................................... 27 4.3.1 Mapping ...........................................................................................................28 4.3.2 Environment .................................................................................................... 29 4.3.3 Gravitmetry......................................................................................................30 4.3.4 Seismology ....................................................................................................... 33

4.4 Payload Suites ........................................................................................................ 35 4.5 Ground Segment .................................................................................................... 37

4.5.1 Ground-Based Ka-Band Receiver.................................................................... 37 4.5.2 Optical Ground Stations .................................................................................. 37

4.6 Lunar Surface Segment..........................................................................................38 5 MISSION ANALYSIS ........................................................................... 39

5.1 Requirements ......................................................................................................... 39 5.2 Assumptions........................................................................................................... 39 5.3 Mission Geometry Options .................................................................................... 39

5.3.1 Launchers and Initial Orbits ........................................................................... 39 5.3.2 Transfer Strategies...........................................................................................40 5.3.3 Phase 1 Orbit Options ......................................................................................40

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5.3.4 Phase 2 Orbit Options......................................................................................40 5.4 Baseline Design ......................................................................................................40

5.4.1 Trade-Offs ........................................................................................................40 5.4.2 Option 1.1.2.1 Mission Scenario....................................................................... 41 5.4.3 Option 1.3.2.1 Mission Scenario ......................................................................42 5.4.4 Option 3.2.2.1 Mission Scenario......................................................................42 5.4.5 Delta-V Budgets ...............................................................................................43

6 OPERATIONS ......................................................................................45 6.1 Assumptions ........................................................................................................... 45 6.2 Real Time Data Requirements ............................................................................... 45 6.3 Ground Station Configuration ...............................................................................46 6.4 Operations ..............................................................................................................46

6.4.1 Preparation ......................................................................................................46 6.4.2 Execution ......................................................................................................... 47

6.5 Remarks.................................................................................................................. 47 7 SYSTEMS............................................................................................ 49

7.1 System Requirements and Assumptions ...............................................................49 7.2 System and Mission Options and Trades ..............................................................49

7.2.1 Trade-Space .....................................................................................................49 7.2.2 Options Identification & Trade-Tree...............................................................49

7.3 Trade-space Qualitative Analysis...........................................................................49 7.4 Options Payload Mass Capacity .............................................................................50

7.4.1 Selected options ...............................................................................................50 7.4.2 Mass Budgets Determination .......................................................................... 51 7.4.3 Options Mass Summary................................................................................... 53

7.5 Trade-tree Rationale .............................................................................................. 54 7.6 Selected Mission Options....................................................................................... 56

7.6.1 Mass Budgets ................................................................................................... 56 7.6.2 Instrument Suites ............................................................................................ 57 7.6.3 Mission Objectives Compliance Matrix...........................................................58

8 LUMETTO SPACECRAFT ..................................................................... 61 8.1 Electrical Propulsion .............................................................................................. 61

8.1.1 Assumptions and Requirements ..................................................................... 61 8.1.2 Options............................................................................................................. 61 8.1.3 Baseline ............................................................................................................ 61

8.2 Chemical Propulsion ..............................................................................................64 8.2.1 Assumptions and Requirements .....................................................................64 8.2.2 Baseline ............................................................................................................64

8.3 Power ...................................................................................................................... 67 8.3.1 Chemical Propulsion Option ........................................................................... 67 8.3.2 Electric Propulsion Option ..............................................................................68 8.3.3 Remarks ...........................................................................................................69

8.4 AOCS.......................................................................................................................69

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8.4.1 Assumptions and Requirements ..................................................................... 69 8.4.2 Baseline Design................................................................................................ 69 8.4.3 Proposed AOCS Experiments .......................................................................... 70

8.5 Data Handling .........................................................................................................71 8.5.1 Assumptions and Requirements ......................................................................71 8.5.2 Baseline DHS Architecture...............................................................................71

8.6 Structures & Configuration.................................................................................... 73 8.6.1 Structures......................................................................................................... 73 8.6.2 Configuration ................................................................................................... 73

9 RISK ....................................................................................................77 9.1 Risk Management Policy – Objectives .................................................................. 77 9.2 Risk Management Policy – Project Goals.............................................................. 77 9.3 Precursor Mission Severity Categorization ........................................................... 78 9.4 Top Risk Log........................................................................................................... 79

10 PROGRAMMATICS ............................................................................. 83 10.1 Technology Development and Schedule:...............................................................84 10.2 AIV Approach.........................................................................................................84

10.2.1 Assumptions ....................................................................................................84 10.2.2 Model Philosophy ............................................................................................ 85 10.2.3 Schedule ........................................................................................................... 85 10.2.4 Programmatics/AIV - Impacts of Identified Options .....................................88

10.3 Conclusions ............................................................................................................90 11 COST................................................................................................... 93

11.1 Costing Assumptions ............................................................................................. 93 12 CONCLUSIONS ................................................................................... 95 13 REFERENCES ..................................................................................... 97 14 ACRONYMS ........................................................................................ 99

APPENDIX A - LUMETTO TRADE TREE .................................................. 101

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1 INTRODUCTION

1.1 Background

A number of technology demonstration missions for exploration were proposed in the frame of activities of the Inter-directorate Exploration Scenario Working Group (ESWG). To reach a better overview and understanding about missions of different nature, a selection was made among which there was the Lunar Mapping and Exploration Technology & Telecommunication Orbiter, or in short LUMETTO, jointly proposed by HSF-EA and TIA-TF.

DG-P (ESWG Chair) requested the CDF to carry out “quick assessments” in a “fast track fashion” and LUMETTO was the first study in this context and was completed by an interdisciplinary team of specialists from ESTEC and ESOC in only 4 sessions (1 per week), starting with a Kick-Off on 26 April 2010 and finishing with an Internal Final Presentation on 19 May 2010. The outcome of this study is an overview of the spacecraft concept compatible with the system requirements of the LUMETTO mission.

1.2 Scope

The study involved an assessment of LUMETTO mission concepts to provide information on the space and ground segments design as well as the trade-offs required for the mission. In addition the study provided an industrial cost, risk and programmatic analysis for the mission.

1.3 Document structure

The layout of this report can be seen in the Table of Contents. Due to the fast track nature of this study this document is a streamlined report providing an overview of the study concept, detailing each domain addressed in the study, either in specific chapters or in sections of the spacecraft chapter.

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2 EXECUTIVE SUMMARY A number of technology demonstration missions for exploration were proposed in the frame of activities of the Inter-directorate Exploration Scenario Working Group (ESWG). To reach a better overview and understanding about missions of different nature, a selection was made among which there was the Lunar Mapping and Exploration Technology & Telecommunication Orbiter, or in short LUMETTO.

The LUMETTO mission objectives are the following:

1. Demonstrate advanced capabilities for future exploration.

2. Prepare and support robotic and human lunar surface operations in two phases

a. Phase 1: Perform lunar mapping in preparation and support of robotic and human lunar surface exploration

b. Phase 2: provide communication and navigation services for robotic and early human exploration missions

3. Optimise utilisation of platform and mission for opportunistic scientific investigations

Different mission design concepts compliant with the mission requirements and cost frame were defined and traded against each other. Out of an initial set of more than 60 options three

• Soyuz-Hybrid-Single-Stage

• Soyuz-Electric-Single-Stage

• Shared-Ariane5-Hybrid-Single-Stage

were identified as the most promising scenarios and are proposed to be looked at in more detail. These options are identified and referenced on the LUMETTO Trade Tree in Appendix A.

Overall the LUMETTO mission is considered to be a pioneering and attractive mission, due to characteristics such as:

• The innovative use of the telecommunication payload combining optical mapping with optical communications.

• An advanced transfer strategy to reach the Moon by using chemical propulsion to pass the van-Allan-belts, go to Sun-Earth WSB point L2 and then heading back to the Moon with electric propulsion provide a very innovative character to the mission.

• A potential follow-up on ESA lunar mission experience such as Smart-1

• The support to future exploration activities by means of TLC service, surface mapping and Lunar environment characterisation.

• Providing valuable science contributions by means of the seismology instrument.

This mission could play a significant role in preparing European critical path contributions for future international exploration missions, through the provision of knowledge on the Lunar environment (mapping) and telecommunications services to

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Lunar surface assets. The mission fits very well into the proposed international lunar exploration scenario by the International Space Exploration Coordination Group (ISECG) by means of the proposed schedule: a launch date of 2017 and at least 10 years of high bandwidth data relay services with a focus on the Lunar South pole region.

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3 MISSION OBJECTIVES AND REQUIREMENTS

3.1 Objectives

In order to support future exploration missions, critical capabilities have been identified and categorised. The concept of technology demonstration missions has been introduced as possible near-term, self-standing milestones preparing Europe for making critical path contributions to global space exploration. A set of mission objectives have been identified to characterise a lunar mission that can provide the aforementioned contribution.

The mission objectives have been grouped into three categories, each containing a set of specific objectives that need to be met.

1. Demonstrate advanced capabilities for future exploration:

• Demonstrate advanced telecommunication technologies (optical) and delay tolerant telecommunication protocols

• Demonstration of tele-operations in combination with lunar surface payloads

• Demonstrate advanced electric propulsion for orbit transfer and orbit and attitude control

• Demonstrate advanced mission trajectories and associated operations

• Provide opportunities for demonstration of advanced generic technologies in areas such as power management.

2. Prepare and support robotic and human lunar surface operations in two phases:

• Phase 1: Perform lunar mapping in preparation and support of robotic and human lunar surface exploration: o Provide opportunities for accommodation of instruments of postponed or

cancelled national orbiter missions (DLR’s LEO, BNSC’s Moonlite, ASI’s Magia) such as for:

i. Global surface coverage with resolution < 1 m (stereo) and spectral resolution of < 10 m (range 0.2 – 14 µm)

ii. Subsurface sounding (global coverage) mapping (few metres deep with mm resolution; hundred metres deep with m resolution)

iii. Lunar space environment and dust characterisation o Support mission planning and operations of early lunar surface robots

• Phase 2: provide communication and navigation services for robotic and early human exploration missions: o Provide lunar data relay service o Provide time synchronisation services o Provide initial lunar surface positioning service.

3. Optimise utilisation of platform and mission for opportunistic scientific investigations:

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• Provide opportunities for accommodation of scientific instruments of failed national orbiter missions: o DLR’s LEO o BNSC’s Moonlite, o ASI’s Magia

• for measurement of: o gravity and magnetic field, o heat-flow o lunar seismology o lunar crater altimetry o nature and origin of lunar polar volatiles.

3.2 Surface Scenarios

In order to provide a reference Lunar surface scenario for LUMETTO CDF study the work performed in the frame of ISECG’s international Lunar exploration scenario has been used as a basis. The key characteristics of this scenario are:

• Phased approach to Exploration

• High level of mobility

as can be seen in Figure 3-1.

This led to the selection of the proposed launch date of 2017 and the focus on the South Pole region.

Figure 3-1: ISECG Scenario

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3.3 Requirements

The Mission objectives have been translated into a set of technical requirements, which are listed in the tables below:

3.3.1 Science Requirements

SCI-010 The image resolution of the South Pole shall be 20cm to 30cm

SCI-020 The image resolution of non South Polar locations shall < 1 meter.

SCI-030 The complete lunar surface shall be mapped during phase 1

SCI-040 The mission will not perform a controlled crash on the Lunar surface

3.3.2 Launch Requirements

Launch - 010 The launch date shall not be later than 2017

Launch - 020 The launcher shall be either a shared Ariane 5 or a Soyuz 2-1b-Fregat M

Launch - 030 The initial orbit for a shared Ariane 5 launch is a GTO

3.3.3 Operational Requirements

OPE - 010 The operations at the Moon shall be split into two phases

OPE - 020 Phase 1 shall be a lunar surface mapping phase

OPE - 030 Phase 2 shall support Data Relay for defined Assets on the lunar surface for at least 10 years

3.3.4 Communication Requirements

COM-010 The Demonstrator communications (payload) between LUMETTO and Earth shall make use of Ka-band and an optical link

COM -020 The Demonstrator communications (payload) between LUMETTO and the Lunar Surface provide the appropriate channels as defined by the lunar surface assets including Ka-band and optical link.

COM -030 The High Bandwidth Communications shall support tele-operation of defined lunar surface assets

COM -040 The High Bandwidth Communications shall support Data Relay of the defined lunar surface assets

3.3.5 Orbit Requirements

ORB - 010 During the mapping phase the selected orbit shall be a frozen orbit that has a theoretical Delta V budget for orbit maintenance of 0 m/s per year, in case no electrical propulsion is present

ORB - 020 The slant range at any point in the phase 2 orbit shall not exceed 10,000 km

3.3.6 AOCS Requirements

AOC - 010 The pointing accuracy of the orbiter shall be at least 0.1 degrees

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4 INSTRUMENT PAYLOAD

4.1 Overview

The instrument payload consists of a “core” telecommunications package which is then combined with a number of other instrument packages to optimise the payload capacity offered by a particular mission option. The “core” optical communications suite was defined to meet the primary mission objective to “demonstrate advanced capabilities for future exploration”. To address the other two high-level mission objectives related to ‘preparing for and supporting both robotic and human surface exploration operations” and “utilisation of the mission for opportunistic scientific investigations”, a selection of instruments were made from three cancelled European national missions, i.e. the German Lunar Exploration Orbiter (LEO) mission, the Italian MAGIA mission, and the UK MoonLITE mission.

4.2 Core Optical Communications Payload

The major building-blocks of the core communications payload consisted of the following packages:

• Ka-Band

• Optical communications

• Navigation: o Doppler Ranging o Precise Ranging & Time Transfer o Clock.

• Optical Imaging Payload

4.2.1 Ka-Band LUMETTO-Moon and LUMETTO-Earth Payloads

The Ka-band link budget is summarized in Table 4-1 and the masses and power for the Ka-band equipment is shown in Table 4-2. A block diagram for the Ka-band payload is shown in Figure 4-1.

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LUMETTO-Moon LUMETTO-Earth

Data Rate 300 Mbits/s Data Rate 300 Mbits/s

Tx Power 10 W Tx Power 100 W

LUMETTO Antenna 0.5 m LUMETTO Antenna 1.5 m

Ground antenna 0.5 m Ground antenna 5 m

Distance 5,520 km Distance 396,000 km

Rx Noise Temperature

260 K Rx Noise Temperature

260 dB

Eb/N0 9 dB Eb/N0 12 dB

Table 4-1: Ka-band link-budget

KaRx

Return Link

Forward Link

APM Tracking receiver

To Return Feeder Link

From Forward Feeder Link

ISL Tracking Antenna Figure 4-1: Ka-band Block Diagram


Repeater Repeater

Mass 15 kg Mass 15 kg

Power 40 W Power 270 W

Antenna Antenna



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Total Total



Table 4-2: Ka-band payload summary

4.2.2 Optical Communications Payload

A block diagram of the optical payload is shown in Figure 4-2: , a summary is given in Table 4-3 and the overall concept is shown in Figure 4-3. A schematic of the case with optical payload only (so no other payload included) is shown in Figure 4-4.

Figure 4-2: Optical payload block diagram

Lumetto-Moon

Mass 55 kg

Power 240 W

Lumetto-Earth

Mass 55 kg

Power 240 W

Total

Mass 110 kg

Power 480 W

Table 4-3: Optical payload summary

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LO

MOD

MOD

MOD

SYN

SYN

SYN

SYN

Modulators & Upconverters

LO

DEMOD

DEMOD

SYN

SYN

SYN

Demodulators & Downconverters

From Forward Feeder Link

To Return Link Feeder Link

LCT

Figure 4-3: Optical payload concept

Figure 4-4 Telecommunications (only) payload

4.2.3 Navigation Payload

The navigation payload, its characteristics and the related technologies are presented in this subsection.

4.2.3.1 Ultra Stable Oscillator

Galileo Rubidium Atomic Frequency:

• Manuf. - SpectraTime + EADS Astrium GmbH

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• Mass - 3.3 kgs

• Volume - 2.4 l.

• Power - 30 W

• ADEV 7×10-14 at 104 seconds

Figure 4-5: Galileo Rubidium Atomic Frequency

Passive Hydrogen Maser:

• Manuf. - Galileo Avionica + SpectraTime

• Mass - 18 kgs

• Volume - 45 l.

• Power - 70 W.

Figure 4-6: Galileo Avionica and SpectraTime Passive Hydrogen Maser

4.2.3.2 Doppler Ranging

• UHF PSK Coherent, USO synch.

• <1 Watt Transmitted Power

• Spread Spectrum 10 MHz Binary Chipping Rate

• Doppler Navigation (Transit) and/or Range to Satellite

• Navigation Data (ephemeris) and PRN modulated (PSK)

• Mass < 5 kg (based on Instrument Presentation input)

• Power < 20 W (based on Instrument Presentation input).

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4.2.3.3 Navigation Accuracy

• Radio navigation is used in conjunction with other aids (INS, Odometer, Terrain Height Map, User Minature Atomic Clock)

• Simulations from BAE/UCL/UoL Limited Applicability (orbits)

• Single-Satellite Transit-Like Doppler Navigation

• Planet-fixed coordinate frame estimation included

• Orbit determination aided by Satellite-Earth link

• Study Long-Range (Transit-Like) Accuracy Requirement: 200 m 2D 1 σ.

• Study Short-Range (WiMax Network) 10 m 2D 1σ.

Figure 4-7: Navigation accuracy (Source: BAE/UCL/UoL Contract No:

20763/07/NL/HE)

4.2.3.4 Related Technologies

Chip-Scale Atomic Clocks Enables one-way Ranging.

-80

-60

-40

-20

0

20

40

60

80

51544 51549 51554 51559 51564 51569

Time (MJD)

ENU

com

pone

nt e

rror

(m)

dE(m)dN(m)dU(m)

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Figure 4-8: Chip-Scale Atomic Clocks Enables one-way Ranging

• 2007 Symmetricom Prototypes o Volume - 15 cm3 o Power - 125 mW o ADEV approximately 10-11 at 1000 seconds

• Evolving to o Volume - 1 cm3 o Power - 30 mW

• New GNSS Pilot Signals – Track C/N0 10 dB-Hz.

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Figure 4-9: New GNSS Pilot Signals – Track C/N0 10 dB-Hz

Figure 4-10: Sketch of the combination of the Telecommunications and Navigation payload

4.2.4 Optical Imaging Payload

Characteristics:

• Use of the Optical Data Terminal for Imaging by reusing the Telescope

• Additional Imaging CCD Sensor of 4 Mega Pixel

• Ground Resolution < 30 cm

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• Data Rate 370 Mbit/s for 3 spectral channels and 8 bit ADC.

It should be noted that the mapping resolution is directly proportional to the orbital height above the lunar surface, e.g.:

• At 50 km around 22 cm

• At 200 km around 88 cm.

For optical mapping, it is proposed to use the Optical terminals Acquisition Sensor as an Imaging Sensor, in which case the following additional components are required:

• An additional CCD or APS Sensor

• Associated drive electronics

• Signal processing electronics.

These additional components are indicated in Figure 4-11.

Figure 4-11: Architecture of the optical mapping instrument (mono)

Note that the Fine Pointing Actuator (FPA) allows steering of the imaging beam within the telescope Field of View, and an additional FPA allows stereo imaging (see Figure 4-12).

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Figure 4-12: Architecture of the optical mapping instrument (stereo)

The components required for mono imaging are:

• Optical beam splitter

• Multi-spectral CCD

• Signal processing electronic.

The components required for stereo imaging are:

• A second Fine Pointing Assembly

• Optical beam splitter

• Two sets of multi-spectral CCD’s

• Signal processing electronics.

Table 4-4 summarizes the mass and power requirements for the mapping instrument package. The architecture of the telecommunications, navigation and mapping package is shown in Figure 4-13.

Mono Mapping Stereo Mapping

Mass 5 kg Mass 8 kg


Table 4-4 Mapping payload summary

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Figure 4-13: Architecture of the telecommunications, navigation and mapping package

4.3 Additional Instruments

In order to address mission objective 2 (Prepare and support robotic and human lunar surface operations) additional instruments have been selected. The additional instrument selection was based on combining the available candidates into an overall complementary set. Thus, national mission payloads relating to optical mapping were not considered, as this capability was included in the core optical communications package described earlier. As part of this process, the remaining candidate instruments were also classified into four distinct categories (see Table 4-5). As can be seen in this table the first three categories address mission objective 2 fully, whereas the fourth category (seismology) is more of a scientific nature.

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Instrument Package/Suite

Mapping Environment Gravimetry Seismology • Synthetic

Aperture Radar (LEOSAR)

• UV Spectrometer (USMI)

• Thermal IR Spectrometer (SERTIS)

• Radiation Monitor (RadMo)

• Dust Particle Detector (LEOPARD)

• Neutral Atoms Detector (ALENA)

• Energetic Particle Spectrometer (RADIO)

• Magnetometer (LunarMag)

• Accelerometer (ISA)

• Subsatellite (active)

• Penetrators

Table 4-5: Instrument packages/suites

These packages were then combined together in various ways (A through N) to provide a range of total payload mass from a minimum “communications only” package (as described in section 4.2.2) to a set that included all of the instruments above together with the core communications package. An instrument package or instrument suite may or may not be compliant with the mission objectives. Table 4-8 in section 4.4 shows the compliance table of the instruments suites with respect to the mission objectives.

Each of the instruments in Table 4-5 are summarised below.

4.3.1 Mapping

4.3.1.1 Synthetic Aperture Radar

LEOSAR measures topographical and radar image data of the lunar surface and meter-deep subsurface structure by using an active polarimetric microwave instrument operating at 23 cm wavelength (L-band). It is independent of solar illumination and can therefore observe regions where insufficient illumination restricts the use of optical systems, leading to the capability of global mapping of the moon surface. The SAR signal is moreover able to penetrate deep (several tenths of meters) into subsurface regions providing unique tomographic information about the regolith. SAR also delivers high-resolution (both geometric and radiometric) wide-swath images for global mapping complementing the optical images from high resolution cameras (e.g. HRSC). The total mass of the instrument is below 23 kg and the average power consumption is below 100W.

4.3.1.2 UV Spectrometer

Ultraviolet Spectral Mapping Instrument (USMI) is a multispectral imaging camera designed for planetary surface mapping in the ultraviolet. It provides complementary observations to optical and infrared imaging and, thus, can be used to

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constrain the mineralogical, petrologic, and chemical composition of geologic material or, in combination with observations in the visual and IR, to determine it in much greater detail. The main components of USMI are the camera optics, transmission and reflection filters covering the 200-400 nm range in several bands, and CCD TDI detector heads. The total mass of the instrument can be as low as 5 kg. A phase-A study was performed in the frame of the German Lunar Exploration Orbiter.

4.3.1.3 Thermal IR Spectrometer

SERTIS is an imaging spectrometer for the thermal infrared (TIS) with an integrated µ-Radiometer (TIR). The spectrometer operates at wavelengths of 7-14 µm with a spectral resolution of less than 200 nm; the µ-Radiometer operates at 7-40 µm. It is the first hyperspectral instrument for the thermal infrared to be operated in orbit around the Moon. The spectral range covered allows the direct identification of plagioclase minerals which have characteristic features here. It complements existing and planned near infrared observations. SERTIS will map the mineralogy of the Moon with a lateral resolution of better than 100m. The radiometer allows mapping the thermophysical properties of the lunar regolith with high accuracy. SERTIS is based on the development of the Mercury Radiometer and Thermal Infrared Spectrometer (MERTIS). The instrument has a mass of <3.5 kg and a power consumption of <15 W. The lateral resolution will be <100 m. Because SERTIS is an adapted derivate of the MERTIS experiment that is already under development [Hiesinger and Helbert, 2008], the technical readiness level (TRL) is high and costs can be reliably predicted.

4.3.2 Environment

4.3.2.1 Radiation Monitor

Radiation Monitor (RadMo) measures the radiation environment of both charged (Ions and electrons) and neutral (neutrons and gammas) high-energy particles over a wide range of energies (up to 200 MeV/nuc for ions). In addition, RadMo provides rudimentary Gamma-spectroscopy and thermal neutron detection. The latter has spatial resolution limited to roughly the height of the orbit which already allows identification of proton-rich (water-rich) deposits. RadMo consists of three sensors and an electronics box. The sensors cover charged particles and gamma rays by a dE/dx-vs total E solid-state telescope with scintillator calorimeter. Non-thermal neutrons are detected with a combination of several hydrogen-rich plastic scintillators, thermal neutrons by a novel and extremely light-weight solid-state-based technique. Total RadMo required resources are 6-7 kg (depending on accommodation) and 12 W (incl. 20% margin).

4.3.2.2 Dust Particle Detector

The Lunar Exploration Orbiter Particle Detector (LEOPARD) determines the surface composition of bodies without an atmosphere such as the Earth’ moon by measuring with unchallenged accuracy the mass, speed vector, electrostatic charge, and the chemical composition of individual surface ejecta particles populating dust clouds around such bodies. It’s trajectory sensor determines the grain’s speed vector by a novel charge-sensitive wire sensor while passing the instrument’s charge-sensitive grid sensor with an uncertainty of below 0.1%, while the instrument’s reflectron-type mass spectrometer measures the grain’s composition with a mass resolution of >200. It is

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able to detect grains with speeds between 1 and 100 km/s and sizes between 100 nm and 100 µm. The instrument’s mass can be as low as 4 kg. Instrument prototypes have been built and successfully tested.

4.3.2.3 Neutral Atoms Detector

ALENA is an energetic neutral atoms detector which measures the flux and velocity of neutral atoms and particles. Exploration objective is to characterise the space environment around the Moon. Science objective is to improve understanding of the interaction between the solar wind and the lunar surface.

4.3.2.4 Energetic Particle Spectrometer

The RADIO instrument measures the flux of protons, neutrons, alpha-particles and electrons around the spacecraft. Such data is important to understand the levels of radiation to which human explorers will be subject (outside of the Van Allen belts).

4.3.3 Gravitmetry

4.3.3.1 Magnetometer

The Lunar Fluxgate Magnetometer (LunarMag) measures the magnetic field 30 to 50 km above the lunar surface. A detailed and global map of the lunar vector magnetic field will give far reaching information and insights into the structure and dynamics of the lunar crust.

An earlier lunar dynamo, impact generated magnetisation or transient magnetic fields could contribute to the lunar magnetic field. In regions with strong magnetic field anomalies a local magnetosphere can be created. LunarMag shall measure the magnetic field on two spacecrafts This allows the application of gradient measurements to separate the external contributions from the crustal magnetic field. Close cooperation and synergies with gravity field measurements is anticipated. The total mass of the instrument is 1.6 kg per satellite. The TRL is 9; similar instruments are flown on Rosetta, Venus Express and Themis.

4.3.3.2 Accelerometer

The ISA accelerometer measures the non-gravitational forces acting on the spacecraft. Data characterises the forces affecting the attitude of the spacecraft in its orbit around the Moon, and could lead to an improved understanding of the internal structure of the Moon.

4.3.3.3 Sub-Satellite

The purpose of this payload component is to track the sub-satellite for the scientific objective of measuring the gravity field of the Moon. Options would be to re-build and fly the GRACE/GRAIL RF equipment or use (already existing) optical coherent technology. An example of the GRACE/GRAIL approach is shown in Figure 4-14.

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Figure 4-14: GRACE/GRAIL tracking approach

In the case of use of a passive sub-satellite, a pulsed-laser radar would be utilised, whereas for an active sub-satellite, two approaches are feasible:

• Ka-Band Ranging (KBR) as for GRACE: o 10 kg 50 W estimated

• Optical Doppler Tracking: o Use coherent optical terminal which is baseline for EDRS o Additional required signal processing electronics only.

The additional components with respect to the Core Communications Payload as discussed in section 4.2 are shown in the architecture diagram in Figure 4-15.

Figure 4-15: Architecture diagram of the additional components

Characteristics:

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• Pulsed Nd:YAG laser

• Same receiver as for communication

• Ranging processing electronic

• Mass 5 kg

• Power 50 W.

Figure 4-16: Block diagram of the complete communications package

The estimated characteristics of the sub-satellite are summarized in Table 4-6 and a conceptual drawing is shown in Figure 4-17.

Sub-Satellite

Bus 20 kg 65 W

Payload 6 kg 50 W

Propellant 4 kg

Total 30 kg 115 W

Table 4-6: Sub-satellite characteristics

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Figure 4-17: Sub-satellite visualisation

4.3.4 Seismology

4.3.4.1 Penetrator

The Penetrator payload requires a UHF receiver onboard the LUMETTO spacecraft. The block diagram for the penetrator receiver is shown in Figure 4-18. The receiver mass is estimated at 5 kg, and the receiver power at 20 W. Figure 4-19 shows the communications payload as discussed in section 4.2 including the penetrator UHF receiver.

Figure 4-18: Penetrator receiver payload block diagram

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Figure 4-19: Telecommunications, Navigation, Mapping and Seismology payload

The penetrator concept is shown in Figure 4-20.

Figure 4-20: Penetrator concept

The characteristics of the penetrator payload are:

• UHF communications

• 30 kbits/day

• Mass ~40 kg

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• Power ~0.06 W.

4.4 Payload Suites

Table 4-8 shows how each of the payload instrument suites/packages map to the LUMETTO mission objectives.

Option A B C D E F G H I J K L M N

IncludesComms Only

Comms & Navigation

Comms & Navigation & Optical Mapping

Comms & Navigation & Full Mapping

Comms & Navigation & Optical Mapping & Environment

Comms & Navigation & Full Mapping & Environment

Comms & Navigation & Seismology

Comms & Navigation & Seismology & Environment

Comms & Navigation & Gravimetry

Comms & Navigation & Gravimetry & Environment

Comms & Navigation & Gravimetry & Seismology

Comms & Navigation & Gravimetry & Full Mapping

Comms & Navigation & Gravimetry & Full Mapping & Environment

Comms & Navigation, Full Mapping, Environment, Gravimetry, Seismology (3 penetrators)

TOTAL PACKAGE MASS 199.5 220.5 231 268.2 249.9 287.1 417.75 436.65 269.43 288.33 466.68 317.13 336.03 485.28

Table 4-7: Instrument Suites Options Masses

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Comms

Comms /

Nav

Mappin

g

Enviro

nmen

t

Gravim

etry

Subsa

ts

Seismolo

gy

(I) Demonstrate advanced capabilities for future exploration

X X1.1 – Demonstrate advanced telecommunication technologies (optical) and delay tolerant telecommunication protocols X X1.2 – Demonstration of tele-operations in combination with lunar surface payloads X X

1.3 – Demonstrate advanced electric propulsion for orbit transfer and orbit and attitude control (tbc)

1.4 – Demonstrate advanced mission trajectories and associated operations (tbc)

1.5 – Provide opportunities for demonstration of advanced generic technologies in areas such as power management

(II) Prepare and support robotic and human lunar surface operations in two phases

X X X X

Phase 1: Perform lunar mapping in preparation and support of robotic and human lunar surface exploration

X X

2.1 – Provide opportunities for accommodation of instruments of failed national orbiter missions (DLR’s LEO, BNSC’s Moonlite, ASI’s Maggia) such as for

X X2.2 - Global surface coverage with resolution < 1 m (stereo) and spectral resolution of < 10 m (range 0.2 – 14 µm)

X2.3 - Subsurface sounding (global coverage) mapping (few metres deep with mm resolution; hundred metres deep with m resolution)

X2.4 - Lunar space environment and dust characterisation X2.5 – Support mission planning and operations of early lunar surface robots X XPhase 2: Provide communication and navigation services for robotic and early human exploration missions

X2.6 - Provide lunar data relay service X X2.7 – Provide time synchronisation services X2.8 – Provide lunar initial lunar surface positioning service X(III) Optimise utilisation of platform and mission for opportunistic scientific investigations

X X X

3.1 – Provide opportunities for accommodation of scientific instruments of failed national orbiter missions (DLR’s LEO, BNSC’s Moonlite, ASI’s Maggia) e.g. for measurement of gravity and magnetic field, heat-flow, lunar seismology, lunar crater altimetry, nature and origin of lunar polar

l til

X X X

Table 4-8 : Instrument Suites vs Mission Objectives

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4.5 Ground Segment

Two types of ground-segment facilities are required to support the LUMETTO mission, i.e.:

• Ka-Band Receivers

• Optical Ground Stations.

4.5.1 Ground-Based Ka-Band Receiver

The link budget is made for a 5m Earth station antenna.

The procurement cost for such a station is estimated to be below 1 M€. Installation at already existing ESA sites (e.g. Australia, Spain, Argentina) is envisaged.

4.5.2 Optical Ground Stations

Identified potential stations for mission support are:

• The 1m OGS of ESA based in Tenerife (Figure 4-21)

• The 17m MAGIC telescope (as a “photon bucket”) in La Palma (Figure 4-22).

Figure 4-21: The OGS facility in Tenerife

Figure 4-22: The MAGIC telescope in La Palma

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4.5.2.1 Optical Ground Station Receiver

A generic block-diagram receiver is shown in Figure 4-23.

Figure 4-23: Generic block-diagram

4.6 Lunar Surface Segment

For the optical terminal on the lunar surface to support either robotic or human operations, the following terminal high-level requirements have been identified:

• The telescope needs to be covered by an inflatable radome, or an otherwise protective cover to mitigate contamination by dust.

• The terminal would only require a power supply and a thermal radiator o Additional mass ≈ 25 kg. o Total terminal mass ≈ 80 kg.

Figure 4-24: Potential lunar optical terminal

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5 MISSION ANALYSIS The task for the Mission Analysis for the LUMETTO CDF study was to analyze the overall mission geometry from launch till EoL. The analysis included launcher selection, initial orbit selection, analysis for three different propulsion principles (SEP, Chemical and a hybrid solution), phase 1 orbit selection and phase 2 orbit selection.

5.1 Requirements

The requirements applicable to mission analysis are summarized in Table 5-1.

Identifier Requirement Description SCI-030 The complete lunar surface shall be mapped during phase 1 SCI-040 The LUMETTO mission will not perform a controlled crash on the Lunar surface Launch – 010 The launch date shall not be later than 2017 Launch - 020 The launcher shall be either a shared Ariane 5 or a Soyuz 2-1b-Fregat M Launch - 030 The initial orbit for a shared Ariane 5 launch is a GTO OPE - 010 The operations at the Moon shall be split into two phases OPE - 020 Phase 1 shall be a lunar surface mapping phase OPE - 030 Phase 2 shall support Data Relay for defined Assets on the lunar surface for at least 10

years ORB - 010 During the mapping phase the selected orbit shall be a frozen orbit that has a theoretical

Delta V budget for orbit maintenance of 0 m/s per year, in case no electrical propulsion is present

ORB - 020 The slant range at any point in the phase 2 orbit shall not exceed 10,000 km

Table 5-1: LUMETTO Requirements applicable to Mission Analysis

5.2 Assumptions

The following assumption has been used in the mission analysis for LUMETTO:

• SEP engine is the PPS5000 Hall Effect Xenon Thruster at: 5 kW, 300 V, 276 mN and Isp of 1763 s.

5.3 Mission Geometry Options

5.3.1 Launchers and Initial Orbits

The launchers considered are the shared Ariane 5 into GTO and a dedicated Soyuz, as dictated by the mission requirements, Table 5-1. For the Soyuz launch the initial orbit options were limited to GTO and HEO 200,000 km. LTO was not considered as initial orbit for LUMETTO due to low flexibility in terms of launch window as well as a lower launch mass.

The Soyuz launch capacity to GTO is 3070 kg and to HEO 2426 kg. Including a 1194SF adapter of 110 kg results in a maximum wet mass for the LUMETTO S/C of 2960 kg and 2316 kg, for GTO and HEO respectively. For the shared Ariane 5 two sharing cases were considered (LUMETTO in both cases being the secondary passenger). For the first case the primary passenger mass is fixed at 6000 kg and for the second case the LUMETTO S/C mass is determined by taking the largest payload suite.

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5.3.2 Transfer Strategies

Transfer orbits considered during the study include a direct transfer, S-E WSB transfer as well a “Smart-1”-like transfer, which is a delta-V optimized solution for the SEP case.

5.3.3 Phase 1 Orbit Options

For the mapping phase two candidate orbits have been identified that fulfil the ground resolution requirements for the mapping and also provide global coverage:

• frozen polar LLO, ca. 45 x 195 km, periselenium over the South Pole, no station-keeping required

• 50 x 50 polar LLO, requires ~150 m/s station-keeping per year.

5.3.4 Phase 2 Orbit Options

For the phase 2 orbit the design is driven by:

• maximum availability and coverage over the South Pole

• maximum slant range of 10,000 km

• minimum maintenance delta-V cost

Furthermore an analysis was done considering a constellation of two S/C in order to have continues availability (with the second orbiter being from a different future mission). These requirements resulted in the identification of the three phase 2 candidate orbits as shown in Table 5-2.

Phase 2 orbit

Transfer delta-V from 50x50 km, 90 deg phase 1 orbit

Transfer delta-V from Frozen LLO, 90 deg

phase 1 orbit Orbital Period

South Pole Pass Duration

3500 x 3500 km, 39.23 deg

1154 m/s 1081 m/s TBD TBD

700 x 8100 km, 51 deg 954 m/s 980 m/s 12 h ~8.5 h 300 x 5600 km, 51.7 deg 1002 m/s 1023 m/s 8 h ~4.5 h

Table 5-2: Phase 2 candidate orbits including the required delta-v for transfer from phase 1 to phase 2

5.4 Baseline Design

5.4.1 Trade-Offs

The transfer time from GTO to the phase 1 orbit with a “Smart-1”-like transfer strategy with the PPS5000 (set at 276 mN) and an initial mass in the order of 3 tons would take in the order of 3 years. Considering Smart-1 suffered a 8.5% in power degradation due to the Van Allen belts during its 18 months transfer, the “Smart-1”-like transfer is discarded. Nevertheless the transfer duration could be reduced down to 2 years if initial mass is optimised down to 2 tons.

As the mission options based on a chemical propulsion system resulted in limited payload masses, a hybrid solution or electrical solution (from HEO) were selected as most promising (in terms of maximum payload mass capacity) candidate transfers.

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During the LUMETTO CDF study the hybrid transfer strategy consisted of a chemical transfer to the S-E WSB from where the SEP takes over. With this transfer strategy the Van Allen belt pass duration is minimized and propellant mass is saved by using the SEP for the rest of the mission, resulting in an increase in maximum payload mass capacity. This transfer strategy has not been analyzed within ESA, meaning no software is available to quantify its performance. It is also likely not the optimal solution for a hybrid system. The optimal solution is likely to be a bi-elliptic rather than a WSB transfer.

For the phase 1 orbit the 50 x 50 km frozen polar orbit is selected as all three final LUMETTO S/C options have a SEP (two hybrids and one fully electric). The consequence of this selected orbit compared to the frozen polar orbit is a better ground resolution at the cost of 150 m/s delta-v per year. For electric propulsion the propulsion mass penalty is considered sufficiently low.

The phase 2 orbit for all three options is selected to be the 700 x 8100 at 51 deg inclination as it has a higher ground coverage, longer South Pole pass duration as well as lower transfer delta-v.

Considering the above described trade-decisions the performance of several options (shown in the overall mission trade-tree shown in Appendix A) were analyzed and the three most promising options were selected as study end-result which will be described in the next paragraphs. Note that the numbering system used in the Appendix identifies the options.

5.4.2 Option 1.1.2.1 Mission Scenario

Option 1.1.2.1(Appendix A) is based on a dedicated Soyuz launch into GTO, using a hybrid transfer. The chemical propulsion system performs the burn to reach the S-E WSB from where the SEP takes over to perform the rest of the mission. The overall mission concept is shown in Figure 5-1. The transfer time from the S-E WSB to the phase 1 orbit could not be quantified due to the lack of knowledge on this transfer type but the target is set at 180 days. The delta-v budget is shown in Table 5-3.

Figure 5-1: Sketch of the mission geometry for option 1.1.2.1

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The second remaining option is option 1.3.2.1 (Appendix A) which consists of a dedicated Soyuz launch into a HEO with an apogee of 200,000 km. This option is a fully electric option using resonances and lunar gravitational capture as transfer strategy. The transfer time is estimated at approximately 200 days with a total delta-v of 3528 m/s. The mission geometry concept is sketched in Figure 5-2 and the delta-v budget is again given in Table 5-3.

Figure 5-2: Sketch of the mission geometry for option 1.3.2.1


The third option is option 3.2.2.1, having similar orbit geometry as option 1.1.2.1 (paragraph 5.4.2, Figure 5-1). This option uses a shared Ariane 5 launch instead of the Soyuz launch used for option 1.1.2.1. The delta-v budget is equal to the 1.1.2.1 case and is again given in Table 5-3.

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5.4.5 Delta-V Budgets

Table 5-3 summarizes the delta-v budget for the three above described remaining mission options for LUMETTO.

Trade option 1.1.2.1 1.3.2.1 3.2.2.1 Launcher Soyuz Soyuz Shared A5 Initial orbit GTO HEO 200k GTO

Propulsion/staging Hybrid, Single Stage

Electrical, Single Stage

Hybrid, Single Stage

Transfer strategy Chem. to S-E WSB, EP for rest

Resonances & gravitational capture

Chem. to S-E WSB, EP for rest

∆V Precapture transfer [m/s] 763.00 600.00 763.00∆V Capture to phase 1 orbit [m/s] 2100.00 1108.00 2100.00∆V Phase 1 maint. [m/s] 300.00 300.00 300.00∆V Phase 1 to phase 2 [m/s] 2000.00 2000.00 2000.00∆V Phase 2 maint. [m/s] 0.00 0.00 0.00∆V EOL [m/s] 120.00 120.00 120.00∆V Total chemical [m/s] 763.00 n.a 763.00∆V Total electrical [m/s] 4520.00 3528.00 4520.00Transfer time to phase 1 [days] 180 (target) 200.00 180 (target)

Table 5-3: Delta-v budget for the three remaining LUMETTO mission options (chemical burns in green, electrical in red)

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6 OPERATIONS

6.1 Assumptions

The following operations phases have been identified:

• Preparation

• Training

• LEOP & transfer to Phase 1 Orbit,

• Remote sensing

• Data Relay (Phase 2) o before Human Lunar Return (HLR) o after HLR

• EOM.

The following Multi Ground station network has been identified:

• Optical

• X/K band.

Uplink

• platform and payload control data

• file transfer to be relayed

• Real time data to be relayed from Earth

• Data rates: 5 -10 Mbit/s

Downlink

• RT data from the Moon

• S/C telemetry

• relayed file transfers

• Data rates: 300 Mbit/s up to 1 Gbit/s.

Before HLR routine during Normal working day; after HLR shift work.

6.2 Real Time Data Requirements • For completeness Real Time data should be relayed: o from earth to Moon surface and back o from Moon surface to Moon surface and back o from Lunar orbiter to moon surface and back.

• Interference of signals arriving at the spacecraft should be avoided (different channels or time management)

• No force feedback loop is possible from Earth, due to the fact that the response time needs to be at maximum 100 ms. It would be possible from lunar surface to

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lunar surface or from lunar orbit to lunar surface. Without force feedback no precise, dynamic or delicate remote operation is possible

• Real time or near real time session duration shall be 15 minutes minimum

• The data rate is at minimum 1 Mbit/s from operator to Robot and 4 Mbit/s from robot to operator per robotic asset.

6.3 Ground Station Configuration

To satisfy the above stated requirements and assumptions the following ground station configuration has been selected:

• For LEOP and lunar the transfer as well as after HLR three 15 m stations in X-Band are selected: KRU, VIL, PER

• Before HLR, for normal working day operations two 35 m DSA (NNO & MAL), ensuring a 2 to 4 hours daily pass, in X and K-Band are selected

• After HLR three 5 m K-Band (CEB, NNO & MAL), ensuring contact whenever spacecraft is not occulted by Moon, are selected for downlink. The Uplink is covered by Three 15 m station in X-Band (KRU, VIL, PER)

• For the optical link ideally 5 optical stations are required, spaced equally in longitude (their location is TBC). This would ensure that 2 to 3 are able to see the Moon‘s vicinity at any time minimising the probability of cloud obstruction

• A Real time space link extension is needed (minimum 1 Mbit/s up to X-Band station, 4 Mbit/s down to K/optical station per robotic asset).

6.4 Operations

6.4.1 Preparation

• The Operations GS preparation and operation execution will be performed by ESOC for the platform and payload (non science part) and by the SOCC for science.

• GS and Operation preparation in phase C/D will consist of the following items:

1. Procedures writing

2. Simulator design and testing

3. SVT tests

4. Flight Dynamics system design and interface testing

5. Mission analysis

6. Station readiness tests

7. Ground network End to End tests

8. FCT build up and training in simulation program

9. Control room set up

10. Mission planning

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6.4.2 Execution

The targeted launch date is in 2017, the operations execution phase therefore starts in 2017. The operations execution phase has been split up into 4 phases as described below:

1. Phase 1 (2017 for 6 months)

a. Launch

b. Transfer to phase 1 lunar orbit

2. Phase 2 (for TBC years)

a. Payload commissioning

b. Mapping phase


a. Transfer to Phase 2 orbit

b. Relay phase before HLR


a. Relay phase after HLR

6.5 Remarks

In order to minimise the operations cost it is proposed to select a 12 hour orbit with the aposelenium at 12:00 UTC and 24:00 UTC. This will avoid shift work for routine operations before HLR. It has to be noted that an increase in the number of payloads and payload diversity included in the LUMETTO mission will increase the preparation and execution costs of the mission. The delta between “telecom only” payload suite and full payload suite will be 4-5 engineers for preparation until commissioning and 3 for routine operations up to the end of mapping phase.

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7 SYSTEMS

7.1 System Requirements and Assumptions

The overall mission objectives and requirements are listed in section 3.1 A system margin of 20% is applied to all S/C dry masses within the LUMETTO study.

7.2 System and Mission Options and Trades

7.2.1 Trade-Space

The following top-level system and mission trades were considered:

• Launcher: dedicated Soyuz 2-1b fregat, Shared Ariane 5 ECA to GTO minus 6 tons, Shared Ariane 5 ECA with full instrument suite

• Propulsion: chemical, electrical, hybrid (chemical + electrical)

• Transfer: direct, S-E WSB (hybrid), “Smart-1”-like

• Soyuz intial orbit: GTO, HEO 200,000 km

• Phase 1 orbit: 50 x 50 km polar, 45 x 195 km frozen polar orbit

• Phase 2 orbit: L2-based, 3500 x 3500 km 60 deg, 700 x 8100 km 51 deg, 300 x 5600 km 51.7 deg, 3500 x 3500 km 39.23 deg

• Platform: existing, dedicated.

7.2.2 Options Identification & Trade-Tree

The options identified in the previous section are presented graphically in the LUMETTO trade-tree Figure 7-1. The trade-tree shows which options were eliminated (red), which options remained open (gray) and which options were selected (green). The rationale behind the table is discussed in section 7.5.

7.3 Trade-space Qualitative Analysis

In order to minimize the amount of trade options all LUMETTO CDF team-members have given their expert opinion on each of the main trades (listed in section 7.2.1). The advantages and disadvantages resulting from this qualitative analysis are summarized in Table 7-1: .

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Trade 1 Trade 2 Trade 4 Trade 5Launch Vehicle Propulsion Type Staging PlatformSoyuz Ariane Chemical Electric Hybrid Single Stage Dual Stage Existing Dedicated

Systems

AOCS

EP and slow extraction from van allen belts is not looked as a major risk, since it has already been dealt with on SMART-1 or SPIRALE with ''cheap'' sensors. State of the art Star Trackers for example are not anymore bothered by the belts. Gyro MEMS is also very resistant to TID.

SGEO P/F is already dealing with Hybrid Main propustion (400N engine and EP - but not the EP thruster we plan to select)

More tuning configurations to deal with, and additional failure cases.

Modifications needed to reuse SGEO P/F : - Minor : positionning of the STRs, which are placed on the P/F to avoid seeing earth & sun on a GEO orbit. - Minor : The gyros used could be changed (without major redesign to save around 7 kg and also to save power) - Major : The wheels can be downscaled - Major : the EP is not the one we need (SPT100 instead of 5000) and its positionning on the S/C (also the xenon tanks) is not effective (accomodated for NS Station keeping in GEO). - Major : a mechanism similar to Smart 1 for the EP has to be implemented (2-axis control of EP thrusting vector) to avoid too many wheel offloading (using chemical propulsion) - In any case, AOCS SW has to be modified.

Pro :'- Redesign to reduce mass- W.R.T the big effort to exchange main propulsion needed, pointing modes will be different than existing p/f due to the mission around the moon, in case a complete p/f redesign is needed it will not necessarily be much more expensive or risky.

Electric Propulsion

A5ME capability (restartable upper stage) was not considered within the study.A5ME would allow also HEO in orbit insertion as in the soyuz case: EP start after van allen belt escape with reduced transfer time to lunar orbit and reduced cumulated radiation dose.

With reference to consider Lumetto as potential candidate demonstrator of technologies for exploration (future cargo missions) I would recommend to consider two PPS5000 firing in parallel (up to 10kW required power). This choice will allow to short the transfer duration and it will allow demonstartion of a technology capable to enable future large cargo missions.


Chemical Propulsion

400N with ISP in excess of 325s is considered applicable to future telecom scenario (ref is Tomorrow's bird)


Power

10kW power available to EP system is already baseline for BepiColombo. 12 to 18kW is the current payload power available on alphabus platform. Finally during the Space Propulsion Conference 2010 NASA presented their next step for exploration (a 30kW platform with EP technology demo mission - transfer)


Structures / Configuration Small GEO Platform a viable solution

Cost

Lower Cost Dedicated Launch (Higher cost out of CGS) (Arianespace alignment of costs with A5)

Variable cost (likely scenario higher cost) New dev (more expensive) existing (cheaper)

EP only more expensive at component but requires more detail for full system. Synergies with Power and AOCS

2 paralell systems most expensive, but it is a technology development

Single Stage, no interfaces required (lower Cost)

Only makes sense if many recurring units will be used over time.

Cheaper (depends on size and the delta changes required) More engineering but could be smaller and more versatile.

Risk

Highly reliable launch vehicle with a successful launch record. Single launch advantage with respect to shared launch scenario (no interference with second payload) (++)

Interference with second payload of a shared Ariane 5 launch. Schedule delays in delivery, launcher integration and flight readiness. Need to find compatible secondary payload, schedule constraints. (+)

Reliable system with heritage. Shorter lunar transfer time is a plus given the reduced exposure to high radiation areas. (++)

Reliable system with heritage assuming fully redundant thruster assembly and control electronics (PPU). Would require long lunar transfer time with negative implications such as higher radiation exposure during van Allen belt crossings with consequential increase in electronics degradation/malfunction (SEU, SEL, etc.) , solar array degradation, and reduction in overall mission lifetime. Soyuz inserted initial HEO with apogee 200,000km would mitigate radiation impact (-)

Higher complexity as compared to fully electric or fully chemical propulsion options. Lower radiation effects risk as compared to fully electric option. (+)

Single stage offers less complexity and higher reliability (+)

Dual stage is more complex (incl. addition of separation mechanisms) (-)

Existing platform reduces development risk (cost/ schedule implications)

Dedicated platform increases development risk (cost/ schedule implications)

Programmatics

Pro:1) lower mechanical test levels at system level (minor).2) single launch : more flexibility wrt launch date and launch facility for combined operations with launcher allocation to the single ProgramSame:1) launch site in French Guyana (same logistics,shared PPF and HPF facilities with Ariane) and Baikonur (tbc).

Same:1) launch site in French Guyana same logisticsCON:1) Higher mechanical system test levels (minor)

CHEMICALPRO : 1) Chemical propulsion has high heritage in verification,

SAME:1) Minor effects on the schedule for system testing and loading/unloading operations.2) During the mechanical testing the tanks shall be filled with simulant and then drained and dried (minor)

ELECTRICALCON : 1) Electrical propulsion S/S needs development but the TRL is high (used on Smart, Goce, under qualification for BepiColombo).(medium)2) During the mechanical testing the Xe tank shall be filled with Xenon and kept till the launch (minor)3) Requires high voltage harness and some safety constraints during operations and testing (minor)SAME:1) Minor effects on the schedule for system testing and loading/unloading operations.

HYBRIDCON : 1) Electrical propulsion S/S needs development but the TRL is high (used on Smart, Goce, under qualification for BepiColombo).(medium)2) During the mechanical testing the Xe tank shall be filled with Xenon and kept till the launch (minor)3) Requires high voltage harness and some safety constraints during operations and testing (minor)SAME:1) Minor effects on the schedule for system testing and loading/unloading operations.

SINGLE STAGEPRO:1) Easier from AIT/AIV point of view (less mechanical and electrical interfaces to be tested, no additional pyrotecnics or mechanisms)

DUAL STAGECON : 1) Integration and testing of additional hardware (separation system, pyrothecnics, umbilicals).2) Separation test at sytem level to be added with shock mesurement (according to the type of explosive actuators).Some impact on schedule and some additional MGSE to support the separation system (impacts but not big)

Anyway note that it will be not a new approach, see BepiColombo (multistages)

PRO:1) Use of an existing Platform will allow to gain the heritage reducing the Models representativity needs in the Model Phylosophy.2) Some schedule and cost saving CON:1) Effort to make design changes opf the existing platform to meet the LUMETTO Mission and Payload Requirements which will actually reduce the benefits of the heritage.

NOTE : THE USE OF AN EXISTING STRUCTURE, VERY LIKELY WILL REQUIRE SOME MODIFICATION IN THE DESIGN TO ALLOCATE THE NEW PAYLOAD AND TO MEET THE DESIGN LOAD FACTORS. IT IS UNLIKELY THAT A PURE PROTOFLIGHT APPROACH WILL BE FEASIBLE WITH DELTA QUALIFICATION DURING THE SYSTEM TESTING. IT CANNOT BE EXCLUDED AT THE PRESENT TIME, BUT MORE LIKELY AT LEAST A SYSTEM STRUCTURAL MODEL SM WILL BE REQUIRED IN ORDER TO PERFORM THE STATIC LOAD TEST ON THE MODIFIED STRUCTURE AND MECHANICAL ENVIRONMENTAL TESTING AT QUALIFICATION LEVEL.

PROPOSED MODEL PHILOSOPHY:

PRO:1) The design and configuration will better fit the specific needs for the mission and payload requiremnts (LoS, stability, accessibility, loads)2) The qualification at system level to be performed through a Model Philosophy including a Structurat Therma Model (STM), Electrical Test Bench (ETB), and Flight Model. This is the less risky approach in term of verification. During the STM will be verified in advance also the MGSE interfaces, test facilities operations, mechanical procedures and personnel training.3) The structural qualification can run on STM with a later need for the Flight units to be integrated on the FM structure. 4) The STM approach would allow to validate the thermal model and the verify the thermal subsystem in advance wrt the Flight Model. CON:1)Wrt existing platform : Design of a new Structure , harness and Propulsion piping routing. Anyhow also an existing Platform could very likely require big effort in re-design phase for what concern new interfaces and load calculation. 2) Cost and potential schedule impact

PROPOSED MODEL PHILOSOPHY Table 7-1: Trade-Space Qualitative Analysis Table (See Appendix A)

7.4 Options Payload Mass Capacity

7.4.1 Selected options

In order to make well-informed decisions regarding the identified options the assessment of the performance in terms of maximum payload mass capacity is essential. Nine options are selected for this analysis in order to have a good overview of the performance of the different mission options. The selected options are shown in Table 7-2, making use of the numbering system defined in the trade-tree Figure 7-1.

Option Launcher

Primary Passenger mass [kg] Propulsion Staging Phase 1 orbit Phase 2 orbit

1.1.2.1 Dedicated Soyuz into GTO n.a Hybrid Single 50x50km polar 700x8100km, 51deg 1.1.2.2 Dedicated Soyuz into GTO n.a Hybrid Dual 50x50km polar 700x8100km, 51deg 1.2.1.1 Dedicated Soyuz into HEO n.a Chemical Single 50x200 frozen polar 700x8100km, 51deg 1.3.2.1 Dedicated Soyuz into HEO n.a Electrical Single 50x50km polar 700x8100km, 51deg 2.1.1.1 Shared A5 into GTO 6000 Chemical Single 50x200 frozen polar 700x8100km, 51deg 2.1.1.2 Shared A5 into GTO 6000 Chemical Dual 50x200 frozen polar 700x8100km, 51deg 2.2.2.1 Shared A5 into GTO 6000 Hybrid Single 50x50km polar 700x8100km, 51deg 2.2.2.2 Shared A5 into GTO 6000 Hybrid Dual 50x50km polar 700x8100km, 51deg 3.2.2.1 Shared A5 into GTO Not fixed Hybrid Single 50x50km polar 700x8100km, 51deg

Table 7-2: Mission trade options selected for maximum payload mass evaluation

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First the approach used in setting up the mass budgets for all cases is described after which the wet mass, dry mass and maximum payload mass are summarized in section 7.4.3.

7.4.2 Mass Budgets Determination

Due to the fast-track approach of the LUMETTO CDF study not all subsystems were able to produce sensible bottom-up mass budgets. Instead, rough S/S mass estimates or masses based on statistical data were used for most S/S’s.

7.4.2.1 Bottom-up S/S mass estimates

The mass of the DHS, TT&C, AOCS, propulsion and to some extent the power S/S are based on a bottom-up approach. For a detailed overview see the relevant subsystem sections.

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Figure 7-1: LUMETTO Trade-Tree (See Appendix A)

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7.4.2.2 Statistical S/S mass estimates

The thermal control S/S mass is based on statistical data, see Figure 7-2. The statistical data implies that the thermal control S/S mass is always close to 3.9% of the total S/C dry mass. As this assumption contains a lot of uncertainty a 20 % S/S margin is applied.

Total S/C dry mass Vs. Thermal S/S mass

y = 0.039xR2 = 0.975

0

50

100

150

200

250

0 1000 2000 3000 4000 5000 6000

Total S/C dry [kg]

Ther

mal

S/S

mas

s

CDF dataLinear fit

Figure 7-2: Thermal S/S mass vs. total S/C dry mass

The mass for the S/C harnessing is estimated to be 6.81%, again based on statistical data. As this assumption contains a lot of uncertainty a 20 % S/S margin is applied.

Some of the mission options consider a dual stage S/C. The first stage is a kick-stage which will provide the first delta-v burn to reach the S-E WSB after which it is discarded. The mass of this kick-stage is estimated using an inert mass fraction finert of 0.20 and the following relation: )1/(, inertinertpropdrystage ffMM −⋅= .

This results in a dry stage mass of 159.2 kg for option 1.1.2.2 and a dry mass of 172.1 kg for options 2.1.1.2 and 2.2.2.2. These masses are in line with the dry mass of the LPF PM which has a slightly higher dry mass of 191 kg due to the fact that it can carry 1277 kg of propellant using 4 tanks with a mass of approximately 14 kg each. The LUMETTO stage tanks only need to carry propellant masses in the range of 636.9 kg to 688.6 kg, meaning the used mass estimation relation is close to the LPF PM dry mass excluding two tanks.

7.4.3 Options Mass Summary

The above described masses and mass determination methods were used to determine the mass budget for each of the nine selected options.

The total wet mass, total dry mass as well as the final payload mass and payload mass fraction are summarized in Table 7-3 for all of the nine options. Options 2.1.1.1 and 2.1.1.2 are shown in grey as these two options are unable to support the lightest instrument suite (which has a mass of 199.5 kg).

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Option Launcher Propulsion Staging Phase 1 orbitS/C Wet mass [kg]

S/C Dry mass (incl margins) [kg]

Payload mass (incl 20% margin) [kg]

Dry mass to payload mass ratio [%]

1.1.2.1 Dedicated Soyuz into GTO Hybrid Single 50x50km polar 2960.00 1691.29 326.20 19.291.1.2.2 Dedicated Soyuz into GTO Hybrid Dual 50x50km polar 2316.00 1110.40 321.77 28.981.2.1.1 Dedicated Soyuz into HEO Chemical Single 50x200 frozen polar 2316.00 1110.40 215.35 19.391.3.2.1 Dedicated Soyuz into HEO Electrical Single 50x50km polar 2316.00 1798.45 463.23 25.762.1.1.1 Shared A5 into GTO Chemical Single 50x200 frozen polar 3200.00 1156.50 182.31 15.762.1.1.2 Shared A5 into GTO Chemical Dual 50x200 frozen polar 3200.00 1060.68 168.17 15.862.2.2.1 Shared A5 into GTO Hybrid Single 50x50km polar 3200.00 1828.53 395.63 21.642.2.2.2 Shared A5 into GTO Hybrid Dual 50x50km polar 3200.00 1973.43 468.17 23.723.2.2.1 Shared A5 into GTO Hybrid Single 50x50km polar 3533.07 2018.74 485.28 24.04

Table 7-3: Mission options S/C wet mass, S/C dry mass (including all margins), payload mass (including margins) and the payload mass fraction

7.5 Trade-tree Rationale

Based on the qualitative analysis (section 7.3) and the payload mass capacities as summarized in subsection 7.4.3, several options are eliminated and three options are selected as baseline as shown in the trade-tree Figure 7-1. The rationale behind the selection or elimination is summarized in Table 7-4.

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Option Description Status Reasoning1 Soyuz Open

1.1 Hybrid Propulsion Open1.1.1 Frozen polar orbit ~50x200km Open

1.1.2 50x50km polar orbit SelectedBetter payload performance (higher resolution mapping), the 50x50 km polar orbit costs more delta V but for options with an electrical propulsion system this consequence is acceptable

1.1.2.1 Single stage Selected Selected as baseline as the available payload mass is promising, "best" smallest baseline option1.1.2.2 Dual stage Open

1.1.2.1.1 Existing platform Open1.1.2.1.2 Dedicated platform Open1.1.2.2.1 Existing platform Open1.1.2.2.2 Dedicated platform Open

1.2 Chemical Propulsion Open1.2.1 Frozen polar orbit ~50x200km Open

1.2.2 50x50km polar orbit Eliminated50x50 km polar orbit costs too much maintenance delta-V and the selected 50x200km frozen orbit has sufficient delta V

1.2.1.1 Single stage Open

1.2.1.2 Dual stage Eliminated

Dual stage chemical option was considered for the shared A5, where the payload mass decreased compared to single stage plus all other disadvantages of the dual stage concept (see rational table: ...\LUMETTO_Study\LUMETTOMiscellaneous\01 Trades\Synthesis.xls)

1.2.1.1.1 Existing platform Open1.2.1.1.2 Dedicated platform Open

1.3 Electrical Propulsion Open1.3.1 Frozen polar orbit ~50x200km Open

1.3.2 50x50km polar orbit SelectedBetter payload performance (higher resolution mapping), tge 50x50 km polar orbit costs more delta V but for options with an electrical propulsion system this consequence is acceptable

1.3.2.1 Single stage Selected See dual stage comment

1.3.2.2 Dual stage EliminatedDual stage is not a good option for electrical propulsion as the propellant mass is very low and mass savings are unlikely (and complexity and cost increase significantly)

1.3.2.1.1 Existing platform Open1.3.2.1.2 Dedicated platform Open

2 Shared Arianae 5 GTO - 6000kg Open

2.1 Chemical propulsion EliminatedEliminated because both a dual stage and a single stage option can not carry enough payload to support the smallest instrument suite

2.1.1 Frozen polar orbit ~50x200km Open

2.1.2 50x50km polar orbit Eliminated50x50 km polar orbit costs too much maintenance delta-V and the selected 50x200km frozen orbit has sufficient delta V

2.1.1.1 Single stage Eliminated Max payload mass of about 170 kg, which is not enough for smallest instrument suite2.1.1.2 Dual stage Eliminated Max payload mass of about 155 kg, which is not enough for smallest instrument suite

2.2 Hybrid Propulsion Open2.2.1 Frozen polar orbit ~50x200km Open


2.2.2.1 Single stage Open2.2.2.2 Dual stage Open

2.2.2.1.1 Existing platform Open2.2.2.1.2 Dedicated platform Open2.2.2.2.1 Existing platform Open2.2.2.2.2 Dedicated platform Open

2.3 Electrical propulsion EliminatedTransfer time from GTO to the Moon are in the order of 2 years which is considered too long, the lifetime of the operational phase will decrease which is undesireble

3 Shared A5 into GTO - Full instrument suite Open

3.1 Chemical Propulsion OpenLUMETTO Launch mass is likely to be much higher than for the hybrid option. No analysis has been done during CDF study

3.2 Hybrid Propulsion SelectedMost promising option in terms of the combination of reaching the mission objectives compared to chemical (as it does not demonstrate SEP)

3.3 Electrical Propulsion EliminatedTransfer time from GTO to the Moon are in the order of 2 years which is considered too long, the lifetime of the operational phase will decrease which is undesireble

3.2.1 Frozen polar orbit ~50x200km Open


3.2.2.1 Single Stage SelectedPreferred baseline option as dual stage has some major disadvantages (see rational table: ...\LUMETTO_Study\LUMETTOMiscellaneous\01 Trades\Synthesis.xls)

3.2.2.2 Dual Stage Open3.2.2.2.1 Existing platform Open3.2.2.2.2 Dedicated platform Open

Phase 2 orbits

O1 L2 based Eliminated

L2 based communication solutions have been eliminated as the coverage to the poles as poor and the distance to the lunar surface is too large, demanding a large amount of power for communcation for both the oribiter as well as the foreseen surface assets

O2 3500x3500km, 60 degrees EliminatedThe 3500x3500km 60degree orbit is unstable, requiring maintenance as it is not a frozen orbit, thus making it a less attractive option compared to the other phase 2 orbits

O3 300x5600km, 51.7 degress OpenGood option but not the baseline as the O4 option is seems better in terms of coverage and availability and delta-v for orbit transfer from phase 1 to phase 2

O4 700x8100km, 51 degrees SelectedThis phase 2 orbit is selected as baseline for LUMETTO as the coverage and pass duration are better compared to O3, also the delta-v is slightly lower for transfer from phase 1 to phase 2

O5 Eccentric Polar Orbit Eliminated The apocenter of the orbit does not remain stable enough over the south-pole, too much maintenanceO6 Other orbit Eliminated No other candidate communication orbits were identified in the coarse of the CDF study

Table 7-4: Trade-Tree Elimination/Selection rationale

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7.6 Selected Mission Options

7.6.1 Mass Budgets

The three selected options as baseline for the LUMETTO mission are:

• option 1.1.2.1: Soyuz, hybrid, single stage

• option 1.3.2.1: Soyuz, electrical, single stage

• option 3.2.2.1: Shared (primary passenger ~5,667 kg) Ariane 5, hybrid, single stage.

The mass budgets of these options are shown in Table 7-5 through Table 7-7. Element 1 1.1.2.1 - Soyuz Hybrid

Target Spacecraft Mass at Launch 3070.00 kgABOVE MASS TARGET BY: 0.00 kg

Input Input Without Margin Margin Total % of TotalMass Margin Dry mass contributions % kg kg

DI 234.90 20.00 Structure 234.90 kg 20.00 46.98 281.88 20.00DI 45.81 20.00 Thermal Control 45.81 kg 20.00 9.16 54.97 3.90EL Communications 7.40 kg 5.00 0.37 7.77 0.55SST Data Handling 20.00 kg 10.00 2.00 22.00 1.56SST AOCS 43.00 kg 10.00 4.30 47.30 3.36SST Propulsion 220.40 kg 4.10 9.04 229.44 16.28EL Power 245.00 kg 30.00 73.50 318.50 22.60DI 101.12 20.00 Harness 101.12 kg 20.00 20.22 121.35 8.61DI 271.83 20.00 Instruments 271.83 kg 20.00 54.37 326.20 23.14

Total Dry(excl.adapter) 1189.46 1409.40 kgSystem margin (excl.adapter) 20.00 % 281.88 kgTotal Dry with margin (excl.adapter) 1691.29 kg

Other contributionsEL SEPM 0.00 kg - - - -

Wet mass contributionsEL Propellant 1268.72 kg N.A. N.A. 1268.72 42.86

Adapter mass (including sep. mech.), kg 110.00 kg 0.00 0.00 110.00 0.04Total wet mass (excl.adapter) 2960.00 kgLaunch mass (including adapter) 3070.00 kg

Table 7-5: Mass budget of option 1.1.2.1: Soyuz, hybrid, single stage

Element 3 1.3.2.1 - Soyuz Electrical

Target Spacecraft Mass at Launch 2426.00 kgBelow Mass Target by: 0.00 kg







Table 7-6: Mass budget of option 1.3.2.1: Soyuz, electrical, single stage

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Element 6 3.2.2.1 - Shared Ariane 5 - Full payload

Target Spacecraft Mass at Launch 3533.07 kgABOVE MASS TARGET BY: 0.00 kg







Table 7-7: Mass budget of option 3.2.2.1: Shared Ariane 5, hybrid, single stage

7.6.2 Instrument Suites

Based on the available payload masses instrument suites are selected to fit the mission option. The instrument suite masses, available masses and selected instrument suites are shown in Table 7-8.

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Comms onlyComms + NavMappingEnvironmentGravimetry (+ subsats)Seismology

Comms Only

Comms & Navigation

Comms & Navigation & Optical Mapping

Comms & Navigation & Full Mapping

Comms & Navigation & Optical Mapping & Environment

Comms & Navigation & Full Mapping & Environment

Comms & Navigation & Seismology

Option LauncherAvailable Mass (kg)

Option A Option B Option C Option D Option E Option F Option G

1.1.2.1 Dedicated Soyuz into GTO 326.20 199.50 220.50 231.00 268.20 231.00 287.10 417.751.1.2.2 Dedicated Soyuz into GTO 321.77 199.50 220.50 231.00 268.20 231.00 287.10 417.751.2.1.1 Dedicated Soyuz into HEO 215.35 199.50 220.50 231.00 268.20 231.00 287.10 417.751.3.2.1 Dedicated Soyuz into HEO 463.23 199.50 220.50 231.00 268.20 231.00 287.10 417.752.1.1.1 Shared A5 into GTO 182.31 199.50 220.50 231.00 268.20 231.00 287.10 417.752.1.1.2 Shared A5 into GTO 168.17 199.50 220.50 231.00 268.20 231.00 287.10 417.752.2.2.1 Shared A5 into GTO 395.63 199.50 220.50 231.00 268.20 231.00 287.10 417.752.2.2.2 Shared A5 into GTO 468.17 199.50 220.50 231.00 268.20 231.00 287.10 417.753.2.2.1 Shared A5 into GTO 485.28 199.50 220.50 231.00 268.20 231.00 287.10 417.75

Comms onlyComms + NavMappingEnvironmentGravimetry (+ subsats)Seismology

Comms & Navigation & Seismology & Environment

Comms & Navigation & Gravimetry

Comms & Navigation & Gravimetry & Environment

Comms & Navigation & Gravimetry & Seismology


Comms & Navigation & Gravimetry & Full Mapping (HRSC) & Environment

Comms & Navigation & Gravimetry & Full Mapping (HRSC) & Environment & Seismology

Option LauncherAvailable Mass (kg)

Option H Option I Option J Option K Option L Option M Option N

1.1.2.1 Dedicated Soyuz into GTO 326.20 436.65 269.43 288.33 466.68 317.13 336.03 485.281.1.2.2 Dedicated Soyuz into GTO 321.77 436.65 269.43 288.33 466.68 317.13 336.03 485.281.2.1.1 Dedicated Soyuz into HEO 215.35 436.65 269.43 288.33 466.68 317.13 336.03 485.281.3.2.1 Dedicated Soyuz into HEO 463.23 436.65 269.43 288.33 466.68 317.13 336.03 485.282.1.1.1 Shared A5 into GTO 182.31 436.65 269.43 288.33 466.68 317.13 336.03 485.282.1.1.2 Shared A5 into GTO 168.17 436.65 269.43 288.33 466.68 317.13 336.03 485.282.2.2.1 Shared A5 into GTO 395.63 436.65 269.43 288.33 466.68 317.13 336.03 485.282.2.2.2 Shared A5 into GTO 468.17 436.65 269.43 288.33 466.68 317.13 336.03 485.283.2.2.1 Shared A5 into GTO 485.28 436.65 269.43 288.33 466.68 317.13 336.03 485.28

Table 7-8: Summary of the instrument suites with corresponding masses and evaluated mission options with respect to the instrument suites

7.6.3 Mission Objectives Compliance Matrix

The mission objectives compliance matrix for the three baseline options with the selected payload suites as indicated in Table 7-8 is shown in Table 7-9. All three mission options fully fulfil all mission objectives but one. The only objective that has not yet been fulfilled is “to provide opportunities for the demonstration of advanced generic technologies in areas such as power management”. Up till now only experiments were identified for the AOCS S/S, subsection 8.4.3. It is likely other S/S demonstration experiments can be identified at a later design phase.

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Option 1.1.2.1 1.3.2.1 3.2.2.1Launcher Soyuz Soyuz Shared A5Propulsion Hybrid Electrical HybridStaging Single Single Single

Instrument package


Comms & Navigation & Gravimetry & Full Mapping (HRSC) & Environment

Comms & Navigation & Gravimetry & Environment & Seismology & Full Mapping (HRSC)

Maximum Allowable Payload mass [kg] 326.20 463.23 485.28Mass of selected Payload Suite [kg] 317.13 336.03 485.28Mission Objectives Compliance Matrix1. Demonstrate advanced capabilities for future exploration Partially Fulfilled Partially Fulfilled Partially Fulfilled1.1 Demonstrate advanced telecommunication technologies (optical) and delay tolerant telecommuncations protocols Fulfilled Fulfilled Fulfilled

1.2 Demonstration of tele-operations in combination with lunar surface payloads Fulfilled Fulfilled Fulfilled

1.3 Demonstrate advanced electric propulsion for orbit transfer and orbit and attitude control Fulfilled Fulfilled Fulfilled

1.4 Demonstrate advanced mission trajectories and associated operations Fulfilled Fulfilled Fulfilled

1.5 Provide opportunities for demonstration of advanced generic technologies in areas such as power management Partially Fulfilled Partially Fulfilled Partially Fulfilled

2.Prepare and support robotic and human lunar surface operations in two phases Fulfilled Fulfilled Fulfilled

2.1 Phase 1: Perform lunar mapping in preparation and support of robotic and huma lunar surface exploration Fulfilled Fulfilled Fulfilled

2.1.1 Provide opportunities for accomodation of instruments of failed national orbiter missions (DLR's LEO, BNSC's Moonlite, ASI's Maggia)

Fulfilled Fulfilled Fulfilled

2.1.1.1 Global coverage with a resolution < 1 m (stereo) and spectral resolution < 10 m (range 0.2 - 14 microns) Fulfilled Fulfilled Fulfilled

2.1.1.2 Subsurface sounding (global coverage) mapping (few metres deep with mm resolution; hundred metres deep with m resolution)


2.1.1.3 Lunar space environment and dust characterisation Fulfilled Fulfilled Fulfilled2.1.2 Support mission planning and operations of early lunar surface robots Fulfilled Fulfilled Fulfilled

2.2 Phase 2: Provide communcation and navigation services for robotic and early human exploration missions Fulfilled Fulfilled Fulfilled

2.2.1 Provide lunar data relay service Fulfilled Fulfilled Fulfilled2.2.2 Provide time synchronisation services Fulfilled Fulfilled Fulfilled2.2.3 Provide initial lunar surface positioning service Fulfilled Fulfilled Fulfilled

3. Optimize utilisation of platform and mission for opportunistic scientific investigations Fulfilled Fulfilled Fulfilled

3.1 Provide opportunities for accommodation of scientific instruments of failed national orbiter missions (DLR’s LEO, BNSC’s Moonlite, ASI’s Maggia) e.g. for measurement of gravity and magnetic field, heat-flow, lunar seismology, lunar crater altimetry, nature and origin of lunar polar volatiles


Table 7-9: Mission Objectives Compliance Matrix

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8 LUMETTO SPACECRAFT

8.1 Electrical Propulsion

8.1.1 Assumptions and Requirements

The electrical propulsion system is to provide the following delta-v’s:

• 4520 m/s for cases 1.1.2.1 and 3.2.2.1

• 3528 m/s for case 1.3.2.1

As one of the mission objectives is to demonstrate an advanced electric propulsion system, a non-flight-proven engine is preferred. The engine should furthermore be ready for a 2017 launch.

8.1.2 Options

During the LUMETTO study four candidate electric propulsion engines were identified. The nominal characteristics of these engines are summarized in Table 8-1.

Engine Power [W] Thrust [mN] Specific Impulse [s] TRL PPS1350 (SNECMA, Smart-1) 1,500 87 1600 9T6 (QINETIQ, Bepi-Colombo) 5,000 150 4500 6PPS 5000 (SNECMA)* 5,000** 276 1763 4-510 kW HET 10,000 440 3125 3

*Nominal operation at discharge voltage of 300V. Higher specific impulse can be achieved with higher discharge voltage for a reduced thrust level.

**Nominal power level. The thruster can be operated up to 6000W with higher thrust and specific impulse.

Table 8-1 Electrical engine options for LUMETTO

8.1.3 Baseline

The electrical propulsion considered for the LUMETTO S/C is the PPS5000 as:

• PPS1350 has flown on Smart-1 (so no technology demonstration is needed) and the thrust level is too low, resulting in long transfers duration and high degradation levels (reducing the operational life-time of the S/C)

• The T6 engine will fly on Bepi-Colombo (so no technology demonstration) plus the available thrust is considered too low

• the 10 kW engine demands too much power considering the overall LUMETTO mission.

The PPS5000 engine can be operated at a range of power settings resulting in a range of specific impulses versus thrust, shown in Figure 8-2 and Figure 8-3.

In the LUMETTO study the PPS5000 is set to operate at 5 kW as a baseline.

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The SEP architecture is shown schematically in Figure 8-1. Redundancy is applied at component level and equipment level (1 thruster unit in cold redundancy). The mass breakdown of the SEP system for option 1.1.2.1 is shown in Table 8-2.

The propellant masses are calculated based on the delta-v budgets provided by mission analysis (Table 5-3), an Isp of 1763 s and Tsiolkovsky’s relation. The initial masses used are 2323.1 kg (option 1.1.2.1), 2316 kg (option 1.3.2.1) and 2772.8 kg (option 3.2.2.1), resulting in propellant masses of 534.3 kg, 491.8 kg and 637.7 kg, respectively.

Figure 8-1: SEP architecture schematic (left) and PPS5000 (right)

Figure 8-2: Specific impulse vs. Power for the PPS5000 at 300 V and at 450 V

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Figure 8-3: Thrust vs. Power for the PPS5000 at 300 V and 450 V

The PPS5000 system preliminary dry mass is reported in the following table.

1.1.2.1

unit mass

number of units

total mass margin

mass with margin

Propellant Module Xenon tank 53.42578 1 53.42578 20 64.11093479Xenon filter 0.12 1 0.12 5 0.126FDV 0.05 1 0.05 5 0.0525HPT/LPT 0.12 2 0.24 20 0.288E-HPR 2.8 1 2.8 10 3.08Precard 1.7 1 1.7 10 1.87Electrical Module PPU(Plasma) 12 1 12 20 14.4Electrical Filter Unit 1.3 1 1.3 10 1.43Hot interconector box 0.6 2 1.2 10 1.32Bus Harness Harness electronic connectors for TU 0.4 2 0.8 10 0.88TSU (internal to PPU) 1.7 1 1.7 10 1.87Thruster Module Thruster (TU) 10.1 2 20.2 20 24.24XFCU 0.4 2 0.8 20 0.96FU 1.2 2 2.4 10 2.64Gimballed mechanism 11.5 1 11.5 10 12.65Thermal Control, MLI 0 Pipework 0 Harness & Connectors 0 Structural Baseplate 0

TOT 110.2358 130

Table 8-2: Mass breakdown for SEP system for option 1.1.2.1

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Finally, if low thrust (in the order of 40mN) is sufficient to perform orbit maintenance, low power (less than 600W) Electrical systems can be considered to complement the PPS5000 system. This is because:

• Small thrusters are easily placed according with thrust vector demand

• A low power EP system could be easily implemented, sharing the propellant storage and management system as well as the power conditioning unit with the already implemented PPS5000 system

• Additional dry mass required by adding thrusters unit is limited.

8.2 Chemical Propulsion


The main chemical propulsion system’s task is to provide the initial delta-v burn of 763 m/s for cases 1.1.2.1 and 3.2.2.1. The design is driven by a high specific impulse to reduce the required propellant mass.

In estimating the total mass of the propellant system a modified Eurostar 2000 tank design is used for the three baseline options. For the other investigated options with larger delta-V requirements the Eurostar 2000 tank proved to be too small and an adapted Eurostar 3000 series tank was used instead. Furthermore the use of a single main engine is assumed.

8.2.2 Baseline

The chemical propulsion system selected for the LUMETTO S/C is the S400-15 bi-propellant engine, with performance shown in Figure 8-4. The engine has a specific impulse of 321 s at a nominal thrust of 425 N.

Figure 8-4: Performance of the S400-12 and S400-15 bi-propellant engines

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The mass of the chemical propulsions system is estimated (shown in Table 8-3 through Table 8-5) by first estimating the total propellant mass required to perform all delta-V manoeuvres. The main engine delta-V is estimated to be 763 m/s for both LUMETTO mission options using a chemical main engine (case 1.1.2.1 and 3.2.2.1). Assuming an additional 2.5% delta-V required for AOCS manoeuvres, the total delta-V adds up to 782 m/s.

With a propellant margin of 5% and an additional propellant residual factor of 1.5% the total propellant mass is estimated to be 708 kg, based on an initial launch mass of 2960 kg (Soyuz into GTO minus 110 kg for the adapter). In the CDF a total propellant margin of 2.5% is applied with an additional 100% margin for the AOCS propellant. The used margins result in approximately similar total propellant masses but are to be changed to CDF standards in a next iteration.

Assuming an oxidiser to fuel ratio of 1.65 the oxidiser volume and mass are estimated in Table 8-3 to be 0.322 m3 and 440.7 kg, respectively, and the fuel volume and mass are estimated to be 0.303 m3 and 267.1 kg, respectively. Assuming four modified Eurostar 2000 tanks are used, the total mass per tank is estimated in Table 8-4 to be 10.8 kg. Using a system architecture as sketched in Figure 8-5, the equipment mass breakdown is shown in Table 8-5. The total chemical propulsion system mass is calculated to be 90.4 kg, excluding margins, for case 1.1.2.1. For case 3.2.2.1 (shared Ariane 5 into GTO) the total propulsion system mass is estimated to be 97.7 kg. A S/S margin of 10 % is applied as most of the equipment only needs to be modified.

Table 8-3: Propellant mass calculation for the S400-15 bi-propellant engine for

cases 1.1.2.1 and 3.2.2.1

Table 8-4: Chemical propulsion system tank mass budget

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Table 8-5: Chemical propulsion system mass budget

Figure 8-5: Chemical propulsion system architecture, showing only two of the four

tanks

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8.3 Power

For the power subsystem two distinct options have been studied based on the main propulsion subsystem selection: chemical propulsion and electric propulsion. This covers options 1.2, 1.3, 2.1, 2.3, 3.1 and 3.3. The other options (hybrid propulsion) are covered within the uncertainty margins.

8.3.1 Chemical Propulsion Option

8.3.1.1 Assumptions and Requirements

• The payload power consumption, regardless of the payload suite selection is assumed to be 1 kW (1020W to 1045W). It is assumed that units associated with mapping and sub satellite tracking will not be active during Phase 2 of the mission (at least at the end of the mission) and these units will not require (or very little) compensation heating for keeping an acceptable S/C thermal balance.

• The nominal platform power requirement will be 400W to 500W, this includes power for thermal control.

• During the mapping phase (phase 1) the eclipse duration is similar to that of the data relay phase (phase 2) but the eclipse ratio is higher. A spacecraft sized for the data relay phase (phase 2) of the mission will provide a 500W power capability during the mapping phase (phase 1) of the mission. A 500W power capability is sufficient for the mapping and tracking payloads required during Phase 1.

• The payload suite can require a 450W power of compensation heating when OFF to keep a safe temperature (50°C below operating temperature). The following assumptions have been made in order to satisfy this requirement:

o The power dissipation of units ON in mapping phase can contribute to keep telecom payload (OFF) at a reasonable temperature.

o The telecom payload thermal dissipation is only a fraction of its consumption, i.e. actual compensation heating will be << 450W.

o The Solar Array will be less degraded for the mapping phase, and some higher IR input could help i.e. various margins exist for accommodating the mapping phase power need within the telecommunication phase sizing.

o The thermal design of the various payloads for their OFF state shall be optimized for the power subsystem. Note: ideal for such a S/C design could be the usage of louvers.

• The mission includes at least 1200 50 minute eclipses with continuous operation.

For the chemical propulsion option the telecommunications phase is the sizing phase.

8.3.1.2 Baseline power architecture

The properties of the baseline power subsystem baseline is listed below:

• Battery bus with Direct Energy Transfer

• The payloads shall have their own pointing mechanisms in order to keep the Solar Array pointing optimal in all phases of the mission.

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• Requirements are similar to an Earth Observation mission therefore heritage from Earth Observation missions shall be taken (e.g. Sentinel).

The proposed equipment is listed in Table 8-6.

Equipment Mass [kg]

PCDU 35

Solar Array (2kW EOL) 2x25

Battery 70

Table 8-6: Chemical propulsion baseline equipment list

8.3.2 Electric Propulsion Option

8.3.2.1 Assumptions and Requirements

• The electrical engine requires 5.5 kW, this includes losses

• The Electrical Thrusters’ Power Supply is connected to the S/C power bus, including Solar Array power conditioning and the two stages of power conversion.

• The thrusters are active only in Sunlight, the impact on the battery sizing is limited.

• Thermal compensation heating for the telecommunication payload is not required when propulsion is ON (a need for the mapping payloads is considered to be included in the margins)

• The higher solar array capability can be used to better optimize the battery SOC management.

• A higher solar array capability can be used to better optimize the battery SOC management.

8.3.2.2 Baseline power architecture

The properties of the baseline power subsystem baseline is listed below:

• 50V regulated bus with Direct Energy Transfer

• The Solar Array pointing can be kept almost optimal in all operational modes, at least during the propulsion ON modes.

• The electrical propulsion option includes potentially more eclipses (luncar transfer). However, it is assumed that the power dissipation of the battery discharge regulator (regulated bus) can provide a part of the needed heat.i.e. this cancels one of the drawback of the architecture with respect to a high number of eclipse cycles and high eclipse ratio.

• The architecture maximizes the heritage from EP in GTO and previous ESA missions. However high power EP on a small platform implies specific adaptations (e.g. rather 100V IF for EP, management of an abrupt thrust interruption).

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The proposed equipment is listed in Table 8-7.

Equipment Mass [kg]

PCDU 45

Solar Array (6kW EOL) 2x65

Battery 70

Table 8-7: Electrical Propuslion baseline equipment list

8.3.3 Remarks

• A 30% margin is considered, critical assumptions such as for thermal control (no significant compensation heating), mission analysis (Solar Array perfect illumination in all phase (e.g. COM, permanent telecom)) and Electrical Propulsion IF for the EP case.

• SA mass is without SADM

• The baseline is a 2 wing Solar Array, but a 1 wing solution would have advantages in mass and fields of view if it is acceptable for attitude control.

8.4 AOCS


The following design drivers were used for the AOCS S/S:

• The platform required pointing accuracy in the order of 0.1 deg

• In case of long transfer through Van Allen belts make use of equipment that can sustain the expected amount of radiation

• If possible make use of Small GEO platform and AOCS S/S.

8.4.2 Baseline Design

Changes for the AOCS S/S with respect to Small GEO include:

• For the Hybrid & EP cases implementation of a different EP Thruster on a gimbaled mechanism is required.

o On SGEO, 4+4 EP thrusters are mounted to deal with NS and EW Station Keeping with their dedicated Xenon Tanks. On LUMETTO, only one tank is needed (central) with a gimbaled mechanism to limit wheel offloading using CP or cold gas.

• 2 REGYS20 Gyros to be changed to MEMS gyros (saving 7 kg)

• CP system is equivalent to SGEO (same equipments and redundancy for both CP main propulsion and Cold gas system)

o In the Hybrid & CP case, 10 N Thrusters alignments are slightly changed to allow better torque capacity around Z axis.

• Major update of AOCS SW and tuning to deal with new equipments and new attitude profiles (cruise to the moon, moon pointing).

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The LUMETTO AOCS S/S then consists of:

• Star Trackers : 2.61 kg each

• Gyros : 0.8 kg each

• Solar Cells assembly : 0.4 kg

• Wheels assembly : 36.6 kg.

The baseline masses for LUMETTO mission options based on a hybrid or EP propulsion system is 43 kg excluding 10 % and 36 kg excluding 10 % margin for the options having only a chemical propulsion system. The power consumption for the AOCS S/S is estimated to be approximately 100 W during nominal operations and 300 W peaks during slew periods with wheel usage at maximal torque.

The pointing performance of 0.1 deg is within reach. (same requirement for Small Geo P/F)

The propellant mass used by the 10 N thrusters is estimated to be 2.5 % of the total propellant mass of the main engine during LAE firing.

Van Allen belts crossing is not considered as an issue for the AOCS equipments.

8.4.3 Proposed AOCS Experiments

For the AOCS S/S two experiments are proposed for the LUMETTO mission:

• Optical navigation to the moon

• GNSS signals at high altitude.

Optical navigation to the moon

• Demonstrate absolute navigation based on landmark matching in lunar orbit o Required in descent orbit within a Moon Lander scenario, from 100 km altitude o Wide FOV required (typ 70°)

• Demonstrate autonomous interplanetary navigation capability based on limb measurements of Moon and Earth

o Required for solar system exploration missions requiring multiple flybys, e.g. Laplace

o Small FOV required (typ 20 °) o On-board autonomous experiment to go further than the a posteriori

experiment already performed with Smart-1 (AutoNav experiment with AMIE camera images)

o On-board autonomy experiment = image processing on-board (requiring more mass and more power)

• Mass : Around 1 kg

• Power : between 1 and 10W (depending on on-board processing needs)

GNSS Signals at high altitude

• Objective : Demonstrate the feasibility of using GPS signal at very high altitude (former baseline of Orion) and confirm the available signal power. Especially for L2C or E1 using pilot channel, allowing low C/N0 carrier tracking.

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• Equipment needed : GNSS receiver (GPS/Galileo) + High gain antenna

• Mass : Around 3 kg

• Power : Less than 20 W.

8.5 Data Handling


A precise requirements definition is not possible for the level of detail present in the proposed design. Therefore, all the traded options will have the same DHS architecture and requirements. The following assumptions have been taken into account for proposing the DHS baseline architecture:

• Instrument data processing will be performed by the spacecraft bus computer (not by the instruments themselves)

8.5.2 Baseline DHS Architecture

The selected architecture involves a service module onboard computer (CDMU) and a payload-dedicated computer (PCMU) for high-power instrument data processing as can be seen in Figure 8-6 and Table 8-8.

8.5.2.1 Service Module (CDMU)

The service module is characterised with the following items:

• Based on the HICDS (Highly Integrated Control and Data System) ESA study

• HICDS is driving the present development path for new technologies in DHS in Europe. HICDS uses standard components in a more compact way. A significant gain w.r.t. current missions is reached: Mass/4 and Power/2

• Study Heritage: used as baseline for other studies such as SOLO2, XEUS, MDL2, Cross-Scale. The HICDS demonstrator (engineering model) has been delivered to ESTEC in 2007 thus is suitable for this mission

• LEON2 FT Performance: 70 MIPS – 10 MFLOPS @100MHz

• FLASH memory is proposed as solution in favour of SDRAM.

8.5.2.2 Payload Module (PCMU)

The payload module is characterised with the following items:

• Responsible for Instrument and mechanisms (shutter, filter wheel, calibration units) control

• Responsible for Housekeeping and interfacing with the Service Module (TC).

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Figure 8-6: DHS Architecture

Table 8-8: DHS Mass budget

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8.6 Structures & Configuration

8.6.1 Structures

The structure S/S of the LUMETTO S/C has not been designed using a bottom-up approach. Instead the total mass of the S/S can be estimated using an approximation relation such as:

• Structures S/S mass is estimated with: 100/)60)(16( drydrystruc MMLOGM ⋅+⋅−= (see Figure 8-7)

• Structures S/S mass is estimated to be 20% of the total S/C dry mass.

The first approximation relation was estimated to be too optimistic for the LUMETTO S/C and the second, more conservative, relation was used instead.

Figure 8-7: Approximation of the structures S/S mass as a function of S/C dry

mass

8.6.2 Configuration

8.6.2.1 Requirements and design drivers

The main design driver for the configuration of the LUMETTO S/C is the placement of:

• Payload: communication, mapping, sub-satellite and 3 penetrators (for case 3.2.2.1)

• Solar array of 25 m2 to provide 6 kW EOL

• Launcher is either a dedicated Soyuz 2-1b Fregat or a shared Ariane 5 ECA with LUMETTO as secondary passenger inside the Sylda module.

8.6.2.2 Assumptions

The Small GEO platform comes close to the requirements of LUMETTO and has been used as reference.

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8.6.2.3 Baseline

During the LUMETTO CDF study the outer configuration of the S/C has been assessed briefly with the goal of verifying the available volume in the fairing for both launchers as well as the overall placement of the payload and solar panels on the S/C.

The S/C outer dimensions are estimated to be approximately 2 x 1.8 x 2.5 m with 25 m2 of solar panels to provide the 6 kW EOL. The outer configuration of option 3.2.2.1, which carries the maximum payload suite, is shown in Figure 8-8.

Figure 8-8: Outer configuration of the LUMETTO S/C for option 3.2.2.1, carrying

all identified instruments

For the shared Ariane 5 option a Sylda +1500 (CUB) module with a diameter of 4000 mm is used to separate the primary passenger from the secondary. The LUMETTO S/C is to be placed inside the Sylda module. The adapter will be one of the following:

• 1194VS, mass = 165 kg, height = 753mm

• 1666MVS, mass = 160 kg, height = 886mm

For the Soyuz 2.1b fregat a dedicated launch is assumed and one of the following adapters is to be used:

• 1194SF, mass = 110Kg, height = 230mm

• 1666SF, mass = 100Kg, height = 457mm

The diameter of the fairing of the Soyuz fregat is 3800 mm and the overall dimensions of the fairing of both launchers is shown in Figure 8-9. In Figure 8-10 it is shown that the LUMETTO S/C fits in both fairings with sufficient margins. The load capability for OTS adapters for a “Small Geo”-type satellite are:

• Soyuz - 1194 SF: 2500 kg at 2300 mm from separation plane

• Ariane 5 - PAS 1194MVS: 7000 kg at 2400 mm from separation plane.

The mass and dimensions of the LUMETO S/C are within these specifications.

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Figure 8-9: Dimensions of the fairings of the Ariane 5 ECA (left) and Soyuz 2-1b

fregat (right) launchers

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Figure 8-10: Fairing of Ariane 5 (left) and Soyuz (right) with conceptual

LUMETTO accommodation

Subsat consideration:

The Subsat dimensions were derived from it’s given mass characteristics of 30 kg (Bus 20 kg payload & propulsion 10 kg). Based on former studies on this size of satellite RD[1], the size for a cubic-shaped spacecraft would be 40 x 40 cm.

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9 RISK

9.1 Risk Management Policy – Objectives

Maximize the probability of achieving LUMETTO’S intended goals and to contribute to the projects’ risk management process.

The CDF risk management policy for LUMETTO aims at handling risks which may cause serious negative cost, schedule, and/or technical value impacts on the project.

Risk Management Process definition:

Figure 9-1: Risk Management Process

An organized, systematic decision making process that efficiently identifies, analyzes, plans, tracks, controls, communicates, and documents risk to increase the likelihood of achieving the project goals.

9.2 Risk Management Policy – Project Goals

Exploration

• Demonstrate advanced capabilities for future exploration (technology demo)

• Prepare and support robotic and human lunar surface operations in two phases o Phase 1: Perform lunar mapping in preparation and support of robotic and

human lunar surface exploration o Phase 2: Provide communication and navigation services for robotic and early

human exploration missions

• Optimise utilisation of platform and mission for opportunistic scientific investigations.

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No/ minimal consequences.

Failure results in a substantial reduction

(<30%) of the mission’s exploration

return

Failure results in an important reduction

(30-70%) of the mission’s

exploration return

Failure results in a major reduction (70-

90%) of the mission’s exploration return

Failure leading to the impossibility of

fulfilling the mission’s technology

development for exploration objectives

Exploration


Significant increase in

estimated cost

Major increase in estimated

cost

Critical increase in estimated

cost

Cost increase result in project

cancellation

Cost


Dependability: Minor degradation of system (e.g.: system is still able to control the

consequences) Safety: Impact less than minor

Dependability: Major degradation of the system.

Safety: Minor injury, minor disability, minor occupational illness. Minor system or

environmental damage.

Dependability: Loss of mission. Safety: Major damage to flight systems, major damage to ground facilities; Major

damage to public or private property; Temporarily disabling but not life- threatening injury, or temporary

occupational illness; Major detrimental environmental effects.

Safety : Loss of system, launcher or launch facilities.

Loss of life, life-threatening or permanently disabling injury or occupational illness;

Severe detrimental environmental effects.

Technical (ECSS-Q-30 and ECSS-Q-40)

No/ minimal consequences

Minimum 1

Launch delayed (TBD)

months

Significant

2


months

Major

3


months

Critical

4

Launch opportunity lost

Catastrophic

5

Schedule Severity

Technical

The spacecraft platform and payloads shall operate correctly during the specified lifetime (e.g. Phase 2 shall support Data Relay for defined assets on the lunar surface for at least 10 years OPE-030, Phase 1 mission time TBD)

Schedule

Launch date shall be no later than 2017 (start 2010)

Cost

The cost at completion shall be ≤ 600 M€

9.3 Precursor Mission Severity Categorization

Table 9-1: Severity/Likelihood Categorization & Risk Index

Will almost never occur, 1 in 10000 projects

Pf=0.0001 R=0.9999 Minimum A (1)

Will occur seldom, about 1 in 1000 projects

Pf=0.001 R= 0.999 Low B (2)

Will occur sometimes, about 1 in 100 projects

Pf=0.01 R=0.99 Medium C (3)

Will occur frequently, about 1 in 10 projects

Pf=0.1 R=0.9 High D (4)

Certain to occur, will occur once or more times per project. Maximum E (5)

Definition Likelihood Score

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9.4 Top Risk Log

The identified risks are valid for all studied mission options. The risk log is organized by mission element and summarized in the Top Risk Index Chart below:

Severity

5 1 1 1

4 4 5

3 2 2 4

2

1

A B C D E

Likelihood

Table 9-2: Top Risk Index Chart

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page 80 of 102

Table 9-3: Top Risk Log (1/2)

Ris

k Ty

peR

isk

inde

x R

isk

scen

ario

Cla

ssifi

catio

n C

ause

Miti

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itiga

ting

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n 2

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gatin

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estim

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n of

gro

und

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ent r

eadi

ness

/inve

stm

ent

requ

ired

to su

ppor

t con

tinuo

us

spac

e op

tical

com

mun

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ions

.

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t/sch

edul

e

Glo

bal o

ptic

al g

roun

d st

atio

n ne

twor

k is

re

quire

d to

supp

ort c

ontin

uous

opt

ical

lin

k. L

imite

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und

stat

ions

un

der E

SA o

r Eur

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ate.

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mita

tions

of s

pace

opt

ical

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n.Te

chni

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ther

eff

ects

on

long

-rang

e op

tical

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stem

s. Su

n po

sitio

n re

lativ

e to

tra

nsm

itter

and

rece

iver

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Net

wor

k of

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ical

gro

und

stat

ions

with

suff

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nt

geog

raph

ical

dis

tanc

e am

ong

them

.

Gro

und

stat

ions

pre

fera

bly

loca

ted

at h

igh

altit

ude

with

stab

le a

nd

clea

n at

mos

pher

ic c

ondi

tions

to

min

imiz

e w

eath

er e

ffec

ts.

Early

star

t of i

nter

natio

nal

coop

erat

ion

agre

emen

ts.

3DLi

mite

d ES

A e

xper

ienc

e in

op

tical

com

mun

icat

ions

(spa

ce).

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nica

l

Euro

pean

spac

e op

tical

com

mun

icat

ions

he

ritag

e lim

ited

to th

e A

RTE

MIS

(GEO

) an

d SM

AR

T-1

expe

rimen

ts. N

o lo

ng te

rm

serv

ice

prov

idin

g ex

perie

nce

at E

arth

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oon

dist

ance

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st in

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nolo

gy a

nd

test

ing.

LUM

ETTO

is c

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an

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ptic

al sp

ace

com

mun

icat

ions

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a re

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sion

.

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ious

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ical

co

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unic

atio

n ex

perim

ents

.

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nche

r

4CIn

terf

eren

ce w

ith se

cond

pay

load

of

a sh

ared

Aria

ne 5

laun

ch. N

eed

to fi

nd c

ompa

tible

seco

ndar

y pa

yloa

d, sc

hedu

le c

onst

rain

ts.

Sche

dule

Sche

dule

del

ays i

n de

liver

y, la

unch

er

inte

grat

ion,

and

flig

ht re

adin

ess.

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rnat

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dedi

cate

d So

yuz

laun

ch.

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to a

rran

ge a

ll le

gal a

spec

ts

rela

ted

to a

shar

ed la

unch

scen

ario

.

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sion

4DC

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ex o

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tiona

l re

quire

men

ts, m

ultip

le m

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on

phas

es/p

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ads,

and

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sion

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ratio

n >1

0 ye

ars.

Mis

sion

Dem

andi

ng m

issi

on o

bjec

tives

incl

. de

mon

stra

tion

of a

dvan

ced

capa

bilit

ies

for e

xplo

ratio

n, lu

nar m

appi

ng, o

ptic

al

and

Ka

Ban

d da

ta re

lay,

nav

igat

ion

and

scie

ntifi

c lu

nar e

xplo

ratio

n.

The

prim

ary

focu

s is t

ech.

de

mo.

It is

und

erst

ood

that

da

ta re

lauy

and

map

ping

se

rvic

es a

re se

cond

ary

obje

ctiv

es

4DD

ata

rela

y m

issi

on o

bjec

tive

high

ly d

epen

dant

on

ESA

ext

erna

l lu

nar s

urfa

ce in

frast

ruct

ure

deve

lopm

ent.

Mis

sion

Rob

otic

/hum

an lu

nar e

xplo

ratio

n pl

ans

are

bein

g le

ad b

y N

ASA

. Unc

erta

inite

s in

futu

re L

unar

exp

lora

tion

stra

tegi

es.

Coo

rdin

ate

mis

sion

ob

ject

ives

with

Lun

ar

surf

ace

data

rela

y re

quire

men

ts.

Act

ive

and

clos

e co

oper

atio

n w

ith

all L

unar

exp

lora

tion

partn

ers.

4C

Dua

l sta

ge te

chni

cal c

ompl

exity

(in

cl. a

dditi

on o

f sep

arat

ion

mec

hani

sms)

. Sm

all

impa

ct o

n sc

hedu

le a

nd so

me

addi

tiona

l M

GSE

requ

ired

to su

ppor

t the

se

para

tion

syst

em.

Tech

nica

l

Add

ition

al fa

ilure

cas

es. M

ore

mec

hani

cal/e

lect

rical

inte

rface

s, ad

ditio

n of

sepa

ratio

n m

echa

nism

s, an

d hi

gher

op

erat

iona

l com

plex

ity. A

dditi

onal

test

ing

and

grou

nd su

ppor

t equ

ipm

ent r

equi

red

to

supp

ort d

ual s

tage

syst

em.

Bas

elin

e a

sing

le st

age

spac

ecra

ft w

hich

has

less

m

echa

nica

l/ele

ctric

al

inte

rfac

es a

nd n

o ad

ditio

nal

pyro

s or m

echa

nism

s.

Sepa

ratio

n te

st a

t sys

tem

leve

l to

be

adde

d w

ith sh

ock

mea

sure

men

t (a

ccor

ding

to th

e ty

pe o

f exp

losi

ve

actu

ator

s).

LUMETTO


page 81 of 102

Table 9-4: Top Risk Log (2/2)

Plat

form

4D

Und

eres

timat

ion

of re

quire

d m

odifi

catio

ns to

exi

stin

g (o

r in

deve

lopm

ent)

tele

com

pla

tform

s (e

.g. S

MA

LL G

EO) t

o ac

com

mod

ate

LUM

ETTO

's pa

yloa

ds a

nd m

eet i

ts sp

ecifi

c m

issi

on re

quire

men

ts. I

mpa

ct o

n co

st a

nd sc

hedu

le.

Cos

t/sch

edul

eM

ajor

stru

ctur

es, p

ropu

lsion

, pow

er,

AO

CS,

com

ms a

nd S

W m

odifi

catio

ns

requ

ired.

Exis

ting

plat

form

redu

ces

tech

nica

l ris

k as

it is

bas

ed

on su

cces

sful

spac

e op

erat

ions

.

Dev

elop

men

t ris

k co

uld

be lo

wer

if

exis

ting

plat

form

is a

n ad

equa

te fi

t.B

asel

ine

a de

dica

ted

plat

form

.

4D

Hig

h ra

diat

ion

expo

sure

dur

ing

van

Alle

n be

lt cr

ossi

ngs w

ith

cons

eque

ntia

l inc

reas

e in

el

ectro

nics

de

grad

atio

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.

LUMETTO


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10 PROGRAMMATICS From the programmatic perspective, the D/TEC approach stipulates, that the technology maturity of the payload to be demonstrated shall be at the appropriate Technology Readiness Level (TRL) before being selected. In addition, a clear development and verification status is required.

The TRLs present a systematic measure, supporting the assessments of the maturity of a technology of interest and enabling a consistent comparison in terms of development status between different technologies. The following definitions do apply:

Table 10-1: Technology Readiness Levels (TRL)

The European Space Technology Master Plan gives the following statement:

“…To prepare the legal basis for multi-annual programmes it takes about 18 month to reach political agreement on financial ceiling…”

In order to achieve a reasonable estimation of the necessary development durations, this additional time period has to be taken into account. The following table presents an indication for the development periods according to the European Space Technology Master Plan:

Actual system “flight proven” through successful mission operations TRL 9

Actual system completed and “flight qualified” through test and TRL 8

System prototype demonstration in a space environment TRL 7

System/subsystem model or prototype demonstration in a relevant TRL 6

Component and/or breadboard validation in relevant environment TRL 5

Component and/or breadboard validation in laboratory environment TRL 4

Analytical and experimental critical function and/or characteristic TRL 3

Technology concept and/or application formulated TRL 2

Basic principles observed and reported TRL 1

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Table 10-2 Development Durations for TRLs

10.1 Technology Development and Schedule:

The development period for the individual payload instruments cannot be predicted in an accurate manner and is based on available information and for some payload on best and general estimation.

Apart from the Communications Payloads (at already high technology maturity), all other payloads candidates are selected from “cancelled” European national missions (i.e. LEO, MoonLITE and MAGIA Mission). In details:

The candidate payloads for LEO:

• Covering mapping and gravimetry objectives are already at TRL 6 or higher

• Covering the characterization of lunar environment (RadMo and LEOPARD) are at TRL 4-5 with advanced design of flight instruments.

For the candidate payloads such as MAGIA ( Mapping/gravimetry/environmental and sub-satellites) or LunarEX (sismology – penetrators) an accurate status of development is currently not available.

For these payloads, it is assumed that they are at minimum TRL 4 at mid of phase A and able to reach level 5/6 at the end of Phase B.

10.2 AIV Approach

10.2.1 Assumptions

Considering that the structure shall allocate specific payload, it is here considered that the platform structure is a dedicated one and its qualification is achieved with one STM (Structural-Thermal-Model).

It is assumed that the S/C bus consists in major parts of already flight proven and off-the-shelf equipment and so platform FM components (e.g. Power, AOCS, Data Handling, etc.) will be delivered in a qualified status and will not be the overall schedule driver. This is considered applicable also for the Electrical Propulsion S/S if this option will be chosen (see high TRL status of the electrical thrusters).

The mechanical and thermal design, as well as the functionality of the payload instrument Flight Models (FM) will be qualified by the suppliers prior to their delivery.

The FM Instruments will be delivered to the System after full acceptance verification at lower level.

12 years + 1,5 year1-210 years + 1,5 year2-38 years + 1,5 year3-46 years + 1,5 year4-54 years + 1,5 year5-6

DurationTRL

12 years + 1,5 year1-210 years + 1,5 year2-38 years + 1,5 year3-46 years + 1,5 year4-54 years + 1,5 year5-6

DurationTRL

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Due to the estimated standard average required time for the development, it has been considered a TRL of the instruments at the current date in line with the expected required launch date (not later than 2017).

It is assumed that the TRL shall reach level 4 minimum at mid of the phase A in order to define the payload suits.

At the end of phase B, the technology maturity shall be well beyond TRL 5 for all payloads.

The qualified flight units shall be available during Phase D for the integration and system testing.

The STM and FTB (Functional-Test-Bench) components of each payload will be delivered at a build status, which is representative for the mechanical, electrical and functional interfaces of the final FM component.

An aadvanced procurement of long lead items shall be considered, as needed by consolidated detailed schedule.

Concerning the System Models Flow the assembly and test activities of the models STM and FM is considered sequential, as per the standard verification approach.

10.2.2 Model Philosophy

Due to the specific mission and configuration a new built platform has been considered and a model philosophy composed of three models at spacecraft system level is proposed. This philosophy allows a decoupling of verification activities on the different S/C models and thus, introduces programme flexibility and reduces schedule risks.

Structural Thermal Model (STM): will ensure the mechanical and thermal qualification of the spacecraft main structural design. Most of the unit assemblies will be represented by structural dummies with heaters.

Functional Test Bench (FTB): will ensure the verification of the electrical, functional and software interfaces of the payload instruments. Beside the thorough end-to end-testing and the operational verification, the FTB will also be used for the software development and the validation of system level procedures. Breadboard units, designed with off-the-shelf equipment and, in case of need, with interface simulators can be used most of the time. These breadboards must be available well before the envisaged delivery date of the flight models

(Proto) Flight Model ((P)FM): will ensure the integration and verification of the flight model structure, equipment and payload at full system level. The (P)FM of the payload instruments are assumed to be delivered in qualified status (structural, thermal and functional qualification campaign performed by supplier).

10.2.3 Schedule

The master schedule exercise gives an overview on the on ground activities, phases and major system reviews for all options apart from options considering a hybrid propulsion system. The impact will be an additional 2 to 3 months due to an increase in the number of interfaces.

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Figure 10-1: Programmatic System Schedule

LUMETTO


page 87 of 102

Figure 10-2: Programmatic System Schedule (details)

LUMETTO


page 88 of 102

The overall schedule is driven by the currently assumed development status of the instruments.

The schedule exercise showed that the on ground sequence phases C and D shall be performed and completed in less than four years considering a reduced development time of the instruments. Only a three years time allocation has been considered for the instruments at current lower TRL (<4) to perform the development from TRL5 to delivery of verified FM units since this task is critical for the overall schedule with the envisaged launch date within 2017.

For payloads with TRL level of 4/5 (or higher) already at the current date, the time window for their development till delivery of qualified and accepted flight unit is around the one presented in the ESA Master Plan.

Higher risks than normal is envisaged in case of longer technological development time required for the instruments.

Furthermore with the above assumption, it can be noted that:

• The System STM and ETB are not on the critical path

• Use of the System MGSE for both models (STM and PFM) in sequential way (meaning one set only)

• The PFM AIT activities could be optimized but no consistent gain in schedule can be obtained.

Considering the Penetrators, they may be the payload requiring the most development effort and time. Assuming the Penetrators as electrically self-standing equipment, with request to the Orbiter or only with simple (limited) electrical interfaces:

• Their impact on the overall schedule could be mitigated considering that the Penetrator’s availability for the integration at System level could be delayed (of about 6 months) with a limited higher risk in term of verification.

• Using STM equipment for the System FM Thermal and Mechanical campaign.

• Providing a successful and complete thermal and mechanical verification at lower level (Penetrator S/S).

• Performing the integration of the flight units at later stage.

Potential safety concerns on the integrated Penetrators during the AIT operations might need to be considered as well.

10.2.4 Programmatics/AIV - Impacts of Identified Options

For clarity the impacts of the identified options are described below based on the LUMETTO trade tree in Appendix A.

Launcher: Ariane vs Soyuz:

• Different static and dynamic requirements for the design and testing. Very limited schedule impact on testing schedule at system level. No problem envisaged for the testing facility capability.

• Similar logistic scenario during the launch campaign.

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• In case of dual launch with Ariane, less flexibility wrt combined operations with launcher because also a second customer is involved and launch window flexibility.

Propulsion S/S (Chemical vs Electrical/Hybrid)

• The Chemical propulsion has high heritage vs the Electrical one in term of AIV. Electrical Propulsion S/S will require development for higher power thrusters versus the already in operations (Smart, GOCE) or under qualification (Bepi-Colombo). High voltage harness is required for the Electrical Propulsion.

• Propellant simulant loading/draining and drying for the Chemical Propulsion during the mechanical system testing. Loading of Xe since the mechanical testing for the electrical propulsion. No significant overall schedule impact.

• For the Hybrid scenario, more a complex system with higher quantity of mechanical and electrical interfaces needs to be verified.

Single Stage / Multi Stage

• The Multi-Stage approach is implemented already in other programmes (see Bepi Colombo). It is more complex from AIT/AIV point of view having more mechanical and electrical interfaces to be checked and implying the use of additional pyrotechnics and mechanisms. Use of additional explosive actuators could imply higher shock environment for the payloads and additional test at system level of separation test at STM and PFM.

• Use of existing / new Platform

• The use of an existing Platform will allow making advantage of heritage reducing the Models representativeness needs in the Model Philosophy. This potentially would allow some schedule and cost saving.

• It is unlikely that an existing platform can be re-used “as is” though adaptation will be required

o The effort to make design changes of an existing platform to meet the LUMETTO Mission and Payload Requirements will actually reduce the benefits of the heritage.

o To adapt the platform to the specific payload configurations and design factors, could likely lead to modification (structure, harness routing, propulsion routing), which cannot be verified with a pure Prototype approach without high programmatic risk.

The approach with a dedicated platform will better fit, with less technical risks, in terms of design and configuration the specific needs for the mission and payload requirements (L.o.S, stability, accessibility, loads, sub-satellites and penetrators presence).

• The qualification at system level would be performed through a Model Philosophy including a Structural Thermal Model (STM), Electrical Test Bench (ETB), and (Proto) Flight Model. This is the less risky approach in term of verification.

• During the STM also the MGSE interfaces, test facilities operations, mechanical procedures will be verified in advance as well as the training of the personnel will be performed.

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• The structural qualification would run in advance on the STM with a later need for the Flight units to be integrated on the (P)FM structure. The STM approach would allow to validate the thermal model and to verify the thermal subsystem in advance w.r.t. the Flight Model.

• In this case the proposed Model philosophy would be : o Electrical Test Bench (ETB) o Structural Thermal Model (STM) o (Proto) Flight Model ((P)FM).

The use of an existing platform would have benefits in term of cost and schedule (even if its design and procurement has not been identified inside the critical path).

• But for the specific mission of LUMETTO, it is likely that an existing platform cannot be re-used and at the minimum it will need to be modified to meet the mission and payload requirements, loosing partially the achieved heritage in term of qualification and with worse condition in term of accessibility and routing.

• A general statement cannot be expressed at this point but the suitability should be evaluated on a case by case basis. It would be advisable to ask the Industry to evaluate the re-use of their existing platforms, for potential cost saving purposes.

• Anyhow for the above reasons, it is very unlikely that a pure PFM Model philosophy can be implemented. A System Structural Model shall be considered as baseline in order to verify in advance the achievement of qualification for the static and dynamic requirements from the launcher and with less programmatic risk.

• In this case the proposed Model philosophy would be:

• Electrical Test Bench (ETB)

• Structural Model (simplified) (SM)

• Flight Model (FM).

10.3 Conclusions

A schedule exercise was performed to identify the critical path in the development of the Programme, based on some assumptions.

The qualitative impact has been evaluated for some identified options as required from the study objectives.

The achievable launch dates are driven by the current development status of the technology implemented in the payload instruments and by their availability during Phase D at system level.

The development period for some payload instruments cannot be predicted in an accurate manner. For the development phase some key points have to be considered for the payload at current lower TRL:

The TRL of the instruments shall have reached at least level 4 at mid of the phase A of the LUMETTO spacecraft mission sequence (July 2011) for the definition of the payload suites.

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At the end of Phase B (June 2013), the technology maturity for the Instruments shall be well over TRL 5.

The completion of the Instruments development (from TRL 5) and qualification up to delivery of Instruments Flight Unit to the System AIT shall be performed in a reduced time and being completed in less than three years (mid 2016).

Launch date: end of 2017 for all options.

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This Page Intentionally blank

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11 COST

11.1 Costing Assumptions

Table 11-1: CBS with cost assumptions

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page 94 of 102

ROM Cost estimates were performed by ESA cost engineering TEC-SYC for the three different options e.g. Soyuz Hybrid, Soyuz Electric and Ariane 5 full P/L Suite.

The resulting range is between 440M€ and 460M€. This includes all cost elements listed in Table 11-1 except for the External Project Events.

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12 CONCLUSIONS The LUMETTO CDF study was the first of a set of TEXP studies performed in a “fast track fashion” by performing a “quick assessment” to gain an overview of the main mission boundaries and drivers.

Different mission design concepts compliant with the mission requirements and cost frame were defined and traded against each other. Out of an initial set of more than 60 options three

• Soyuz-Hybrid-Single-Stage

• Soyuz-Electric-Single-Stage

• Shared-Ariane5-Hybrid-Single-Stage

were identified as the most promising scenarios and are proposed to be looked at in more detail.

A set of supportive analysis was produced to better understand the decisions and to facilitate the follow up also on options other than the three mentioned above.

Overall the LUMETTO mission is considered to be a pioneering mission. Elements like the advanced transfer strategy to reach the Moon by using chemical propulsion to pass the van-Allan-belts, go to Sun-Earth WSB point L2 and then heading back to the Moon with electric propulsion; will lead potentially to a dedicated GSP study and being beneficial for many future lunar missions.

The innovative use of the telecommunication payload adds another original point to mission.

Furthermore LUMETTO:

• Will be an evolution or follow-up on ESA lunar mission experience such as Smart-1

• Will support exploration activities direct (TLC service, surface mapping…) and indirect (technologies…)

• Will deliver valuable science.

The next logical steps would be to:

• Identify a baseline and (the two) options

• Perform a bottom-up, equipment level S/C design

• Perform a more detailed configuration and structure analysis

• Perform a more detailed mission analysis & (innovative) transfer strategies

• Perform a more detailed radiation assessment

• Perform a life cycle costs analysis

• Identify further technology needs of platform equipment

• Look at mission alternatives.

The LUMETTO mission can potentially provide the demonstration of necessary technologies to enable future Lunar missions.

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13 REFERENCES RD[1] NANOSAT, CDF Study Report: CDF-84(A), February 2009

RD[2] Space Project Management, risk management, ECSS-M-ST-80C, 31 July 2008

RD[3] ESA Cost Engineering Chart of Services, Issue 3

RD[4] Cost Risk Assessment Procedure TEC-SYC/GRE/SA/2006/021_02

RD[5] Technology Readiness Levels, NASA White Paper, J.C. Mankins, April 1995

RD[6] A Combined Moon Data Relay Orbiter and a Moon Lander Mission Concept, M. Wittig and B. Hufenbach, ESA

RD[7] Moon Exploration, ASI presentationto 11 ISECG Meeting – Montreal – July 2008

RD[8] LEO – A Pacemaker for German Exploration, F. Claasen, DLR Bremen, September 2008

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14 ACRONYMS Acronym Definition

ADC Analogue Digital Converter

AIT Assembly Integration & Test

AIV Assembly Integration & Verification

AOCS Attitude and Orbit Control System

CCD Charged Coupled Device

CDF Concurrent Design Facility

CDMU Command Data Management Unit

CoM Centre of Mass

DHS Data Handling System

DSA Deep Space Antenna

EADS European Aeronautic Defence and Space Company

EES Earth Exploration Satellites

EoL End of Life

EoM End of Mission

EPS Electrical Propulsion System

ESA European Space Agency

ETB Electrical Test Bench

FCV Flow Control Valve

FDV Fill/Drain Valve

FoV Field of View

FPA Fine Pointing Actuator

GEO Geostationary Earth Orbit

GNSS Global Navigation Satellite System

GPS Global Positioning System

GS Ground Segment

GSP General Studies Program

GTO Geostationary Transfer Orbit

GSE Ground Support Equipment

HEO Highly Elliptical Orbit

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Acronym Definition

HICDS Highly Integrated Command and Data System

HLR Human Lunar Return

HRSC High Resolution Cameras

IDM Integrated Design Model

INS Inertial Navigation System

ISECG International Space Exploration Coordination Group

LEO Lunar Explorer Orbiter (German Mission)

LEOP Launch and Early Operations Phase

LLO Low Lunar Orbit

MAGIA Missione Altimetrica Geochlmica LuAre

MIPS Million Instructions Per Second

MLI Multi-Layered Insulation

OBC On-Board Computer

PCDU Power Conditioning and Distribution Unit

ROM Rough Order of Magnitude

SADm Solar Array Drive Mechanism

SDRAM Synchronous Dynamic Random Access Memory

SEP Solar Electric Propulsion

S-E WSB Sun-Earth Weak Stability Boundary

SOCC Science Operations Control Centre

STM Structural Thermal Model

TBC To Be Confirmed

TBD To Be Determined

TEXP Technology for Exploration

UHF Ultra High Frequency

LUMETTO

CDF Study Report: CDF-103(A) Month Year

page 101 of 102

APPENDIX A - LUMETTO TRADE TREE

Figure A-14-1: LUMETTO Trade Tree

LUMETTO

CDF Study Report: CDF-103(A) Month Year

page 102 of 102

Trade 1 Trade 2 Trade 4 Trade 5Launch Vehicle Propulsion Type Staging PlatformSoyuz Ariane Chemical Electric Hybrid Single Stage Dual Stage Existing Dedicated

Systems

AOCS

EP and slow extraction from van allen belts is not looked as a major risk, since it has already been dealt with on SMART-1 or SPIRALE with ''cheap'' sensors. State of the art Star Trackers for example are not anymore bothered by the belts. Gyro MEMS is also very resistant to TID.

SGEO P/F is already dealing with Hybrid Main propustion (400N engine and EP - but not the EP thruster we plan to select)

More tuning configurations to deal with, and additional failure cases.

Modifications needed to reuse SGEO P/F : - Minor : positionning of the STRs, which are placed on the P/F to avoid seeing earth & sun on a GEO orbit. - Minor : The gyros used could be changed (without major redesign to save around 7 kg and also to save power) - Major : The wheels can be downscaled - Major : the EP is not the one we need (SPT100 instead of 5000) and its positionning on the S/C (also the xenon tanks) is not effective (accomodated for NS Station keeping in GEO). - Major : a mechanism similar to Smart 1 for the EP has to be implemented (2-axis control of EP thrusting vector) to avoid too many wheel offloading (using chemical propulsion) - In any case, AOCS SW has to be modified.

Pro :'- Redesign to reduce mass- W.R.T the big effort to exchange main propulsion needed, pointing modes will be different than existing p/f due to the mission around the moon, in case a complete p/f redesign is needed it will not necessarily be much more expensive or risky.

Electric Propulsion

A5ME capability (restartable upper stage) was not considered within the study.A5ME would allow also HEO in orbit insertion as in the soyuz case: EP start after van allen belt escape with reduced transfer time to lunar orbit and reduced cumulated radiation dose.



Chemical Propulsion



Power



Structures / Configuration Small GEO Platform a viable solution

Cost

Lower Cost Dedicated Launch (Higher cost out of CGS) (Arianespace alignment of costs with A5)

Variable cost (likely scenario higher cost) New dev (more expensive) existing (cheaper)

EP only more expensive at component but requires more detail for full system. Synergies with Power and AOCS

2 paralell systems most expensive, but it is a technology development

Single Stage, no interfaces required (lower Cost)

Only makes sense if many recurring units will be used over time.

Cheaper (depends on size and the delta changes required) More engineering but could be smaller and more versatile.

Risk

Highly reliable launch vehicle with a successful launch record. Single launch advantage with respect to shared launch scenario (no interference with second payload) (++)

Interference with second payload of a shared Ariane 5 launch. Schedule delays in delivery, launcher integration and flight readiness. Need to find compatible secondary payload, schedule constraints. (+)

Reliable system with heritage. Shorter lunar transfer time is a plus given the reduced exposure to high radiation areas. (++)

Reliable system with heritage assuming fully redundant thruster assembly and control electronics (PPU). Would require long lunar transfer time with negative implications such as higher radiation exposure during van Allen belt crossings with consequential increase in electronics degradation/malfunction (SEU, SEL, etc.) , solar array degradation, and reduction in overall mission lifetime. Soyuz inserted initial HEO with apogee 200,000km would mitigate radiation impact (-)

Higher complexity as compared to fully electric or fully chemical propulsion options. Lower radiation effects risk as compared to fully electric option. (+)

Single stage offers less complexity and higher reliability (+)

Dual stage is more complex (incl. addition of separation mechanisms) (-)

Existing platform reduces development risk (cost/ schedule implications)

Dedicated platform increases development risk (cost/ schedule implications)

Programmatics

Pro:1) lower mechanical test levels at system level (minor).2) single launch : more flexibility wrt launch date and launch facility for combined operations with launcher allocation to the single ProgramSame:1) launch site in French Guyana (same logistics,shared PPF and HPF facilities with Ariane) and Baikonur (tbc).

Same:1) launch site in French Guyana same logisticsCON:1) Higher mechanical system test levels (minor)

CHEMICALPRO : 1) Chemical propulsion has high heritage in verification,

SAME:1) Minor effects on the schedule for system testing and loading/unloading operations.2) During the mechanical testing the tanks shall be filled with simulant and then drained and dried (minor)

ELECTRICALCON : 1) Electrical propulsion S/S needs development but the TRL is high (used on Smart, Goce, under qualification for BepiColombo).(medium)2) During the mechanical testing the Xe tank shall be filled with Xenon and kept till the launch (minor)3) Requires high voltage harness and some safety constraints during operations and testing (minor)SAME:1) Minor effects on the schedule for system testing and loading/unloading operations.

HYBRIDCON : 1) Electrical propulsion S/S needs development but the TRL is high (used on Smart, Goce, under qualification for BepiColombo).(medium)2) During the mechanical testing the Xe tank shall be filled with Xenon and kept till the launch (minor)3) Requires high voltage harness and some safety constraints during operations and testing (minor)SAME:1) Minor effects on the schedule for system testing and loading/unloading operations.

SINGLE STAGEPRO:1) Easier from AIT/AIV point of view (less mechanical and electrical interfaces to be tested, no additional pyrotecnics or mechanisms)

DUAL STAGECON : 1) Integration and testing of additional hardware (separation system, pyrothecnics, umbilicals).2) Separation test at sytem level to be added with shock mesurement (according to the type of explosive actuators).Some impact on schedule and some additional MGSE to support the separation system (impacts but not big)

Anyway note that it will be not a new approach, see BepiColombo (multistages)

PRO:1) Use of an existing Platform will allow to gain the heritage reducing the Models representativity needs in the Model Phylosophy.2) Some schedule and cost saving CON:1) Effort to make design changes opf the existing platform to meet the LUMETTO Mission and Payload Requirements which will actually reduce the benefits of the heritage.

NOTE : THE USE OF AN EXISTING STRUCTURE, VERY LIKELY WILL REQUIRE SOME MODIFICATION IN THE DESIGN TO ALLOCATE THE NEW PAYLOAD AND TO MEET THE DESIGN LOAD FACTORS. IT IS UNLIKELY THAT A PURE PROTOFLIGHT APPROACH WILL BE FEASIBLE WITH DELTA QUALIFICATION DURING THE SYSTEM TESTING. IT CANNOT BE EXCLUDED AT THE PRESENT TIME, BUT MORE LIKELY AT LEAST A SYSTEM STRUCTURAL MODEL SM WILL BE REQUIRED IN ORDER TO PERFORM THE STATIC LOAD TEST ON THE MODIFIED STRUCTURE AND MECHANICAL ENVIRONMENTAL TESTING AT QUALIFICATION LEVEL.

PROPOSED MODEL PHILOSOPHY:

PRO:1) The design and configuration will better fit the specific needs for the mission and payload requiremnts (LoS, stability, accessibility, loads)2) The qualification at system level to be performed through a Model Philosophy including a Structurat Therma Model (STM), Electrical Test Bench (ETB), and Flight Model. This is the less risky approach in term of verification. During the STM will be verified in advance also the MGSE interfaces, test facilities operations, mechanical procedures and personnel training.3) The structural qualification can run on STM with a later need for the Flight units to be integrated on the FM structure. 4) The STM approach would allow to validate the thermal model and the verify the thermal subsystem in advance wrt the Flight Model. CON:1)Wrt existing platform : Design of a new Structure , harness and Propulsion piping routing. Anyhow also an existing Platform could very likely require big effort in re-design phase for what concern new interfaces and load calculation. 2) Cost and potential schedule impact

PROPOSED MODEL PHILOSOPHY Table A-14-1: Trade-Space Qualitative Analysis Table

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